falcon 900 pilot training manual

8/15/2019 Falcon 900 Pilot Training Manual

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Courses for the Falcon 900 and other Falcon aircthe following FlightSafety learning centers:

FlightSafety InternationalTeterboro Learning CenterTeterboro International Airport100 Moonachie AvenueMoonachie, New Jersey 07074Phone: (201) 528-0100



INSERT LATEST REVISED PAGES, DESTROY

LIST OF EFFECTIVE PAG

Dates of issue for original and change

Second Edition .......0 ..............June 2007

NOTE

For printing purposes, revision numbers in fotom of every page that has changed in any wagraphical revisions, reflow of pages, and oth

necessarily affect the meaning of the manual

THIS PUBLICATION CONSISTS OF TH

Page *RevisionNo. No.

Cover ...................................... 0

Copyright ................................ 0iii–vi ii ...................................... 0EC-i–EC-ii .............................. 0NP-i–NP-vi .............................. 0NP-1–NP-52 .......................... 0AP-i–AP-vi .............................. 0

AP-1–AP-122 ........................ 0EP-i–EP-iv .............................. 0EP-1–EP-50 .......................... 0LIM-i–LIM-iv .......................... 0LIM-1–LIM-22 ........................ 0MAP i MAP i 0

CRM-i–C

CRM-1–SR-i–SRSR-1–SMWS-i–MWS-1–APP-i–A

APP-1–A

PageNo.



FOR TRAINING PURPOS

NOTICE

The material contained in this training manual is obtained from the aircraft manufacturer’s Pilot Man

Manuals. It is to be used for familiarization and tra

At the time of printing it contained then-current infoof conflict between data provided herein and that

by the manufacturer or the FAA, that of the manufatake precedence.

We at FlightSafety want you to have the best train

come any suggestions you might have for improviother aspect of our training program.



CONTENTSEXPANDED CHECKLIST

Normal Procedures

Abnormal Procedures

Emergency Procedures

LIMITATIONS

MANEUVERS AND PROCEDURES

WEIGHT AND BALANCE

PERFORMANCE

CREW RESOURCE MANAGEMENT

RECURRENT

System Review

Master Warning System

APPENDIX



EXPANDED CHECK

CONTENTS

NORMAL PROCEDURES ..................................

EMERGENCY PROCEDURES...........................

ABNORMAL PROCEDURES.............................

FALCON 900 PILOT TRAINING MANUAL



NORMAL PROCED

CONTENTS

BEFORE START (POWER OFF)........................

COCKPIT CHECK...............................................

Overhead Panels..........................................

CONSOLE/OVERHEAD PANEL .......................

Left Console................................................

Pilot Instrument Panel.................................Center Instrument Panel..............................

Copilot Instrument Panel ............................

Right Console..............................................

Pedestal .......................................................

BEFORE START (POWER ON) .........................

After APU Start or Engine 2 Startor if a Ground Power Unit Is Used .............

STARTING ENGINES.........................................

Starting Problems........................................

Start.............................................................

BEFORE TAXI ....................................................

TAXI ....................................................................

BEFORE TAKEOFF




ADVERSE WEATHER CONDITIONS ........................

Operating in Icing Conditions ..............................

Cold Weather Operation .......................................

Severe Turbulence Penetration .............................

Windshear Situation..............................................

F




ILLUSTRATIONS

Figure Title

NP-1 Overhead Switch Panel ..................

NP-2 Warning Panel ................................

NP-3 Buffet Onset Envelope....................

TABLE

Table Title

NP-1 Landing Computations ..................

NP-2. Three Engines Operative .................




NORMAL PROCED

NOTE

This Normal checklist is designed onlpurposes. Where checklist procedure

the Airplane Flight Manual, the Ai Manual takes preference.

NOTE

Items marked with an asterisk are to beon the first flight of the day with the s

BEFORE START (POWER OFF)1. *Preflight Checklist..........................................

The aircraft exterior and interior preflight chethe chocks removed before closing the door an

2. *Documents and Keys .....................................

The aircraft flight, performance, weight amanuals must be on board before dispatch. Theand Airworthiness must also be readily availabAny other documents as required by company

this time. Any keys and/or security items requon board as well.

COCKPIT CHECK




3. FMS Master (LH and RH) .........................................

In order to save battery power when turning on the bchecklist, these master switches should be placed oswitches are “in” or flush with the surrounding panel

4. Bus-Tied Switch..........................................................

By separating the buses, this allows an individualturning on the batteries later in the checklist. Sepallows for detection of low battery voltage and proper

5. APU Generator Switch................................................

This action satisfies APU start interlock circuitry reqand APU generator field circuit.

6. BAT 1–BAT 2 Switches ..............................................

This prevents the needless discharge of the batteries

checklist is complete.

7. GEN 1–GEN 2–GEN 3 Switches ...............................

The generator switches must be placed in the onprovide engine start interlock circuitry.

8. *IRS1–IRS 2–IRS 3 Battery Voltages ........................

These test buttons are depressed individually to terespective auxiliary batteries on the overhead pMinimum voltage is 24 volts.

F




12. CMPTR 1–2–3 Switches ..................................

The engine computers must be in the auto pothe engines. Ensure that the CMPTR lights oare out when electrical power is applied to the

13. Start Selector Switches (All 3).........................

These three-position switches must be placedposition in order to provide ignition and painterlock for starting the engines.

14. XTK Switch.....................................................

Ensure that the switch used to connect Groupfuel tanks is in the neutral position.

15. Booster (Pump) Switches (All 3) .....................

This limits battery discharge when they are pla

16. X-BP Rotary Switches (All 3) .........................

These three rotary switches are closed to proconfiguration for starting the engines and takeo

17. XTK 2 Switch (If Installed)..............................

This switch, if installed, is placed to the AUTOoperation of the valve between the forward and

18. HP and PRV Bleed Switches (3) ......................




20. Isolation Valve Knob...................................................

This switch is in the horizontal (open) position in bleed-air manifold can be supplied by all three engine

21. Crew and Passenger Air-Conditioning Valve Switche

Placing these three-position switches to the AUTO pcircuit to automatically close the valves when takethen open the valves slowly after takeoff.

22. BAG Switch................................................................

This three-position switch allows normal pressurizatthe baggage compartment.

23. Windshield Switches (3) .............................................

This position ensures that the demand on the batterwhen they are turned on.

24. Anti-ice Switches (4)...................................................

This precludes unwanted anti-ice system operation after

25. Pitot Heating Switches (3) ..........................................

This position ensures minimal battery discharge whand prevents overheating of the pitot static componen

26. Wiper Switches (2)......................................................

F




CONSOLE/OVERHEAD PANEL

LEFT CONSOLE

1. Data Loader ......................................................

The FMS data loader should be turned off unti

powered by other than batteries and the data lothe flight management system.

2. Oxygen Mask....................................................

The pilot’s oxygen mask must be checked as100%, and with both the hose and communi

the proper positions. Press the test button on toxygen.

3. IRS 1 (and Optional IRS 3) ..............................

IRS 1 should be off until the aircraft is powerunless the checklist calls for the IRS to be tur

IRS 3 is installed, it also should be off.

4. Audio Control Panel .........................................

a. SPK, ST Audio VHF 1 and VHF 2.............

b. Microphone VHF and C’PIT......................

c. All Other Pushbuttons.................................

d. VOR–DME Potentiometer ..........................

Set the pilot’s radio jackbox for the proper ra




CENTER INSTRUMENT PANEL

1. Radar (2)......................................................................

If two radar control panels are installed, the masteturned off until the checklist calls for the radar toaction will save the batteries and prevent damage

injury to personnel.

2. Fuel Shutoff Switches (3)............................................

These three switches should be in the guarded positiofuel to the engines for starting. These switches are tevent of an engine fire.

3. Fire-Extinguisher Switches (5) ...................................

These five switches must be in the zero/safetied inadvertent discharge of the fire bottles.

4. Normal L/G Control....................................................

It is imperative that the normal landing gear hanposition to preclude the inadvertent retraction of thpower is applied to the aircraft.

5. Gear Pull Handle.........................................................

This ensures normal electrical retraction sequencingand, when required and operated, allows landing gear

6. Brake Selector Switch .................................................

F




9. Thrust Reverser Switch.....................................

This switch is usually kept in this position tothe thrust reverser while on the ground. The swstowing of the thrust reverser either on the grou

COPILOT INSTRUMENT PANEL1. Automatic Cabin Pressure Controller ..............

This three-position switch on the controller pthe pressurization system based on a predetebarometric control.

2. DN/UP Knob....................................................

The DN position (green index) allows that commanded to an increased pressure conditiofailure of the automatic cabin pressurizatiocommands a cabin altitude rate of descent oper minute.

3. Auto/Manual Selector Switch...........................

This allows for the normal operation of the cab

4. NORM/EMERG Selector Switch.....................

The NORM position permits full normal operand pressurization systems. The EMERG abnormal and/or emergency checklist procedur

5 D S i h




8. GPWS Flaps O’RIDE Switch .....................................

This switch, when not guarded and in the override “TOO LOW FLAPS” warning given by the GPWS usually occurs when on final approach and wheconfigured in the SLATS + 40° flaps configuration.

RIGHT CONSOLE

1. Passenger Oxygen Valve Selector...............................

Placing the passenger oxygen valve in the normaautomatic deployment of the passenger oxygen maaltitude exceed 10,000 feet during flight.

2. Oxygen Pressure..........................................................

Minimum oxygen pressure with no passengers oassuming the flight stays under 10,000 feet altitude. Tthe consumption requirements of each crewmemberecommended that if passengers are carried, the oxcharged to its capacity. This is especially true if the oceanic areas or over desolate terrain. The aircraft perbe used as reference as to the minimum pressure requimission with passengers and is based on the following

• The flight is being flown either above or below above 41,000 feet, one of the pilots must breathe o

• An emergency descent to 10,000 feet is made witaircraft breathing oxygen on NORMAL.

• Further flight is conducted at 10,000 feet or belousing first aid oxygen

F




5. Autoload Shed Switch (If Installed) ................

This switch allows for the resumption of elebus items if a generator disconnects from its bu

6. IRS 2.................................................................

IRS 2 should be off until the aircraft is powerif the checklist calls for IRS programming.

7. Oxygen Mask....................................................

The copilot’s oxygen mask must be checked

to 100%, and both the hose and communicthe proper positions. Press the test button onof oxygen.

8. Cond Lever (If Installed)..................................

This lever must be placed in the full forward

the crew and passenger air-conditioning systThis lever has been removed on SN 165 and su

9. Nose Lever (If Installed) ..................................

This lever connects the cabin with the nose cothe nose cone in flight. On later aircraft (SN 7lever has been removed.

PEDESTAL

1. Power Levers (All 3) ........................................




4. Normal Tailplane Circuit Breaker ...............................

The mechanical circuit breaker, located aft of the emswitch, must be in the down position in order to profor operation of the normal elevator trim system.

5. Flap-Slat Handle..........................................................

Ensure that the flap-slat handle is fully forward in the starting the engines or applying pressure to either the N

6. Emergency Slats Switch..............................................

This switch controls the emergency operation of thshould be used only in the event of a main system mthe switch should be kept in the off and guarded posit

7. MMO Switch (If Installed) .........................................

This switch, if installed, allows adjustment of the M

warning based on aircraft gross weight.

8. Landing Gear Emergency Extension Handles (3) ......

Check that the nose landing gear emergency extensithe left aft side portion of the center pedestal, is fullythe clips provided. The main landing gear extension hfloor area to the right and left of the pilot and copilmust be stowed by pushing down on the handles access doors.

BEFORE START (POWER ON)

F




Check the battery temperatures as fol

CAUTION


Figure NP-1. Overhead Switc



• If either battery temperature reaches 1160°F on aircraft SN 132 and subsequentred HOT BAT lights are on, the battery turned off, monitored while it cools, and rprior to takeoff.

NOTE

On the average, the rate of battery coolingground is 1°F per minute.

2. Battery Voltages..........................................................

Minimum battery voltage for APU start is 23 volts aengine start.

3. *HRZN Battery Voltage .............................................

The standby horizon battery voltage test button isleft voltmeter on the overhead panel. The reavolts minimum.

If the No. 2 emergency battery is installed, its voltagand the reading should be 24 volts as well.

4. Bus-Tied Switch ..........................................................

This rotary switch controls the bus tie relay, which in TIED light on the master warning panel. The switchposition to tie the left and right main DC buses togethe

5. Park Brake/No. 2 Park Brake Light ............................

Pull the parking brake handle to the first detent hes

F




6. Fire Detection ...................................................

Set the test control switch, located on the mFIRE position. All detection and warninsimultaneously and the aural warning will soumalfunctioning, the corresponding light will nthe test control must be held in the FIRE po

COMP smoke detection light comes on, that mDo not silence the aural warning until the FIRon. After silencing the aural warning, releasecenter. It may take up to a maximum of 10 COMP light to extinguish.

7. COND BAT Pushbutton Light (SB 125) ..........

Aircraft with SB 125 will have a toggle switnormally located on the copilot’s instrument by the toggle switch must be closed before tak

8. Navigation Lights.............................................

Although not a mandatory requirement, the Fturn on the aircraft position or navigation lighis applied to the aircraft.

9. and Lights

and Wording “NO TAKEOFF”.........................

Move the No. 2 power lever from the cutotoward the takeoff position. The T/O CONFshould illuminate. Move the power lever backthe lights should extinguish.

T/O CONFIGENG 2 FAIL




3. APU Start Switch (1 Second)......................................

Do not hold the APU start pushbutton in for moreholding coil, if operating properly, should hold through the start sequence. The starter will stay engOIL light should extinguish by 60% N1. The GEN liat 97% N1 + 4 seconds, signifying the generator is

APU master light should start blinking after the stone of the start interlock circuit requirements may nor the APU may shut down for activation of a protect

4. APU N1-T5 Indicators—APU Generator Volts/Amps

After the APU is on speed and the generator has gon

APU voltage output is 28.5 volts. The amperage drout at the maximum of 350 amps. If desired, it mayoff one of the batteries to ease the demand on the APcharge draw for the battery being charged is less thaother battery may be turned on for charging. Closelcharging and the APU generator load before procechecklist. The load demand on the APU should bebefore performing an engine start.

Discontinue start (STOP PUSHBUTTON)does not rise within 10 seconds. Wait 5 minutto attempting a second start.

5. APU Bleed-Air Switch................................................

It is recommended that a 1-minute minimum waitinbefore turning on the APU bleed. If possible, wait uncharged to less than a 50-amp draw per battery. Apreserve the life of the APU which can be plac

CAUTION

F




A ground power unit providing 28.5 volts DCmust be connected at the receptacle, located onof the aircraft. Move the two-position switcPOWER. The BAT 1 and BAT 2 lights on thilluminate, as well the BUS TIED light. The Ncompletely isolated from the electrical systemDC buses are automatically tied together.

AFTER APU START OR ENGINE 2 STAOR IF A GROUND POWER UNIT IS US

1. LH AV Master and RH AV Master ..................

Light (LH and RH)....................

2. FMS Master (LH and RH)................................

Lights (LH and RH) ..................

If all four switches are installed, turn them on

3. Maintenance Test Panel ...................................

The test panel at the right side panel of the cochecked for any red indications at one or more the panel. Push the reset button at the center oindicators, or address the malfunction if the red

4. Voice Recorder .................................................

5. IRS (3)/FMS (2) VLF-Omega .........................

POS Sensors

FMS

MASTER

AV

MASTER




22. Rudder/Aileron Trim........................................

Check the operation of the aileron and ruddeactuating the double switches on the center peobserving the trim indicators and movement othe direction in which the respective trim checking movement, center the trim to zero fo

check that the trim does not move when only each surface is actuated.

23. No. 2 Stall Test Pushbutton..............................

Depress the No. 2 stall button on the center pestall warning system. The following actions sh

• The aural stall warning sounds and cannot b

• The IGN lights on the overhead panel illum

• The outboard slats extend. A red light follolight should be observed on the slat/flap ind

24. Standby Pump Switch ......................................

Place the standby hydraulic pump switch terminates the prestart hydraulic checks.

25. Cabin Pressure Controller ................................

Cabin Aural Warning........................................

Push the test button located on the bottom of tThe cabin altitude aural warning should be hThe red CABIN light should illuminate. Th


F



Check battery temperatures:

• if either battery temperature is above 12the amber WARM light is on, do not attemtery start. The APU/engines must be starte

a ground power unit.

• if either battery temperature exceeds 120ing starting, monitor temperature changfew minutes after starting.

• if either battery temperature exceeds 140ing starting, wait until it drops to 120°F

takeoff.

• if either battery temperature reaches 150prior to 132) or 160°F (SN 132 and subsand the red HOT BAT lights are on, themust be turned off, monitored while it coreplaced prior to takeoff.

27. ADC 1, then ADC 2 Pushbuttons ...............................

VMO /MMO Aural Warnings ........................................

Individually test ADC 1 and ADC 2 by depressing tcenter pedestal. These tests allow the various funccomputers and the VMO /MMO aural warnings to be t

the test buttons, the following should be observed as w

• Predetermined values appear on the indicators adata computer:

l i d f

CAUTION

F




28. EFIS Annunciators/GPWS ...............................

Push and hold the PUSH TEST button on top the instrument remote controller, which is locacenter pedestal. This tests the radio altimeter results in display of a height of 100 feet ancomparators for the first 4 seconds. The co

followed by the flashing red illumination ofboth the EADIs and EHSIs. After 10 seconds,steady red for all warnings, except for CAT II,amber TEST light is displayed on the EADI ftest. Release the test button to restore normal o

If the aircraft is in the air, only the radio altimred RA flag on the EADI. The EFIS system teRA test is inhibited after glide slope capture.

The GPWS system is tested by pushing theon either GPWS annunciator panel, located instrument panels. Hold the button in for ththe PULL-UP light goes out. The system aural warnings.

29. Altimeters and ASEL .......................................

Four altimeters must be set to the local statiopilot and copilot altimeters, the standby altimcontroller altimeter.

The ASEL box, located in each EADI, is seadjusting the ALT SEL knob on the instrumenon the aft portion of the center pedestal. Toaltitude, pull up on the knob before turning it t1,000-foot increments in altitude, push down o


F



31. Radios/E. Bat 2 ...........................................................

Any radios that have not already been turned on byswitches should be turned on at this time.

Some aircraft have an option including an emergenadditional electrical power capability in the eve

normal electrical power. The switch to turn on this located on the middle portion of the center pedethrottle quadrant.

32. VHF 1 .........................................................................

33. and Lights...................

The red CABIN and amber REAR DOORS lights, lwarning panel, should be out prior to taxiing the airrecommended that the doors be closed prior to starCABIN light is wired to the main cabin door (two minstalled, the forward lavatory door (one proximity sw

34. Seats and Rudder Pedals ........................................ A

STARTING ENGINES

STARTING PROBLEMS

Discontinue a start whenever any one of the following c

• The ITT does not rise within 10 seconds after moto idle.

• Oil pressure does not rise within 10 seconds afte

REAR DOORSCABIN

F




NOTEPerform a dry motoring whenever fueto have accumulated in the tailpipe.

1. Power Lever .....................................................

This completes motoring circuit interlock requ

2. Start Selector Switch........................................

This allows engine motoring while the staprevents starter engagement after releasing the

3. Start Button....................................... HOLD D

The starter remains engaged as long as the stathe exhaust pipe of fuel.

If the amber IGN light remains on, although N 2 sp and all idle parameters are within l imits:


This allows the operator to manually disengagfailure of the aircraft’s automatic dropout circu

If the ignition l ight goes out:


If the ignition l ight remains on:

2 Associated GEN Switch


F



illuminated when the pilot in command determioperating conditions, it would be in the interest of lights, whether on the ground or in flight.

2. Booster 2 Switch ........................................................

Place the No. 2 boost pump switch to the standby

ensure that the FUEL 2 light on the master warning p

3. DC Power Selector ......................................................

a. Battery Start or APU Assist Start ..........................

b. GPU Start...............................................................

If starting the engines with the batteries or with an DC power selector switch in the NORMAL positiground power unit, place the DC power selectorPOWER position.

4. Start Button............................... DEPRESSED LESS

At 12 to 15% N2 and indication of N1 rotationwithout indication of LP spool (N1) rotation.

5. Power Lever.................................................................

This introduces fuel flow and ignition.

a. 2 Light............................................

6. ITT, N1, Fuel Flow, and Oil Pressure..........................

IGN

IGN

F




7. Idle Speed Parameters ......................................

8. Power Selector Switch .....................................

Subsequent engine starts should be made usiTherefore, at this point, the DC power NORMAL position.

9. Ground Power Unit (As Applicable) ...............

Light .........................................

Removing ground power from the groundgenerator connection to the respective main bu

Engine 3 and 1 Start

1. Booster Switch (3 or 1).....................................

Light .........................................

2. GEN 2 Ammeter...............................................

Use same starting procedure as used for eng

Engine Anti-ice ................................................

BEFORE TAXI1. Circuit Breakers...............................................

After the engines are started and stabilized

FUEL

GEN 2


F



3. Battery Amps/Temperature.........................................

Check the battery amperage and temperatures are normprobably show a charge since they assisted in the start

4. Generators Volts/Amps...............................................

Check the bus voltage on each side by observing ththeir output at 28.5 volts. Move each generator amrespective generator positions to read the ampegenerator. The amperage should not exceed 95 togenerator. Normally, the amperage draw for thgenerators is 110 to 125 amps for each generatgenerator provides 125 to 150 amps. After the check

switch to the BAT 1 and BAT 2 positions.

5. Booster 2 Switch.........................................................

Light ....................................................

6. *Isolation Valve Knob...............................................

Light ...................................................

This rotary switch on the pneumatic section of thselected to the closed position and the ISOL light ilclosure of this valve. This confirms the operatiseparates the bleed air from the No. 1 and No. 3 enengine and APU. This valve is then opened and the IS

7. Compass Headings (5)................................................

Check the headings on the EHSIs, the RMIs, and thbe sure they are all the same and that the IRS systems

ISOL

FUEL 2

F




NOTE

On aircraft without transfer valve

the amber lights labeled “XTK 2 OPE2 CLOSED” are not used.

11. Hydraulic System 1 and 2 Pressure and Quanti

Check the hydraulic quantities and pressures t

no loss of hydraulic fluid during activation of t

12. Standby Pump Switch......................................

Place the standby hydraulic pump switch in


Figure NP-2. Warning Pa

F



d. Release the antiskid test button, and after appseconds, the green L and R brake lights will again

e. Release the brakes and the green lights will go ou

14. *Airbrakes ..................................................................

The airbrake handle in the trim well on the center pNo. 2 position, allows the extension of the centerpanels. Check the illumination of the AIRBRAKE configuration panel.

15. *No. 1 Stall System....................................................

Now that the No. 1 hydraulic system is powered by eand 3, the No. 1 stall system is tested to ensure that tbe extended by the No. 1 hydraulic system. A flashinlight will be seen, the aural stall warning will sound,for all three engines will illuminate. The airbrakes wand the AIRBRAKE light will flash.

16. *Airbrakes...................................................................

Move the airbrake handle to zero to extinguish the fand arm the system for deployment as required. Tretracted for takeoff.

17. Flaps/Slats .................................................................

18. No. 1 and No. 2 Stall Systems....................................

After the slats and flaps have extended to the proabove, depress the STALL 1 button. The aural stallthe igniter lights will illuminate and the inboard




21. AP Transfer by (LH FCS and RH FCSPushbuttons on ID-802)....................................

22. COMM/NAV/Radar/IRS ..................................

Before taxiing the aircraft, check that alnavigation radios are on, the radar is in the st

systems are fully aligned with the mode selecNAV position.

23. Flight Recorder .................................................

24. EFIS/MFD Displays .........................................

Select the desired mode(s) for the EHSmultifunction display. Set the heading and coupon the departure course to be flown.

25. All Flags...........................................................

26. COND BATT Switch (SB 125)........................

Operators with this service bulletin must closeto the batteries in the rear compartment before

TAXI

NOTEThe taxi check should be accomplishedthe blocks and when clear of congestpilot must maintain an outside watcduring taxi operations


F



Check the No. 1 brakes simultaneously, and gentlyright brake pedals until the green L and R lights indicate that a hydraulic pressure of at least 225 psi ibrakes. A slight feedback of pressure should alsopedals. Maximum pressure delivered to the brakes is

• 1,595 psi for aircraft SNs 1, 75 and subsequent, a

BA SB F900-42 incorporated

• 2,175 psi for aircraft SNs 2 through 74, exAMD-BA SB F900-42 incorporated

Check the No. 2 brakes individually, and gently deprpedals until each brake pedal is depressed enough to light. Release one brake check before checking the o

illuminates when the hydraulic pressure to the No. 2 psi. A slight feedback of pressure should also be felThe maximum pressure delivered to the brakes bysystem is 1,080 psi.

A general rule to be followed is that whenever a diffselected, the brakes should be checked to ensurworking properly.

4. Brake Selector .............................................................

After the foregoing brake checks have been made, seNo. 1 system with the antiskid on.

5. Thrust Reverser ....................................................... C

With the No. 2 throttle at idle, pull the reverser leveand observe that the amber TRANS light illuminatesthe green DEPLOYED light illuminates. Place the em




illuminate, and a fluctuation in the temperaturmay be observed. Slowly advance the throttl40% N1 to ensure that manual control of the ethrottle to idle, return the computer switch to Afor all engines.

7. Pilot Window....................................................

Slide the left direct-vision window forward anwindow. While holding the window closed alocking lever forward and downward to locEnsure that the green mark on the groovedlocated on the end of the handle is in view.

8. APU Stop Button .............................................

Push the APU STOP button, which simulatcommands APU shutdown by closing the fuel

9. APU Master (N1 Zero).....................................

When the APU low oil pressure light illuminamay be turned off by pushing in on the switch. Tof the APU by removing electrical power from t

10. APU Bleed .......................................................

The APU bleed switch should be turned o

window to preclude any adverse pressure bumair supplied by the APU.

11. Engine Anti-ice/Wing(or Wing BRK) Anti-ice...................................


F



12. Engine Anti-ice...................................... AS REQUIR

If visible moisture is present and the outside air t+10°C, the engine anti-ice system must be switched system must not be used on the ground. Wing anti-until after takeoff when the landing gear is retracted.engine power for takeoff, ensure that the green an

above each engine anti-ice switch have turned gretakeoff roll.

13. Flight Controls............................................................

Actuate all three primary flight controls over thecontrols should be completely free and automaticallyposition when released. Normally, the copilot cheelevator, while the captain checks the rudder.

14. Takeoff Briefing ..........................................................

a. Confirm the V-speeds and N1 to be used for takairspeed bugs are properly set.

b. Discuss the departure with respect to turns, initirequirements for noise and/or obstacles as publbeing used.

c. The takeoff roll should be aborted if any of before V1:

• Fire/overheat warning

• Engine malfunction

• Illumination of the ENG 2 FAIL light




3. Airbrake ...........................................

4. Thrust Reverser ...................... AS N

d. The captain starts the throttles forward, andto takeoff N1 while calling “Power set.”

e. The copilot calls “80 knots.” At this call

his left hand from the nosewheel steering tthe rudder should be effective for directrunway or wind conditions dictate otherwcan be safely used up to rotation speed.

f. The copilot calls “V1” as the charted speedindicator. Any malfunction after V1 wilproblem, with the proper actions and chealtitude is reached. The captain’s intentionsreturn to be requested, if necessary, should

g. The copilot calls “Rotate” at the chartedwill then apply back pressure to the yrotation attitude.

BEFORE TAKEOFF1. Radar (2)/Transponder.....................................

The radar should be tuned and ready, especiconvective weather in the area of the departumust be turned on with the appropriate code se

2. Parking Brake Handle ......................................

Disengage the parking brake handle by push


F



5. Landing Lights............................................................

Place the landing lights on or, in the case of some airnight conditions dictating.

Do not use the landing l ights for more tminutes while on the ground, as damage mayA 45-minute ground cooling period mustserved at the end of 15-minute ground opThere is no time restriction for use of the lights in flight, as they are individually ventiram-air scoops.

6. Start Selector Switches (3) (If Necessary)..................

Select the AIRSTART position for takeoff if the runwwater, snow or slush on it, or if the presence of birds

NOTE

It is advisable to use ignition for all takeolandings, as it may provide an immediate rean engine should an inadvertent flameout ocing these critical phases of flight.

7. Pitot Heat Switches (3) ...............................................

Just prior to beginning takeoff roll, turn on all three obtain anti-icing of the pitot probes, stall warningRosemont probe, and, if installed, the Teledyne AOA

CAUTION

CAUTION




AFTER TAKEOFF1. Landing Gear....................................................

When a positive rate of climb is indicatedvertical speed indicator, and upon the captaplaces the landing gear selector up. Confirm landing gear until the gear is up and the door li

After takeoff from a snow- or slush-covered retraction 15 seconds, provided that obstaclerespected in case of an engine failure. If nebelow 190 knots, cycle the gear up and dowHowever, gear cycling does not apply to aircheating system (SB F900-32).

2. Wing Anti-ice Switch.......................................

If the total air temperature is below +10°C amoisture, place the wing anti-ice switch on after

For aircraft equipped with a brake heating systakeoff from a snow- or slush-covered runwaapplies. Once the landing gear is retracted, ssystem by selecting the WING-BRK position Leave the switch in this position for at least depending on atmospheric conditions, the antiWING or OFF.

3. Flaps-Slats ....................................... CLEAN/

Flaps-slats are retracted at V2 +25 knots,configuration. Upon the captain’s command, tslats incrementally, one notch at a time. Conf


F



7. Fasten Belts/No Smoking Switches ............................

These switches may be turned off at any time the flig

8. Cabin Pressure and Temperature ................................

Check the cabin rate of climb, cabin altitude, and

indicator to confirm a normal pressurization schedairflow through the gaspers. Rotate the temperature desired temperature in the cabin and in the cockpit. Iis a remote temperature control for the cabin, the remselected at this time for passenger convenience.

9. Entrance Curtain..........................................................

It is recommended that the entrance curtain be closedfrom the cold and to reduce the noise level at this loca

After 10,000 feet:

1. Landing Lights ............................................................

The landing lights may be turned off at this time undictates leaving them on until a higher altitude is reac

After 18,000 feet or transit ion level:

1. Altimeters....................................................................

Set the altimeter to QNE to comply with the FARs.

2. Oxygen ........................................................................




CRUISE1. Fuel Management.............................................

Check that all tanks are feeding properly and maintained. If the takeoff was made with a reaching the filed altitude, open the boost pumvalves. The No. 2 fuel boost pump must be selin order that both boost pumps in group 2 tanfrom the group 2 tanks to all three engines. Wshown a decrease of approximately 900 poundtank groups show the same level, return the valves to the closed position. The No. 2 boostnormal position during all phases of flight.

2. Station Check...................................................

The station check should be performed at leaensure that all aircraft systems are operatrecommended that systems readings be kept asystems operations, especially fuel consumptio

DESCENT1. Cabin Pressure Controls ...................................

Set the automatic pressure controller mode landing field elevation in the LAND ELV winaltimeter setting for the field at which the laQNH window of the controller.

2. Fasten Belt .......................................................


F




4. Altimeters............................................................. SE

When passing the transition level in the descent, fouset to the landing airfield’s QNH; the pilot’s and altimeter, the standby altimeter, and the QNH on the ca

5. Landing Computations/Bugs.......................................

Determine the landing weight, and compute the reqdistance/field length, VREF, and go-around for that on the EADI by rotating the IAS knob on the DC-8on the center pedestal. This VREF setting should additives regarding the final approach configuratiois, for flap settings less than 40°, if the airbrakextended position or if there is a flight control pairspeed additive (Table NP-1).

When landing with any wind, or especially if gusty at the landing airport, add half the steady wind comgust component, not to exceed a 20-knot additive. Dobug setting to incorporate wind, but carry the windspeed above VREF. One of the analog airspeed indic

on this target speed for final approach to touchdowairspeed indicator bugs may be set on VFR and 1.43of a go-around.

Wind additives are to be carried as an additive all the


Table NP-1. LANDING COMPUTATIO

WT/1,000 LB 24 26 28 30 32 34 36

VREF 100 104 108 112 115 119 122 1




b. Discuss the approach to be used:

(1) Type and runway direction

(2) Approach frequency and identification

(3) Airport elevation

(4) Minimum safe altitude

(5) Any transition altitude

(6) The inbound magnetic course

(7) The final approach fix altitude

(8) All missed-approach information

(9) The final approach speed to be used

(10) The missed approach point, timing, D

(11) All added information to clarify the ap

(12) All lighting that is available

(13) All runway information: length, wtouchdown runway remaining, etc.

c. The pilot not flying will make calls on fiwith the following:

(1) One dot left or right—“Localizer”

(2) One dot above or any below—“Glide


F




(10) At minimums:

• With runway in sight—“Minimums-land

• With no runway in sight—“Minimums-

Below 10,000 feet:

1. Landing Lights............................................................

Turn the landing lights on or to pulse (those so equipseen by other aircraft.

APPROACH

1. Entrance Curtain .........................................................

As soon as practical prior to entering the approaentrance curtain, as it is an FAA requirement that clear of obstructions prior to landing.

2. No Smoking Sign........................................................

Place the no smoking sign on prior to landing, as it isthat all smoking materials be extinguished before lasignal to the passengers that landing is imminent.

3. Altimeters/Radio Altimeters ................................ SE

Once again, confirm all altimeters are set to the curensure proper altimeter readings at the DA or MDA.HAT/HAA on the radio altimeter for the approach tsetting is for backup information only and not to be d i i l i d i i d l i d I





Whenever moving the flap-slat handle iit one increment at a time to ensure thement to the position selected.

BEFORE LANDING1. Landing Gear ...................................................

Select the landing gear handle to the downoperation of the landing gear doors and extensia positive effort to watch the gear until all threthe proper three green arrow light indication

should be out as long as the gear was lowered in

2. Antiskid ............................................................

The brake selector switch must be in the #1landing gear control handle must be in the dhold the brake pedals. The green L or R brake

While holding the brake pedals depressed, morelease the antiskid test button. The green illuminate about one second after depressingthen extinguish about one second later. Releabrakes after the green lights extinguish. The lig

3. Hydraulics.........................................................

Check that the hydraulic pressure and quantityfinal activation of all the hydraulic componenlanding. This is a final check to ensure that yavailable for activation of airbrakes, brakes, anl di ll

CAUTION




2. Anti-ice—Wing (or Wing-BRK)......................

Wing anti-icing must be turned off as soon icing conditions still exist after landing, enginuntil parking the aircraft.

If wing anti-icing is inadvertently lefting, do not retract the slats until the winhas cooled sufficiently.

3. Pitot Heat Switches (3).....................................

In sequence with the preceding items, it is ne

pitot heat switches as soon as landing roloverheating of the pitot-static components and e

4. Start Selector Switches (3)...............................

Lights (3) ...................................

If the igniters were placed to airstart for the should be placed to the ground start position du

5. Anticollision Lights..........................................

Move the three-position anticollision light swhich turns off the white wingtip strobe lights

illuminated until the engines are shut down at t

6. Landing Lights .................................................

If h l di li h d d f i

IGN

CAUTION



PARKING1. Park Brake/No. 2 Brake Light..........................

2. Flight Data (Fuel-IRS) .....................................

3. AVIONICS/FMS/EFIS Masters .......................

For those aircraft equipped with one, two, orthem off at this time. It is recommended thnavigation radios be turned off at their respecsave the LED from early failure when AVIONand off during ground operations.

4. Engine Anti-ice Switches (3) ...........................

Prior to shutting down the engines, the engineturned off and the engines stabilized.

5. Taxi Light.........................................................

6. IRS (2 or 3).......................................................

7. Radar (2)/Transponder .....................................

8. Standby Horizon ..............................................

Pull the knob on the emergency horizon, andthe mechanism.

9. Standby Pump ..................................................

F




14. Fasten Belt/No Smoking/Emergency Lights...............

Turn the cabin warning and emergency exit lights electrical power from the aircraft. Although this doewarning lights, it does affect the emergency lights, when power is removed from the main buses.

15. Interior Lights..............................................................

Turn off all extraneous cabin lights to prevent excessbatteries when power is again applied to the aircraft.

16. APU/No. 2 Booster Pump ...........................................

Depress the stop button on the APU to shut it down, booster pump, as it is no longer needed. Turn the Awhen the low oil pressure light comes on.

17. Batteries (2).................................................................

As soon as possible after supplementary electrica

from the aircraft, turn off both batteries to prevent battery power.

18. Engine Computers (Last Flight).................................

19. Generator Switches (Last Flight) ................................

20. Aircraft Chocked ........................................................

The park brake should not be left on for extended paproper chocking before release of parking brake.




1. Engine anti-ice systems (ENG ANTI-ICE)flight or on ground when icing conditions exiduring climb and cruise when the temperatureTAT more than +10°C (50°F).

However, flying in vicinity or through a “cumrapid variation of SAT with SAT increasing anticipate icing conditions by selecting the ant

Do not rely on airframe visual cues to turn antemperature and visible moisture criteria speci

Conclusion:

During climb and cruise, the pneumatic anti-ic

• Below +10°C (50°F) TAT and above –40°C

• In visible moisture

If both of these conditions are not met, the anti

2. Wing anti-ice system (WING ANTI-ICE) shoprior to entering visible moisture whenever th

3. Encounter with icing conditions is evidencedthe non anti-iced area around the windshioperation, lights switched on by WING (EXilluminate the wing leading edges to allow the

4. Comply with engine and wing anti-ice systemminimum N1 speed values.

5. If necessary during the approach , increase thethe airbrakes to help keep N1 speed to no less t

The N1 speed of the operative engines must not be lues as shown in table below.

Three Engines Operative



When the aircraft has been left on the ground for in ambient temperatures of 5 °F ( -15°C) or belowporating a heating network must be heated as follo

• Place both WINSHIELD PILOT/COPILOposition (medium heating )

• Also switch on the side window heating sw

• Keep the heating ON for 15 minutes befor

COLD WEATHER OPERATION

Fuel Selection

Prior to prolonged parking in very low temperatureplacement if necessary) that the freezing point of the anticipated minimum ambient temperature.

In-flight tank fuel temperature must be maintained aing point of fuel being used. I f necessary, increasealtitude to raise the total air temperature.

This should be achieved if the total air temperature isbelow the fuel freezing point .

As water may freeze as it settles out of the fuel whthe cold, draining must not be carried out until thon heated areas of the airfield or in a heated hanga

While moisture in the fuel is not exclusively a coldfrequently cause trouble during engine starts in be

Water Servicing

On ground when cold temperatures are expected wa

F




SEVERE TURBULENCE PENETRATIOFlights in severe turbulence should be avoided wh

If necessary, reduce the speed to 280 KIAS max oseat belts, decrease altitude to increase buffet bou

NOTE

Autopilot or yaw damper operation is

With the autopilot disengaged:• Fly attitude• Avoid using the stabilizer trim• Do not chase altitude and spee

WINDSHEAR SITUATION

If windshear is anticipated:

• Do not take off—Wait

• Do not land—Wait or fly to an alternate ai

Pilot reaction time must be very low—3 to 5 secondonds or more to understand the situation, it will be

If a windshear encounter is imminently anticipate

1. Go around pushbutton......................................

2. Level the wings

3. Pull up to stall warning onset

4. Power levers .....................................................

5. AIRBRAKES handle........................................



Wing Anti-ice Inoperative Without Brake HWing Anti-ice Unwanted Operation

Without Brake Heating ...............................

Wing Anti-ice Inoperative With Brake Heat

Wing Anti-ice Unwanted Operation With Br

Engine Anti-ice Inoperative........................Engine Anti-ice Overpressure.....................

Engine Anti-ice Unwanted Operation.........

Late Activation of Systems.........................

AIR CONDITIONING.........................................

Bleed-Air System Overheat........................

ECU Overheat..............................................

Cabin Air-Conditioning Unit Overheat (OnTurbocooling Unit Anti-icing Emergency C

Cabin Air Conditioning Overheat On Aircra

Equipped With Anti-icing Emergency Cont

Battery Conditioning Failure (SB-125) .......

Nose Cone Overheat....................................

PRESSURIZATION..............................................

Improper Cabin Vertical Speed ...................

Too High Cabin Pressure.............................

Too High Cabin Altitude or Slow Depressu

Door Unlocked Indication




FIRE PROTECTION SYSTEMS...................................Failure of Engine Detection System .....................

Failure of APU Fire Protection System ................

EFIS ..............................................................................

Either EADI CRT Failure .....................................

Either EHSI CRT Failure......................................

Simultaneous Failure of EADI and EHSICRTs on the Same Side ........................................

Successive Failure of EADI and EHSICRTs on the Same Side ........................................

Loss of ASCB Control..........................................Invalid Attitude and/or Heading Data ...................

IRS Attitude Comparison Annunciation withor without a Heading Comparison Annunciation.

IRS Heading Comparison Annunciation without

an Attitude Comparison Annunciation.................IAS/Mach Data Invalid.........................................

IAS Disparity Annunciation .................................

LOC or GS Disparity Annunciation .....................

Failure of Either FMS...........................................




ILLUSTRATIONFigure Title

AP-1 Inflight Relight Envelope................

TABLESTable Title

AP-1 Landing Data ...................................

AP-2 Pitot-Static System—Either Air

Data Computer Inoperative .............

AP-3 Pitot-Static System—Both Air

Data Computers Inoperative ..........

AP-4 Flight In Icing Conditions ...............




ABNORMAL PROCE

NOTE

This abnormal procedures checklist itraining purposes only. Where checklidiffer from the Airplane Flight Manual

Flight Manual takes precedence.

LANDING DISTANCE ADDITIONSThe following are additives to be made to the landia 40° flaps + slats landing under normal circumsta

NOTE

LD—Landing distance

EMERGENCY CHECKLIST

Loss of Both Hydraulic SystemsClean configuration, VREF + 30 knots. Twice theLanding field length is not addressed.

Approach and Landing—Two Engines

If using No. 2 brakes, add 50% to both the LD.

After making the additive for use of No. 2 brakes a

• 7° flaps + slats, fly VREF + 20 knots; add 8




Loss of No. 1 Hydraulic SystemAdd 60% to the LD.

Loss of No. 2 Hydraulic System

The following additions must be made to the landing ditations.

• 0° flaps + slats, fly VREF + 20 knots; add 800 fee



• 40° flaps + s la ts , land a t VREF.

After making the additions for flaps and slats add 10% tbrakes.

Inoperative Stabilizer

Use 20° flaps + slats, and fly VREF + 20 knots; add 800 f

Inoperative Elevator

Use 40° flaps + slats, and fly VREF + 10 knots; add 1,800

PITCH FEEL Light On (Arthur UnitFailed in Heavy Force Position)

Fly VREF + 10 knots; add 800 feet to LD.




Landing with Slats or Flaps MalfunctioIf using:

• 40° flaps + outboard slats only, fly VREF +LD.

• Flaps/slats clean, fly VREF + 30 knots; add

Landing with Airbrake Malfunct

If :

• Airbrakes are extended to position 1, fly VRto LD.

• Airbrakes are extended to position 2, fly VRto LD.

• Airbrakes do not extend, add 10% to the L

Landing with Antiskid Inoperative

Use brakes in #2 A/SKID OFF selected; add 50%

Landing with Parking Brake Only

Add 50% to the LD.

ENGINES

NOTE




ENGINE FAILURE BEFORE V1Reject the takeoff:

1. Brakes................................................................. MA

Use the brakes to their maximum pressure, deprunway, during any aborted takeoff. The brakes, in

antiskid on, are the most important factor in stoppinshould be applied as other actions are taken to decele

2. Power Levers ...............................................................

Immediately upon the abort call, retard the power lbring the power levers to cutoff; otherwise, hydraulic

the aircraft may be lost.

3. Airbrake Handle ..........................................................

Ensure that the airbrake handle is placed to positairbrakes decreases lift on the wings and makes braWhile on the ground, if the failure was the No. 2

hydraulic pump will automatically operate when thethe No. 2 system drops to approximately 1,500 psi.

4. Thrust Reverser...........................................................

Use the thrust reverser as an added benefit to stoBefore ordering reverser deployment, ensure that the

the ground on all three gears, the airbrakes are extendis effectively at idle power. The thrust reverser can taxi speed is achieved without risk of reingestion of stalls. The thrust reverser is most effective, howeverduring deceleration. It should be used to slow the



7. ST–BY Hydraulic Pump Switch (If No. 2 Engine Fai

If the No. 2 engine has failed, there may not be enougoperate the engine-driven hydraulic pump. In order tNo. 2 hydraulic system components, position the staswitch to ON. In flight, with the standby pump switch pump will operate only when the airbrakes are selected

8. At V2 + 25 knots—FLAPS–SLATS Handle ...............

When an airspeed of V2 + 25 knots is achieved, regconfiguration, retract the slats and flaps incremposition. Check that the flap position indicator on this at 0°.

9. Enroute Climb Speed..................................................

Compute the enroute climb speed from the performanis 1.43 VS in the clean configuration. It is also the beif on one engine, the maximum lift over drag speed, use for driftdown if all engines have failed and canno

10. Failed Engine..............................................................

Analyze all engine indications. Attempt an airstartunless a greater emergency exists with the engine. Bopositive and confirmed identification of the affectefurther actions are taken. Shut down the inoperativeFailure In Flight,” this chapter).

Five minutes after brake release:

11. Crew and Passenger Bleed Switches ..........................




NOTE

An attempt airstart may be tried on the(see “Airstart” section, this chapter). Iunsuccessful, complete the engine shudure (see “Engine Failure In Flight” schapter).

ENGINE FAILURE IN FLIGHTThe OIL, GEN, and possibly PUMP lights illumin

Yawing tendency (failure of No. 1 or No. 3 engine

Determine which engine has failed. Analyze all enan airstart on the failed engine unless a greater eme

gine. Both pilots must make a positive and confirmfected engine before any further actions are taken

Engine Shutdown

1. Power Lever..........................................RETARONE MINUT

If the engine is still running, and a precautnecessary, if possible, allow the engine to cbefore shutdown.

2. Power Lever......................................................

After positive identification of the proper engthe affected engine power lever to cutoff.

3. Booster Switch .................................................




6. Fuel Shutoff Switch (If Engine Cannot Be Restarted)

If the engine cannot be restarted, switch tshutoff valve off. In icing conditions, operaengine anti-icing even with the No. 2 engidown. The isolation valve must be open.

This will allow bleed air from the bleed-air mto anti-ice the S-duct. If the No. 1 or No. 3 eshut down, turn off the engine anti-icing switcwhen in icing conditions.

If Engine Shutdown procedure was initiated due to engineoff (after V1) 5 minutes after brake release and if obstacle

7. Crew and Passenger Bleed-Air Switches....................

8. Maximum Continuous Thrust .....................................

If required, see One Engine Inoperative Drift Down cha

NOTE

After an unsuccessful airstart attempt of eng3, select bus-tied switch to FLIGHT NORcheck that the BUS TIED light is out.

If the No. 2 engine is shut down:

8. Bus-Tied Switch ..........................................................

Since the No. 2 generator is no longer supplying elright DC electrical buses, tie the buses to save the

CAUTION

i h




10. Booster 2 Switch ..............................................

This will allow normal booster pump operatibooster pump for automatic operation oncrossfeed switches are selected to crossfeed.

11. X–BP 1–2 and/or 2–3.......................................

a. X–BP light(s) checked................................

If tank 2 level is higher:

b. BOOSTER 1 or 3........................................

In this case, BOOSTER 2 also supplies the

If tank 2 level is lower:

c. BOOSTER 2 ...............................................

In this case, BOOSTER 1 or 3 also supplies

ONE ENGINE INOPERATIVE—APPROA1. Hydraulic Implications .....................................

Depending on which engine is shut down andaffected, the hydraulic implication and specchecked, evaluated, and applied before beginni

2. Electrical Implications......................................

Depending on which engine is shut down andaffected electrical implications and bus load

5 EGPWS FLAPS O’RIDE S it h




5. EGPWS FLAPS O’RIDE Switch...............................

The GPWS FLAPS O’RIDE switch prevents the TOOwarning from sounding when a less than 40° flap land

6. Crew Briefing .............................................................

The crew should be fully briefed on all plans and approach and landing to be flown. Special considerathe special requirements dictated by the emergency that caused the engine to be shut down. The creapproach configuration. The approach speed andadjusted accordingly. With 20° flaps + slats, increaseVREF + 5 knots. When 7° flaps + slats, increase the L+ 15 knots.

7. Approach Checklist.....................................................

a. Passenger Door Curtain.................................

Open the passenger door curtain to permit use of

an emergency exit if necessary, and also ensurhave attached their seat belts correctly and that thin the required position for landing.

b. No Smoking Sign...........................................

c. Altimeters ..................................................

Set QNH or QFE as instructed by local air trafonly be selected if the pressure altitude of the dinside the altimeter setting limits.

If the landing lights incorporate a p lse




If the landing lights incorporate a pulse

position is selected.

g. Approach with 20° flaps + slats:

• Flap/Slat Handle.............................

At the usual point on approach, as wapproach flown with three engines, seflaps + slats incrementally.

• Landing Gear Control ................

At the usual point on approach, as wapproach flown with three engines, pla

• Airspeed .....................................

Fly the final approach at the normal 4plus 5 knots. Increase the landing standard correction for wind must be a

OR

g. Approach with 7° flaps + slats:

• Flap/Slat Handle .........................

At the usual point on approach, as wapproach flown with three engines, sflaps + slats.

• Landing Gear Control ................

At the usual point on approach, as w

Check for normal operating pressure of approxim




Check for normal operating pressure of approxim

No. 2 engine is inoperative, required hydraulic phydraulic system operating components can bstandby hydraulic pump (1,500 to 2,150 psi).

c. Anti-Skid ....................................................

Ensure that the braking selector switch is selecON position. Depress the brake pedals and cheNORM brake lights remain out. Depress the test pthat the L and R NORM brake lights come on afteout one second later and remain out after releasing

d. Airbrake Handle .........................................

NOTE

Airbrakes may be used during landing approavided airspeed is at least VREF + 10 knots. Ilanding distance by 15%.

e. Start Selector Switches (If Necessary) .......

Check for appropriate IGN lights on.

f. CAT 2 Mode Armed ...................................

Depress the CAT 2 pushbutton on the AP controthe CAT 2 message appears on the ID 802.

g. FLAP + SLAT HANDLE .........................SL

NOTE

If l di ith th N 2 i i ti

It is imperative that the landing climb attitude




It is imperative that the landing climb attitude

descent is stopped and the climbout is begun.with the advancement of the power levers to ta

3. Airbrake Handle...............................................

If the airbrakes were extended for some reasomust be retracted as power is applied and th

landing climb attitude.

After completing the above checklist items, followdepending on the flap configuration flown on the

Procedure A—If the Approach

was Flown with 20° Flaps + Slats1. Flap/Slat Handle ..............................................

Ensure that the FLAPS–SLATS handle is set togo-around.

2. Landing Gear Control.......................................

As soon as a climb is indicated on both the altindicator, raise the landing gear.

3. Airspeed to 400 Feet AGL ...............................

The best climb speed to maintain to a minimuground level is VREF + 5 knots. If a higher srotation, maintain that speed to 400 feet AGL. D

3 Airspeed to 400 Feet AGL




3. Airspeed to 400 Feet AGL ..........................................

The best climb speed to maintain to a minimum altituground level is VREF + 15 knots. If a higher speed irotation, maintain that speed to 400 feet AGL. Do not o

For all go-around configurations, when reaching 400 feet Aaltitude:

4. Level Flight Acceleration ...........................................

At an altitude no lower than 400 feet above groundflight while accelerating to the slats-flaps retract spee

5. At V2 + 25 Knots—FLAPS–SLATS ...........................

This is the normal retraction speed for the FLAPS–VREF as the base speed, which is a conservative speweight of the aircraft, VREF speed is 7 knots higher th20° flaps + slats takeoff. Moreover, for a given gross wis equal to the V2 speed for a 7° flaps + slats takeofonly one VREF for the aircraft at a given gross weight.

normal landing configuration, which is 40° flaps + slat

6. Enroute Climb Speed..................................................

Once the slats-flaps are retracted, accelerate to the e1.43 VS. Reduce engine thrust to the maximum contin

Table AP-1. LANDING DATA

LANDING DATA 40° FLAPS + SLATSSTANDARD TEMPERATURE UNCORRE

LANDING DISTANCE/LANDING FIELD LENGTG W

AIRSTART




AIRSTART

General

Do not attempt to relight an engine afire if the engine integrity is questionrotation is not observed (Figure AP-1

Wait ten seconds between two consecuttempts. Do not make more than threairstart attempts.

CAUTION

WARNING

30

20

10

0

ALTITUDE (X 1,000 FT)

MAXIMUM AIRSTART ALTITUDE IN MANUAL MO

The best technique is to apply this procedu




q pp y p

timely fashion, as soon as it has been ascertaithe engine malfunction that caused the flamenot present a danger if an airstart attempt is mapower lever must be immediately set to idleairstart ignition selected in order to take advathe high rotational speed of the compressors

If the relight attempt is unsuccessful, it is neto descend to an altitude which is within theairstart envelope.

1. Power Lever ..................................................... IMME

The power lever must be immediately retarded to input of fuel for an immediate relight, which might re

2. Start Selector Switch ..................................................

As soon as the power lever is retarded to idle, plaswitch to AIR START. The above two actions mbefore the N2 rpm decays to 15% or less to increasesuccessful start. This provides ignition for engine

light on.

3. ITT Rise within 10 Seconds .......................................

This is the indication of a successful relight of the enrises normally, being careful the temperatures stay wenvelope. All other engine instruments must be che

proper indications.

4. Power Lever................................................................

If an airstart is unsuccessful, complete th




p

procedure in this manual.

Abnormal Airstart

Abort an airstart whenever any one of the followin

• The ITT does not rise within 10 seconds aft

to idle.

• The oil pressure does not rise within 10 se

• The ITT rises rapidly and approaches the 978°C (TFE-731-5BR) limit.

• An N1 remains close to zero when N2 spee

• N2 speed is not rising rapidly and smoothl

• If, during an airstart with the computer inceeds 80% with the power lever at idle.

If any of the above are observed during the airstarfollows:

1. Power Lever......................................................

This is done to shut off the fuel at the engine fu

2. Start Selector Switch ........................................

This action disengages the starter if it was uattempt.

3 Complete the Engine Shutdown checklist

3. Fuel Shutoff Switch.....................................................




If the fuel shutoff switch was placed off at someshutdown, place the switch, located on the fire panelthe fuel supply from the fuel tanks to the engine.

4. GEN Switch ................................................................

Place the generator switch on, to provide a start interassisted airstart. This will also configure the gimmediate electrical power to the buses when the eng

5. Engine Computer Switch............................................

The engine computer switch, unless the computer is icause of the engine shutdown in the first place, shoulCheck for CMPTR light out. If the computer is inAirstart—Abnormal—Computer Off checklist, procedure.

6. Booster Switch............................................................

Place the booster pump switch on, checking that thgoes out, to supply fuel under pressure from the tank

7. Engine and Wing Anti-ice Switches ...........................

Place the anti-ice switches off to prevent any unwa

air from the engine being started. The unwantedmight impede the start and/or cause an overtempebeing started.

Select either procedure A or B below, based on N




than zero.

Abort airstart when anyone of the abnormal conditiAirstart,” this chapter).

Procedure A—Windmilling Airstart(N

2more than 15% and Indication of N

1Ro

NOTE

If N2 speed is greater than 15%, a starnot be required.


The start selector switch will provide ignitionAIR START in flight. Check IGNITION light

2. Power Lever......................................................

Move the power lever from cutoff to idle. Tsupplied through the engine fuel control to t

engine.

3. ITT Rise within 10 Seconds .............................

This is the indication of a successful relight oITT rises normally, being careful the temperaother engine instruments must be checke

indications.

4. Engine Acceleration to Idle..............................

If the ignition light remains on after the engine is on spe




7. Start Selector Switch ............................................ MO

This action disengages the starter if it was used toattempt. Check that the ignition light on the overheis out.


After the airstart attempt has been aborted, move theback to its normal position of GRD START.

OR

If the BUS TIED light stays on after the bus-tied sFLIGHT NORM:

This means the buses are still tied together. Attempt motied switch in the opposite direction. If the BUS TIED l

7. Generators Volts and Amps ........................................

Closely monitor the buses for proper power suppbuses are no longer isolated in case of generacomponent malfunctions.

Procedure B—Starter-Assisted Airstart—N2 les


The start selector switch will provide ignition for reAIR START in flight.

4. ITT Rise within 10 Seconds .............................




This is the indication of a successful relight ofITT rises normally, being careful that the temp

5. N1, Fuel Flow, and Oil Pressure.......................

All other engine instruments must be checkindications.

When N2 is above 50%:


Place the start selector switch from AIRSTART

the ignition circuit.

IGN–GEN–OIL–PUMP Lights........................

These lights go out when the normal starting se

7. Engine Instruments ...........................................

Continually monitor the engine instruments for

8. Bus-Tied Switch...............................................

Place the bus-tied switch to FLIGHT NORM

buses to an isolated operation, the normal infliTIED light should go out.

If the ignition light remains on after the engine is

OR




If the BUS TIED light stays on after the bus-tied sFLIGHT NORM:

This means the buses are still tied together. Attembus-tied switch in the opposite direction.

If the BUS TIED light is still on:

9. Generators Volts and Amps ........................................

Closely monitor the buses for proper power supply anare no longer isolated in case of generator, battermalfunctions.

Airstart—Computer in MANUAL Mode

Preparation Phase

1. Airstart Envelope ........................................................

If an immediate relight was not possible, establish t

airstart envelope as prescribed in the chart (see Figure

2. Power Lever.................................................................

The throttle must be placed in cutoff to prepare the en

3. Fuel Shutoff Switch.....................................................

If the fuel shutoff switch was placed off at someshutdown, place the switch, located on the fire panelthe fuel supply from the fuel tanks to the engine

6. Booster Switch.................................................




Place the booster pump switch on, checkinggoes out, to supply fuel under pressure from th

7. Engine and Wing Anti-ice Switches ................

Place the anti-ice switches off to prevent any ufrom the engine being started. The unwantedimpede the start and/or cause an overtempestarted.

8. Bus-Tied Switch ...............................................

The DC electrical buses must be tied toget

assisted airstart is necessary. This, along withprovides the proper start interlock in case thillumination of bus-tied light.

Abort airstart when any one of the abnormal conditAirstart,” this chapter).

Windmilling

After preparing the engine for the computer in mamilling airstart procedure as used for the computeHowever, the N2 speed must be greater than 15%,greater than 10%.

Start-Assisted Airstart—N2 less than 1


Move the power lever from cutoff to idle. This act

supplied through the engine fuel control to the inje




supplied through the engine fuel control to the injeengine.

4. ITT Rise within 10 Seconds........................................

This is the indication of a successful relight of the enITT rises normally, being especially careful the tem

limits since the computer is in manual mode.

NOTE

If any abnormal situations occur as describebeginning of the airstart checklist, or if the N180% with the power lever at idle, abort the s

When N2 is above 50%:

5. Start Selector Switch (As Required)..................... MO

Move the start selector switch to MOTOR–START Sstarter function of the starter-generator, which caucome online to power its bus and stops the ignition to

IGN–GEN–OIL–PUMP Lights...................................

These lights go out when the starting sequence is com

6. Engine Instruments.....................................................

Continually monitor the engine instruments for noensure that the CMPTR light remains on.

7 S S l S i h

This action disengages the starter if it was u

attempt




attempt.

Ignition Light....................................................

Check that the ignition light on the overhead e


After the airstart attempt has been aborted, mback to its normal position of GRD START.

OR

If the BUS TIED light stays on after the bus-

FLIGHT NORM:

This means the buses are still tied together. bus-tied switch in the opposite direction.

If the BUS TIED light is still on:

9. Generators Volts and Amps .............................

Closely monitor the buses for proper power suare no longer isolated in case of generator, malfunctions.

FUEL CONTROL COMPUTER INOPER

CMPTR

2. Engine Computer Switch ............................................




Cycle the computer switch to MAN, to OFF, and banormal circuitry can be restored. Monitor the engguard the power lever while cycling the computer sw

If the CMPTR light stays on:

3. Engine Computer Switch ............................................

If the attempt to reset the computer fails, set the compand operate in that position for the remainder of control circuits will be lost, except for N1 and N2necessitating close scrutiny of engine instruments and

NOTEDo not let the ITT indications of the affectedexceed the operating ITT of the other engin

Avoid rapid displacements of the power levsurge bleed valve cannot open completely, anpressor stalls might result.

Maximum thrust may not be attainable.

Idle thrust may be higher than normal. Thisbe taken into consideration, when landing

the increase in residual power of the engine

The fuel flow on the affected engine mayproximately 5% higher when its N1 speed is mt th N d f th th i

CAUTION

If the indicated oil pressure is greater than 25 psi:

ll i i f h li h i d b




Illumination of the OIL light is caused by tdetection of a chip in the oil may indicate engicase it would be prudent to reduce thrust on potential of further damage at high operatioContinue to monitor the oil pressure and tempeflight if the engine is kept running.

2. Engine Thrust (If Possible)...............................

If the indicated oil pressure is less than 25 psi

3. Retard the power lever and shut down the possible.

4. Complete engine shutdown by check list .

Engine lubrication is no longer available anmay result.

In icing conditions, operate the No. 2 eeven with the No. 2 engine shut down.valve must be open.

If the engine 2 is shut down, the bus B remains supa limited period. The buses A and B have to be tied anilluminated.

NO. 2 ENGINE INLET DOOR OPEN

ENG 2 FAIL

CAUTION

If the engine surges or if any abnormal engine instrum

observed:




observed:

2. Complete the engine shutdown by checklist in thFlight” section in this chapter.

TAKEOFF CONFIGURATION

This is a ground warning and is activated when at least onis advanced above a position of 82° to 84° or greater andfollowing conditions have not been met:

1. Airbrakes.....................................................................

The airbrakes must be fully retracted and the airbrathe zero position.

2. Slats ............................................................................

The flap/slat handle must be placed to the desired conand both inboard slats are not deployed to the fully ex

3. Flaps ...........................................................................

The trailing-edge flaps must not be extended beyond

4. Stabilizer Trim.............................................................

The horizontal stabilizer trim must be in the green ra4 5° d 7 5° h i i di

+ AURAL WARNING “NO TAKEOFF”T/OCONFIG

HYDRAULIC SYSTEMS




LOSS OF NO. 1 SYSTEM

1. Hydraulic Pressure and Quantity......................

A loss of the No. 1 system may also be indicathe hydraulic pressure gage. The hydraulic fzero.

2. Airspeed ..................................................... 26

This is the maximum speed for flight in the epressure is lost to the flight controls.

3. New Bug Speed................................................

Set the appropriate bugs to VREF plus additiv1.43, respectfully.

4. Flaps + Slats Handle (On Approach)...............

Ensure that the flaps + slats handle is sele

approach. This action will provide electrical cprevent any flap operation until the outboard sfunction of the emergency slat switch. Checkslat light.

AND POSSIBLY

PUMP 3

PITCH

FEEL

PUMP 1

monitoring the deceleration rate on the EADI, optim

dry runway being between 0.25 g and 0.30 g, increasi




y y g g g,

7. Landing Gear (At Appropriate Time) .........................

Complete the “Landing Gear—Emergency Extfollowing the “Airbrakes Do Not Extend In Flightlater in this chapter.

System Status

Operative Systems

• Servoactuators barrel 2

• Flaps

• Airbrakes

• Emergency slats (add 5 knots to VREF)

• Aileron Arthur Q

• No. 2 brakes

• Nosewheel s teering

• Parking brake

• Thrust reverser

Inoperative Systems• Servoactuators barrel 1

• Pitch Arthur unit

FAILURE OF A NO. 1 SYSTEM PUMP

PUMP 1




This indicates failure of only one of the No. 1 hyd

1. Hydraulic Pressure and Quantity .....................

Check both hydraulic system gages to ensure pThe other engine-driven hydraulic pump lightpressure and quantity indicated.

If the pressure and quantity are normal, expect long1 system components.

LOSS OF NO. 2 SYSTEM

A loss of the No. 2 system may also be indicated bhydraulic pressure gage. The hydraulic fluid quantiprocedure differs from the No. 1 hydraulic system drives the No. 2 system. Illumination of the PUMPthe loss of the No. 2 engine-driven pump. The standbable if it was the engine pump that failed, provided hy

Depending on airspeed, the AIL FEEL light may i

1. No. 2 Hydraulic Pressure and Quantity...........

AND POSSIBLYPUMP 2AIL

FEEL

PUMP 3

PUMP 1

3. ST–BY Hydraulic Pump Switch .................................

If the hydra lic q antity is ero t rn off the p mp to




If the hydraulic quantity is zero, turn off the pump toand possible damage. If the quantity is normal, use tonly for those key phases of flight.

4. New Bug Speed..........................................................

Set the airspeed bug to reflect the approach beinbenefit of flaps.

5. EGPWS FLAPS O’RIDE Switch...............................

This switch prevents the TOO LOW FLAPS audio w

if a less than 40° flap landing is made.

System Status

Operative Systems

• Servoactuators barrel 1

• Normal slats

• No. 1 brakes (ant iskid)

• Landing Gear (normal and emergency control sy

• Pitch Arthur unit

• Thrust reverser (available if accumulator is char

• Parking brake (available if accumulator is charg

NOTE

If the No. 2 hydraulic system is lost fofollowing additives must be made to th




following additives must be made to thtance and landing field length comput

• 0° flaps + slats, fly VREF + 20 knotsto LD and 1,335 feet to LFL.

• 7° flaps + slats, fly VREF + 15 knotsto LD and 1,000 feet to LFL.

• 20° flaps + slats, fly VREF + 5 knotsto LD and 335 feet to LFL.

• 40° flaps + slats, land at VREF

• After making the additive for flap10% to both the LD and LFL for n

UNWANTED OPERATION OF STANDB

On the ground, the ST BY PUMP light may indicatein the rear compartment is out of the IN-FLT deten

Before correcting the situation by modle to the IN-FLT position (No. 2 hydrensure that all hydraulic pressure is rboth the No. 1 and No. 2 hydraulic sy

h d li i h

CAUTION

ST BY

PUMP

FLIGHT CONTROLS

LANDING WITH INOPERATIVE STABILIZER




LANDING WITH INOPERATIVE STABILIZER

This condition is indicated by the inability to trim the hby either the normal or emergency trim system. The horizremain in the last position selected prior to the trim failucolumn and then perform the following procedures:

1. Autopilot .....................................................................

If the autopilot is engaged, and if it hasn’t already bto the loss of normal trim capability, disengage it fromby depressing the yaw damper pushbutton on the action will disengage the autopilot and yaw damper ato come on and show a disengagement message on th

NOTE

The failure of the horizontal stabilizer causeevator Arthur unit to lock in the position it is time of failure. When the flaps and slats are exthe Arthur unit returns to the low-speed pwhich results in a significant, but very gradcrease of elevator feel force.

If the stabilizer is jammed in the +2 to –4° range:

2. Flaps + Slats Handle .................................................

This is the recommended final flap setting for appro

is a normal trim setting for a 20° flaps + slats lansufficient elevator control for the pilot.

3. Airspeed ......................................................................

by both pilots before landing. The land

ing horn, normally activated by the 40circuitry, will not sound if one or mor




ying gears are not fully extended and lo

If the stabilizer is jammed in the –4° to –10° rang

5. Flaps + Slats Handle ........................................

This is the normal range for the elevator trim wlanding. Make a normal approach.

6. Airspeed ...........................................................

LANDING WITH INOPERATIVE ELEVAT

The elevator portion of the control column will be frwill be unable to control movement about the pitchhorizontal stabilizer trim. Fly the approach and laprocedures:


It is recommended that the normal 40° flapapproach and landing. Check proper positionin

2. Airspeed ...........................................................

The extra speed will improve controllability oelevator.

3. Increase the landing distance by 1,800 feet.

ARTHUR UNIT INOPERATIVE

ORAIL

FEEL

PITCH

FEEL




These lights illuminate when either of the Arthur unit motects a failure. Normally, the autopilot can still be used. light is on, reduce airspeed down to 260 KIAS or MI 0.7

1. FASTEN BELTS Light Pushbutton ............................

The pitch and/or roll control forces may be hlower than normal, depending on whether the

unit has failed in the high- or low-speed pFollow either procedure below, dependingcontrol forces experienced.

Light forces—Avoid large displacements and rapid movcontrols.

High forces—Use normal or emergency trim systems. Ifis on, execute an approach at VREF. If the PITCH FEELan approach at VREF + 10 knots, and increase the landing d

FLAP ASYMMETRY OR JAMMED FLAPS

This light illuminates when a flap asymmetry is detectethe relative position of the outboard portion of the left

tb d ti f th l ft d i ht fl t i t

FLAP

ASYM

CAUTION

FEEL FEEL

2. EGPWS FLAPS O’RIDE Switch....................

The EGPWS FLAPS O’RIDE switch cancel




audible warning when landing with flaps less t

Increase the landing distance by 800 feet.

With the flaps extended between 7° and 20°:

1. Approach Speed ...............................................

This higher approach speed is flown to allowaircraft with any degree of asymmetry in thaircraft as well as possible under the circumsta

2. EGPWS FLAPS O’RIDE Switch....................

The EGPWS FLAPS O’RIDE switch cancelaudible warning when landing with flaps less t


With the flaps extended between 20° and 40°:

1. Approach Speed ...............................................

This higher approach speed is flown to allowaircraft with any degree of asymmetry in thaircraft as well as possible under the circumsta

2. Increase the landing distance by 200 feet.

CAUTION

SLAT MONITORING SYSTEM

AUTO

SLATS




Illumination of this light indicates that information fromcircuits (ground/flight proximity switches, angle of attadisparity.

If the light comes on after takeoff or at a speed lower th

1. Airspeed Envelope............................. BETWEEN 1.3

Do not allow the airspeed to go outside of thesesituations, as the stall warning system may not oindication of a stall and automatic deployment

deliberately perform stall tests.

If the light comes on when at an airspeed of 280 knots o

2. Reduce Airspeed ........................................................

The airspeed must be limited to less than 270 knots

270 knots, the safety systems that lock out the autothe slats are not operational, and unwanted and untimoutboard slats could occur, causing damage to theaircraft control.

SLAT SYSTEM ABNORMAL OPERATION

If using the flaps + slats handle on the center pedestal, ction of the red arrow light on the gear/ flaps + slats indicthat at least one of the four slats is being maneuvered, nenor fully retracted If using the emergency control or aut

SLATS

• During slat retraction, when the FLAPS

CLEAN, the green light flashes during retrand flaps. As soon as the inboard slats anoutboard slats retract at which time the gre




outboard slats retract, at which time the greand the red light comes on. The red light gfully retracted.

• During automatic or emergency extensioncomes on until the outboard slats are fu lly

the red light goes out and the green light flaretracting after auto deployment, the green fthe red light comes on until the slats are fu

• During automatic retraction of the inboardeployed and an AOA stall angle of 23° isgoes from a steady state to a flashing state of the retraction signal. This really means t

are fully extended. When a stall angle belowflashing light goes out, and the red light cof the inboard slats. When the inboard slatslight goes out, and the green light comes oall four slats are fully extended.

• If the Bus A1 failure occurs in the cockpit

extended through the normal slats controbus). In this case the green light flashes.

• Green and red light can never illuminate siing test on the indication panel.

In Cruise

Normally, in the cruise regime of flight, the flaps +CLEAN position. There should be no movement ocaused by selection of the handle or activation of tIf one or more of the slats fail to retract after the han

i i h d l i li h ill i Th

During Approach

Depending on indications, follow the instructions for either c




Case 1

A flashing green slat light with the red light off indioutboard slats are extended. This is the indication that onare extended. At this time the power used to supply the outb

solenoid is from the battery bus.

Proceed as follows:

1. Flaps + Slats Handle .................................................

The flaps may be extended to 40° as long as th

extended.

2. Approach Speed..........................................................

Due to the loss of the lift normally provided by the must be added to the VREF speed to provide the prostall speed.


Case 2

Assume that the flaps + slats handle has been placed SLATS position. The green slat light off, the red slat ligh

tendency for the aircraft to roll due to asymmetry indicaof the slats are not fully extended.

C O

If outboard slats are visually not extended an

with the following item 2.

2 Emergency Slats Switch




2. Emergency Slats Switch ..................................

The use of this switch will provide outboard sthe No. 2 hydraulic system.

Note the slats indications after placing the switchA, B or C below, depending on the indications you

Procedure A

If the green light comes on steadily, all slats exten

1. Flaps + Slats Handle......................................

A normal full 40° flap approach and landadditive or penalty.

2. Approach Speed ...............................................

Procedure B

If the red light stays on and only the outboard sextended, it can be assumed safe to use other flap coand landing.


The flaps may be extended to 40° as longextended.

If the slats cannot be extended by any or all means, th

of flaps for approach and landing.

2 EGPWS Flaps O'Ride Switch




2. EGPWS Flaps O Ride Switch.....................................

3. Approach Speed ..........................................................

To compensate for the loss of lift normally provided

the VREF must be increased by 30 knots to ensure astall during final approach and landing.

4. Increase the landing distance and the landing field len

The increased speed to be maintained on final appincrease in runway requirements to allow for a longer

Do not change the emergency slats switch p

The landing-gear-not-extended aural warnimally activated by the 40° flap warning cimay not sound if the gear is not fully downflaps are not selected to 40°.

UNWANTED OUTBOARD SLAT EXTENSION

The red slats transit light comes on and then goes out, indare in transit. The green slat light then comes on flashing,slats extension The audio warning for stall warning sou

AND IGNAUTO

SLATS

CAUTION



NOTE

Keep the flap/slat handle in this position foing in order to preclude the possibility of aasymmetry that can be particularly bothers




asymmetry that can be particularly bothersfinal approach.

With the airbrakes extended to position 1:

2. Airspeed ......................................................................

To compensate for the loss in lift due to the airbraposition 1, add 10 knots to the normal VREF to maintfrom stall.


With the airbrakes extended to position 2:

4. Airspeed ......................................................................

To compensate for the loss in lift due to the airbraposition 2, add 15 knots to the normal VREF to maint

from stall.


NOTE

Airbrakes may be used during landing approavided airspeed is at least VREF + 10 knots. Ilanding distance by 10%.

LANDING GEAR

ABNORMAL LANDING GEAR EXTENS




If, after the LANDING GEAR handle has been placand one or more of the following indications occur, LANDING GEAR EXTENSION.

• One or more green gear down light is out

• The landing gear handle light is flashing

• Landing gear not extended GEAR voice w

NOTE

The Landing Gear Abnormal Extensio

to be applied as soon as one green ligIt gives no alleviation to any red indicatgear symbol combination as long as osymbol is off.

The landing gear handle must be main

Emergency Landing Gear Extension

1. EMERGENCY GEAR PULL Handle .............

Unlatch and pull this T-handle, which is locahandle. Pulling this handle shuts off the ele

exposes the retraction side of the gear and dallows No. 1 system hydraulic pressure to be pside of the main doors and all three landing ge

CAUTION

Rapidly alternating large rudder applicationsbination with large side-slip angles may r

CAUTION




g p g ystructural failure at any speed.

2. LH MAIN MANUAL GEAR RELEASE Handle......

Gently apply up to full rudder to the left while accmax) until the left green gear down light is illuminate

Maintain wing level with appropriate aileron input .

NOTE

Illumination of the green gear down light mmore than 30 secs with full rudder deflectio

Gently come back to neutral rudder.

3. RH MAIN MANUAL GEAR RELEASE Handle .....

Gently apply up to full rudder to the right while ac

max) until the right green gear down light is illumina

Maintain wings level with appropriate aileron input.

NOTE

Illumination of the green gear down light m

more than 30 secs with full rudder deflectio


If, after applying all the methods possible to try to

problem with one or more landing gears still existassumption that at least one landing gear is locked an emergency, and make a shallow approach and as




If the nose gear is the one gear whose extension ithe main wheels as for a normal landing. Hold thelong as possible and then bring it gently into contthe elevators remain effective. Do not use the

emergency exists, until after the nose touches the gtouches the ground, begin normal, or differentialmaintain directional control. Use No. 1 brakes, Nbrake handle, depending on the availability of systems.

If extension of either of the main landing gears isthe side of the runway corresponding to the exten

level for as long as possible. Control direction wnosewheel steering. When the wing touches the gropedals and counteract veering.

CONTROL HANDLE JAMMED IN DOW

It is impossible to retract the landing gear.

1. Airspeed ...................................................... 24

Jamming of the landing gear control handle icaused by misalignment of the nosewheels; cto free the control handle by pressing the pushbutton located above the control handle

handle if the modification M1688 is installed)if the gear is retracted in this configuration.

Keep the landing gear extended to the landing

a. If the red landing gear lights fail to go out up

landing gear, ice may be preventing the main landin the UP position.

2 Cycle the gear down and up to get rid of the ice




2. Cycle the gear down and up to get rid of the ice.

In nonicing conditions or if takeoff was made withthe runway:

a. Extend and keep the landing gear down.

EMERGENCY RETRACTION ON GROUND

Emergency retraction of the landing gear on the ground mered if it is imperative to stop the aircraft immediately (ranother aircraft or a large obstacle) and conventional mea

Ground retraction of the landing gear is achieved using tdure, with the first two operations being performed simu

1. Ground Safety OverridePushbutton (If Installed) ....................................... PR

This is a red pushbutton located above the normalhandle. On aircraft with modification M1688, this pinstalled. Unlocking of the gear control handle is perf

2. Landing Gear Control Handle.....................................

3. Power Levers...............................................................

Normal shutdown of engines.

4 FUEL SHUT OFF i h ( ll h )

• The #2 P BK light illuminates at 232 psi.

• Brake pressure for the No. 1 brake system

• 1 595 psi for SNs 1 75 and subsequen




• 1,595 psi for SNs 1, 75, and subsequenService Bulletin F900-42

• 2,175 psi for SNs 2 to 74, except for thF900-42

NOTE

Service Bulletin 42 provides a doubletem, that assures a better brake-holding cing pretakeoff, full-engine-power run

• Brake pressure for the No. 2 brake system

• The first detent of the parking brake hand2 hydraulic system pressure. This detent swithout locking the brakes.

• The second detent of the parking brake prhydraulic system pressure. This detent wil

This malfunction is indicated by an abnormgreen No. 1 system L and R brake lights dbrakes are tested in either the #1/ASKIDpositions of the brake selector switch.

1. Brake Selector Switch......................................

Set the three-position selector switch to #2achieved using the pedals that are suppliesystem; the antiskid system is inoperative. Tbraking by pressing the LH and RH brake pe



NOSEWHEEL SHIMMY

Hold the nosewheel steering control depressed.

This applies hydraulic pressure to the nosewheel s




pp y psome shimmy dampening.

FUEL SYSTEM

LOW BOOST PUMP PRESSURE

Case 1

(4.64 psi)

1. No. 2 Booster Switch .......................................

If the No. 2 booster switch was in the NORST–BY position to turn on the other boost pCheck the FUEL 2 light. If it goes out, continstandby boost pump.

If the FUEL 2 light stays on:

2. Associated Fuel Quantity.................................

Carefully observe the No. 2 fuel quantity. Cis dropping normally through normal engeither procedure A or B below, depending onnot evident.

Procedure A

FUEL 2

5. GEN Switch ................................................................

The generator for the shutdown engine is no longerturned OFF to preclude electrical anomalies aelectrical system




electrical system.

6. Engine Anti-ice Switch ...............................................

This closes the anti-icing valves (air intake and ENG

In icing conditions, operate the No. 2 engine aneven with the No. 2 engine shutdown. The ivalve must be open to allow air bleed from thair manifold to anti-ice the S-duct.


7. Bus-Tied Switch ..........................................................

Since the No. 2 generator is no longer supplying elright DC electrical buses, tie the buses to save the

depletion. Check the volts and amps on the two opeillumination of the bus-tied light.

8. ST–BY Hydraulic Pump Switch.................................

If needed to supply hydraulic power to the No. 2 hydthe standby hydraulic pump switch. A windmillin

upon its rpm, may not be able to supply enoughoperate No. 2 system components.

9 No 2 Fuel Tank Fuel

CAUTION

To prevent an engine flameout, carefully monitoduring this operation. The goal is not to run out obefore returning the fuel panel to a configuration, can be kept running from the fuel supplied by groEngine Inoperative Approach and Landing Proced




Procedure B

If no fuel leak is evident:

The flight may be continued, as required, by feedinby gravity, from the group 2 fuel tank only, excabove 31,000 feet that must be performed with X-Bthat X-BP 1-2 and 3-2 lights are illuminated.

Case 2

1. X-BP 1–3..........................................................

Rotate the X-BP 1-3 switch to connect group 1both lateral engines. The corresponding X-BP li

2. Associated Booster Switch...............................

Turn off the boost pump switch that correspopressure warning light.

If the fuel pressure l ight stays on after opening t

3. X-BP 1–3 .........................................................

FUEL 1 FUEL 3OR (4.64 PSI)

If a fuel loss is evident:

5. Associate Engine Power Lever...................................

Since the area of the fuel leak is unknown, an isolatio




Since the area of the fuel leak is unknown, an isolatioshutting down the associated engine. This action engine at the fuel control.

6. Associated Fuel Shutoff Switch..................................

Further isolation of the engine and fuel system is action. Continue to monitor the fuel quantity to seecontinues. In any situation involving a fuel leak, a pmight be appropriate. Check for TRANS light on, the

7. GEN Switch ................................................................

The generator for secured engine is no longer useful off to preclude electrical anomalies associated with t


This closes the nacelle anti-icing valve.

OR

If the fuel pressure warning l ight goes out:

The flight may be continued using fuel management prorate with flight requirements.

3. XTK Switch Set to Low Level Side............................

2. X-BP 1-2 and X-BP 3-2 ...................................

Crossfeed the lateral engines using the two Xlower portion of the fuel panel. Check thatlights go out.




g g

If the FUEL 1 and FUEL 3 lights are out:

3. No. 1 and No. 3 Booster Switches ...................

4. Fuel Quantity Indicators ...................................

To consume fuel in side tanks when no leak is

5. Flight Altitude ..................................................

Since gravity flow of fuel, assisted by air prefuel can be transferred from the tank groupaltitude of the aircraft.

6. X-BP 1–3, X-BP 1–2, and X-BP 3–2 ..............

FUEL 1 and FUEL 3 Lights .............................

7. Fuel Quantity Indicators ..................................

The range of aircraft may be severely affecaltitude. Continually monitor the fuel gages capability and for any abnormal consumption o

If the FUEL 1 or FUEL 3 light remains on (and bly come on):

7. GEN Switch ................................................................

The generator for the secured engine is no longer turned off.





FUEL TRANSFER SYSTEM MALFUNCTIONAIRCRAFT EQUIPPED WITH XTK 2 SYSTE

Case 1

This light indicates that the XTK 2 valve is open when At times, this situation can be corrected by merely movinthrough each of its respective positions, then back to AUcorrect the indication, follow either procedure A or B be

Procedure A

If group 2 fuel total quantity indicates approxima green range):

1. No. 2 Rear Tank Quantity...........................................

Depress the button next to the group 2 tank fuel gagquantity in the rear tank.

If the group 2 rear tank level is 3,300 pounds and steadyrear tank is full:

XTK 2

OPEN

3. Manual XTK Valve ..........................................

This manually positioned valve is located inaircraft, in the floor on the left aisle, in line wwindows. Lift the carpet cutout to gain access f h l d i h l d i i




of the valve and rotate it to the closed positioreturn the flap to the stowed position. Nowwarning panel.

XTK 2 OPEN Light..........................................

The XTK 2 OPEN light should go out after th2 valve has been closed. Fuel management of trear tank must be carefully followed. It may bevalve once the rear tank level has burned down

OR

Procedure B

I f g roup 2 fue l quan t i t y i nd i ca te s approx(amber range):

1. No. 2 Rear Tank Quantity................................

If No. 2 rear tank level is above 1,400 pounds:

2. XTK 2 Switch ..................................................

This is done to shut off the fuel transfer from t

XTK 2 OPEN Light..........................................

NOTE



TANK LEVEL ABNORMALLY LOW

ON A/C WITHOUT XTK 2 SYSTEMLO

FUEL 1

LO

FUEL 2

LO

FUEL 3OROR




This light indicates that a fuel level below 200 pounto a cruise flight time of 15 minutes at an altitude

1. Associated Fuel Quantity Gage .......................

Verify the indication given by the warning lighfuel quantity gage. However, the warning ligaccurate of the actual fuel quantity in the fuel t

2. Associated X-BP..............................................

In order to supply fuel to the low-side engine, gained from a fuel tank containing more quanrespective valve opened should come on.

3. Booster of Affected Tank .................................

Turn off the low-side boost pump until it becouse the fuel from that low tank.

TANK LEVEL ABNORMALLY LOWON A/C WITH XTK 2 SYSTEM

LOFUEL 1

LOFUEL 2

LOFUEL 3

OROR

3. Booster of Affected Tank ............................................

Turn off the low-side boost pump until it becomes abuse the fuel from that low tank.

If h LO FUEL 2 li h d h N 2




If the LO FUEL 2 light and the No. 2 totais higher than the No. 2 rear tank quanindicates a malfunction of front-to-rear t

4. XTK 2 Switch .............................................................

If the rear tank quantity is still decreasing:

5. If necessary, MANUAL XTK 2 Valve .................... O

Disregard illumination of the XTK 2 open light.

When the rear tank quantity has increased:

6. No. 2 Booster Switch..................................................

7. X-BP............................................................................

Check that the X-BP light is out.

FUEL ASYMMETRY

This condition is indicated by asymmetric fuel indicationormal application of aileron trim to one wing versus th

the asymmetry is not caused by a fuel leak. Follow eithbelow, depending on the type of asymmetry:

LO

FUEL 2



Avoid the following:

• High noseup or nosedown pitch attitudes

• Rapid changes in pitch or roll




2. Fuel Quantities............................................................

Monitor fuel quantity indicators to detect any fuel lo

the aircraft as soon as possible in the event a fuel lethe fuel tank vent system or if structural damage is su

The FUELING light comes on if the following occurs:

• One of the two fuel tank vent valves is not closed

• The defueling/refueling valve is not closed.

• The refueling connector door is not closed.

• The refueling control panel door is not closed.

• The gravity-fueling switch is on.

• The defueling switch is on.

• The vent valve lever is not stowed.

• The B2 bus is not powered.

• PRESSURE FUELING circuit breaker

ELECTRICAL SYSTEMS

Case 1—Any Generator Switch is Trip

1 Bus Voltages

GEN 1SWITCH OFF GEN 2OR OR




1. Bus Voltages .....................................................

Check that the voltages are within prescribed l

2. Batteries and Generators Load .........................

Check the ammeter readings for each battery and generator loading is within limits. Eacrunning record of normal bus loading and havepower demands on the electrical system.

3. Shed the load, as necessary, to limit the load operating generator.

Do not attempt to reset a tripped genSince the generator switch has tripped

age condition will exist if a reset attem

If the No. 2 generator is inoperative and BATT 2 l

4. Bus-Tied Switch ...............................................

This will allow a sharing of the load between t

will preclude abnormal demands on the No. 2 bus-tied light illumination and normal bus load

CAUTION

Case 2—GEN 1 Switch is Not Tripped

1. Bus Voltages ...............................................................

GEN 1




1. Bus Voltages ...............................................................

Check that the voltages are within prescribed limits.

2. Batteries and Generators Load ...................................

Set the ammeter selector to the position correspondconcerned. On aircraft fitted with an auto-load systembe indicated on the generator still connected to the bnecessary, to keep the load on each remaining genethe aircraft without the auto-load feature.

If the left main bus voltage is normal:

The absence of an overvoltage condition for bus A is anonline generator is the one associated with the extinguis

3. GEN 1 Switch........................................... OFF, THE

ATTE

If resetting cannot be achieved:

4. GEN 1 Switch .............................................................

OR

If the left main bus voltage is above the green range:

The existence of an overvoltage condition not high enoug

If voltage and amperage indications are norma

OR

If the GEN 1 and GEN 3 lights remain on, represemal No. 1 generator did not automatically reconne




4. GEN 1 Switch........................................... OFF


5. GEN 1 Switch ..................................................

The generator is no longer useful and should belectrical anomalies.

6. Left and Right Main Bus Volts/Amps ..............

Never tie the buses together without previouslamperage are within limits.

Check the volts and amps on the left main bus

for normal valves. Normal voltage and ampgenerator should be shown for the right main b

7. Bus-Tied Switch ...............................................

Ensure that the left and right main buseequalization on voltmeters and ammeters

prescribed limits. Check for bus-tied light illum

8. Bus Load and Voltage.......................................



OR

If the GEN 1 and GEN 3 lights remain on, represemal No. 1 generator did not automatically reconne

4. GEN 3 Switch........................................... OFF





5. GEN 3 Switch ..................................................



Never tie the buses together without previouslamperage are within limits.

Check the volts and amps on the left main busfor normal valves. Normal voltage and ampgenerator should be shown for the right main b

7. Bus-Tied Switch ...............................................

Ensure that the left and right main buseequalization on voltmeters and ammetersprescribed limits. Check for bus-tied light illum




2. Batteries and Generator Load ..........................

Set the ammeter selector to the position corrconcerned. On aircraft fitted with an auto-loadbe indicated on the generator and/or battery stibus and normal loading on the right main bus.




Shed the load on the bus, if necessary, toon the operating generator and/or battthe buses together without previouslythe voltage and amperages on each busprescribed limits.

Do not attempt to reset the generatotripped switch, as an overvoltage condiwhen a reset attempt is made.

On the generator for which the switch has not trwithin the green range:

3. GEN Switch(es) ...............................................

Attempt two resets maximum of the generatorGEN switch off, and then on, to see if the generator will not reset, and main bus voltagrange, then:

4. Both GEN Switches .........................................

Check to see that amperage indications on bot

CAUTION

CAUTION



If the GEN 1 and GEN 3 lights remain on, normal No. 3 generator did not automatically r

4. GEN 3 Switch........................................... OFF






5. GEN 1 Switch ..................................................


OR

Procedure C

If GEN 2 light is on:

1. GEN 2 Switch ..................................................

Turn the GEN 2 switch off and then on a maxthe generator will reset. If the generator cannot

2. GEN 2 Switch ..................................................

3. Left and Right Main Bus Voltage ............ NOT

Check the right battery load for normal indicat

4. Bus-Tied Switch ...............................................

Ensure the left and right main buses are tied byvoltmeters and ammeters and that they are witf b i d li h ill i i

Case 2—GEN 2 and GEN 1 or GEN 3 Switches H

Two generators have been disconnected from the maThe respective generator switches have tripped. This psume a simultaneous tripping of the reverse current

GEN GENAND




sume a simultaneous tripping of the reverse current sponding generators.

1. Bus Voltages ...............................................................

Check that the voltages are within prescribed limits.

2. Batteries and Generator Load.....................................

Set the ammeter selector to the position correspond

concerned. On aircraft fitted with an auto-load systembe indicated on the generator and/or battery still connbus and normal loading on the right main bus.

Shed the load on the bus, if necessary, to limit

on the operating generator and/or battery. Nthe buses together without previously checkthe voltage and amperages on each bus are wiprescribed limits.

3. Bus-Tied Switch...................................... CHECKED

Ensure the buses are not tied when attempting to resswitch has tripped. The objective is to protect the buthe generator being reset.

CAUTION

Perform only one reset attempt of generator attempt should be made, for the reason th

indication that the fault still exists, and adetrimental effects.

Carefully observe the voltmeter/ammeter wswitch on. Be prepared to immediately return




p p yovervoltage condition exist. In fact, the overvothe switch to trip.

If No. 2 generator cannot be reset:


If the volts and amps are normal (in the green rang

8. Bus-Tied Switch ...............................................

Check that the bus-tied light is on and monitokept within limits.


10. Power Lever of No. 2 Engine afterGenerator Reset Attempt .................................

Case 3—GEN 1 and GEN 3 Switches Have

Two generators have been disconnected from the respective generator switches have tripped. This pi lt t i i f th t l f th

GEN GENAND

Shed the load on the bus, if necessary, to limit on the operating generator and/or battery. Nthe buses together without previously checkthe voltage and amperages on each bus are wiprescribed limits.

CAUTION




3. Bus-Tied Switch...................................... CHECKED

Ensure the buses are not tied when attempting to resswitch has tripped. The objective is to protect the buthe generator being reset.

4. BAT 1 Switch..............................................................

The battery will not only provide electrical power tofunction as a buffer when attempting to connect the g

5. Power Lever of No. 1 Engine.....................................

This will minimize the effects, should the reset overvoltage condition. Consequently, the reset at

performed during a critical phase of flight when engreduced.

Engine Idle Setting .....................................................

6.GEN 1 Switch .......................................................................

Perform only one reset attempt of generator 1. Noattempt should be made, for the reason that a findication that the fault still exists, and a new d i l ff

8. Power Lever of No. 3 Engine ...........................

This will minimize the effects should the overvoltage condition. Consequently, the reperformed during a critical phase of flight whreduced.




Engine Idle Setting ..........................................

9. GEN 3 Switch ..................................................

Perform only one reset attempt of generator attempt should be made, for the reason thindication that the fault still exists, and adetrimental effects.

Carefully observe the voltmeter/ammeter wswitch on. Be prepared to immediately returnovervoltage condition exist. In fact, the overvothe switch to trip.

10. Power Lever of No. 3 Engine afterGenerator Reset Attempt .................................

If No. 1 and No. 3 generators cannot be reset:

11. Right Main Bus Volts/Amps.............................

Check to see that the voltage and amperage aand its associated No. 2 generator will pick

when connecting the main buses. If the indicat

12. Bus-Tied Switch ...............................................

BATTERY OVERHEAT

Additionally, the battery temperature indicator red light

HOTBAT




In flight:

Determine which battery is overheating by referring to tture gage and checking the individual battery temperaturtermining which battery is overheating:

1. Associated BAT Switch...............................................

This action disconnects the battery from the elecshould reduce the load or charging action that may becondition.

If the battery temperature keeps rising:

2. Land as soon as possible.

NOTEIf required, the faulty battery may be switchon for landing, provided the HOT BAT light hout. Closely monitor the battery temperature in

On the ground:

If the batteries are warm (temperature higher than 120°F)is on, a battery start must not be attempted. Use an extstarting the engines.

BATTERY FAILURE

The associated battery switch is tripped. If groundprotection device occurs, or if a battery has an interna

i h i d h i f h

BAT 1 BAT 2OR




switch trips open under the action of the reverse currswitch flips off, and the corresponding BAT light on t

If these indications occur:

1. Associated Battery Switch................................

No more than two resets of any electrical systeTry to reset the associated battery by moving thof one or both batteries in flight will not hoperation of electrical systems. Howevergenerators fail, the emergency power normalwill be lost if the batteries cannot be reset.

PITOT-STATIC SYSTEM

EITHER AIR DATA COMPUTER INOPERATI


Table AP-2. PITOT-STATIC SYSTEM—EITHAIR DATA COMPUTER INOPE



FUNCTION LOSS OF INDICATIONS

“On-side”EFIS

Cousou

Airspeed scale

ASEL

IAS flag on EADI

ASEL flag on EADI

Datrec

“On-side”AP-FD CPLD

Vertical modes Reversion to basic modeMSG: CPLD DATAINVALID on ID 802.

Cousou

“On-side”flight instruments

Altimeterrate-of-climbindicators

Flags Useothside

ID 802 CPLD“on-side”

SAT-TAT-TAS Dashes oncorrespondingline

Dat“x s

Horizontalstabilizer trim

If ADC 1 hasfailed the –4°

Normal trim limitedto –4°

Usetrim

Autoslats Extensioninhibition athigh speed

Possibility

light

SeemaAUTO

SLATS

BOTH AIR DATA COMPUTERS INOPER



EFIS M h

Table AP-3. PITOT-STATIC SYSTEM—DATA COMPUTERS INOP



light

AP

MACH

TRIM

flashing onEADI,

APon warning

panel and

AP-M TRIMYD-FD

Autopilot,Mach trim, andyaw damper

MSG: DADC DATAINVALID in ID 802

command bars go ouof view.

Autoslats Extensioninhibition athigh speed

ID 802 SAT-TAT-TAS Dashes oncorresponding line

Flight instruments Altimeters,rate-of-climbindicators

Flags

AUTO

SLATS

Horizontalstabilizer trim

–4° limitation

at high speedor stop over-riding at

Normal trim limitedto –4°

EFIS MachAirspeed scale

AS

ELWind on EHSI

IAS flag on EA

ASEL flag on EA

JAMMED OR ABNORMAL PILOT,COPILOT, AND POSSIBLY STANDBYIAS/MACH INDICATION AT HIGH ALTITUDE

This problem may come about as a result of aircraft operatiin colder than normal altitude temperatures and whilecrystals, which may block the pitot heads. The problemdescent is made to an altitude where air temperatures wil




descent is made to an altitude where air temperatures wilCross-check normal instruments with standby instrumen

The warning are as follows:

The VMO /MMO audio warning sounds, the IAS EADI cotor illuminates, the AP disengagement and/or DADC DATAappears on ID 802, and a disagreement with standby IAS/Ma

If it is certain that the VMO /MMO warning is fnot modify flying parameters.

If the above indications are experienced:

1. Autopilot and Yaw Damper ........................................

Due to possible spurious information inputs to adisconnect the autopilot and yaw damper to preventcontrol inputs. Hand-fly the aircraft, and avoid any

and rapid movements of the flight controls.

If the VMO/MMO warning sounds:

CAUTION

AUTO

SLATS

MACH

TRIMAP

AIL

FEEL

weight and temperature. Since SAT/TAT indicuse the last known temperature observed, or us

6. Aircraft Attitude ...............................................

This would be a normal indicated attitude fowise for every pilot to learn the normal attitudof flight




of flight.

After the fault has been identified, follow Procedof flight, Procedure B for the cruise phase of fligdescent phase of flight.

Procedure A

Climb

1. N1 RPM ............................................................

Set maximum climb thrust N1 in accordancethrust charts contained in the Airplane Perform

2. Aircraft Attitude ...............................................

This would be a normal climb attitude if a clhigh altitudes.

Procedure B

Cruise or Level Flight

Set N1 rpm according to the setting for Mach 0.75 cflight alti tude, aircraft weight, and TAT (if known

Without Anti-icing

1. N1 RPM ......................................................................

Reduce the power levers to idle rpm for the depressurization requirements. If at high altitude, ensuremaintained with the power levers at idle.




2. Vertical Speed........................................................ –2

3. Aircraft Attitude .................................................... 0°

This attitude range should assure a –2,000- to –3,000

With Engine and Wing Anti-icing On

1. N1

RPM ........................................ MINIMUM BY A

Use the three engines operative minimum N1 anti-iNormal Procedures section of the Airplane Flighwhich is printed in this manual, or in the abbreviated

2. Airbrakes .....................................................................

Deploy the airbrakes to position 1 to provide drag in engine power is kept at a higher power setting for ant

3. Vertical Speed in the Descent................................ –1

4. Aircraft Attitude .................................................... 0°

This attitude range should assure a 1,500- to 2,00engines at a higher rpm, while the airbrakes are deplo

b. 25,000 to 34,000 feet ................

c. 34,000 to 51,000 feet ................

3. Reset AUDIO WARN A and AUDcircuit breakers at frequent intervaaudio warning has stopped. Leavbreakers engaged if the warning h




breakers engaged if the warning h

4. If IAS/Mach indications are unrelicated static air temperature may als

PROBE ANTI-ICING MALFUNCTION

Compare instrument readings with the readings of thbus power and proper switch positioning.

ICE PROTECTION SYSTEMS

WING ANTI-ICE INOPERATIVEWITHOUT BRAKE HEATING

The following four cases, with procedures, assuswitch was turned on for anti-ice protection and th

Case 1

L. PITOT R. PITOT ST BY

PITOTOROR

If the light remains amber:

3. No. 3 Engine N1 Speed...............................................

While maintaining a higher rpm on the No. 1 engine,the No. 3 engine to attempt to supply more high-presthe PRV 3 valve.

If h li h




If the light turns green:

4. No. 3 Engine N1 Speed...............................................

Keep the thrust increased on both No. 1 and No. 3 enconditions, to keep the wing anti-ice light green and anti-icing capability.

If the light still remains amber:

Set the N1 No. 1 and No. 2 engines to the minimum rpmAP-4 for operations in icing conditions.

Avoid or leave icing conditions as soon as possible. Sinceis no longer provided, structural icing may result.

Case 2

The wing anti-ice switch is on, and the wing amber ligindicates a system malfunction in which there is too muin the system

NOTE

D o no t go be l ow t he m i n i m um e

prescribed for flight in icing conditionstemperature (Table AP-4) .


Table AP-4. FLIGHT IN ICING CO




Maintain that engine speed on No 1 and No 3 eng

THREE ENGINES OPERATIVE—AN

ONE ENGINE INOPERATIV

TAT –30° TO –20° TO –1 –20°C –10°C –

ON AIRCRAFT WITH WING-BRAKE

>20,000 80% 76%

>20,00076% 73%

>10,000

<10,000 68% 65%

Increase the values of the above by the fo

• 9% if N1 is equal to or higher than

• 6% if N1 is lower than 65%

When the heating system is used:

• The minimum required N1 speed wi

engines operating must be increas

Case 4

The wing anti-ice switch is on, and the wing green light is othat the green wing anti-icing light is not operative for onThe corrective action to determine if it is an indication

i l bl i




operational problem is to:

1. Wing Anti-ice Switch..................................................

Turn off the wing anti-ice switch, and observe thwatching the wing anti-ice lights. Normally, when tice switch off, the amber light will flash a couplextinguish. Turn the switch back on.

2. Wing Anti-ice Switch .................................................

If you observe the normal indications of the amberthen going out, you can assume that the wing valvesbeing supplied sufficient bleed air for anti-icing purpo

If you do not observe these proper indications, the wmust be considered inoperative. You must:

3. Avoid or leave icing conditions as soon as possible.

WING ANTI-ICE UNWANTED

OPERATION WITHOUT BRAKE HEATING

OR

If the green light comes on, it means the opersystem is out of sequence. High-pressure bleedthe wings even when the wing switch is off. operation of the wing anti-ice system or sesystem electrovalve in the open position. If thfollowing procedure must be followed to elimhigh pressure bleed air to the wings




high-pressure bleed air to the wings.

2. Wing Anti-ice Switch.......................................

The amber light may still be flashing.

3. Isolation Valve Knob........................................

Move the rotary switch on the overhead panelamber ISOL light should come on. This separthat the No. 1 and No. 3 engines feed one sfeeds the other side of the bleed-air manifold.

4. HP 1 and PRV 3 Switches................................

Moving these switches to off should keep anythe No. 1 and No. 3 engines, from entering theside of the manifold from which the wing anits air. Therefore, only low-temperature, lowdelivered to this side of the manifold. To minimlow-pressure bleed air might have, perform the

5. Reduce No. 1 and No. 3 engines power setting

WING ANTI-ICE INOPERATIVE WITH B

Procedure A

Switch is in the WING position:

1. No. 1 Engine N1............................................ INCRELIG

Increase the power on the No. 1 engine to attempt pressure bleed air through the HP 1 valve.





2. No. 1 Engine N1..........................................................

OR

If the amber light remains on and steady:

2. No. 3 Engine N1 ..........................................................

While maintaining a higher rpm on the No. 1 enginon the No. 3 engine to attempt to supply more higthrough the PRV 3 valve.


3. No. 3 Engine N1..........................................................

Keep the power increased on both No. 1 and No. 3 econditions, to keep the wing anti-ice light green and anti-icing capability.

OR

If th b li ht i d t d

OR

If the amber light remains on and steady:

4. WING–BRK Switch.........................................

If in icing conditions, do not reduce N1 spee(see Table AP-4).




5. Avoid or leave icing conditions as soon as poss

Wing anti-icing is no longer available.

Procedure B

Switch is in the WING–BRK position:

1. WING–BRK Switch.........................................

If the green light comes on, it is an indicasystem is inoperative.

If the green light stays out:

2. WING–BRK Switch.........................................

Increase the N1 speed of the No. 1 engine untilthe attempt is unsuccessful, repeat similar N1engine. Maintain this N1.

If the light does not turn green and if in icing cond

3. Maintain an N1 rpm not less than 1% above ti i i diti ( T bl AP 4)

Case 2

The WING–BRK switch is in the position as indicated ibelow and the amber light is flashing with the green ligh




below, and the amber light is flashing with the green lighexcessive anti-icing to the wings.

NOTE

When reducing N1 rpm as directed below, doduce the N1 below the minimum speed requoperations in icing conditions.

Procedure ASwitch in the WING position:

1. Reduce N1 speed of No. 1 engine until the amber fland the green light comes on.

If this does not correct the problem:

2. No. 1 Engine N1 Speed ..............................................

3. Reduce N1 speed of No. 3 engine until the amber fland the green light comes on.

Procedure BSwitch in the WING–BRK position:

Case 3

The wing anti-ice switch is on, and both the ambeare on. It would be quite rare for such an indication the aircraft manufacturer feels that this is a coverany other indescribable indications might be seen




any other indescribable indications might be seen anti-icing switch. This indicates a general total mtem, and you must:

Avoid or leave icing conditions.

Case 4

The WING–BRK switch is in the position as indicateand the amber and green wing anti-ice lights are o

Procedure ASwitch in the WING position:

1. WING–BRK Switch.........................................

The amber light should flash and then go out. anti-ice valves were open and that they closed

2. WING–BRK Switch.........................................

Procedure B

Switch in WING-BRK position:

1. WING–BRK Switch...................................................

If the green light stays out:

2. WING–BRK Switch....................................................




The amber light should flash and then go out. This inanti-ice valves were open and that they closed normal

3. WING–BRK Switch....................................................

Carefully watch the indicator lights as the switWING–BRK position. The amber light should come

out. This is a normal indication; however, the green If the green light did not come on, there is a malfuncsystem, and it can be assumed that the wing anti-icproperly.

If the amber light did not come on steady and then go ou


If in icing conditions, do not reduce engine rpm below

WING ANTI-ICE UNWANTEDOPERATION WITH BRAKE HEATING

OR

If the green light comes on, it means the operation ofis out of sequence. High-pressure bleed air may stieven when the wing switch is off. The cause is eithewing anti-ice system or seizing of the wing anti-iceopen position. If the TAT is above +10°C, the folfollowed to eliminate, or to limit, the flow of high-pre

2 WING BRK S i h




2. WING–BRK Switch.........................................

The amber light may still be flashing.

3. Isolation Valve Knob........................................

Move the rotary switch on the overhead panel

amber ISOL light should come on. This separthat the No. 1 and No. 3 engines feed one sfeeds the other side of the bleed-air manifold.

4. HP 1 and PRV 3 Switches................................

Moving these switches to off should keep anythe No. 1 and No. 3 engines, from entering theside of the manifold from which the wing anits air. Therefore, only low-temperature, lowdelivered to this side of the manifold. To minimlow-pressure bleed air might have, perform the

5. Reduce No. 1 and No. 3 engines power setting

ENGINE ANTI-ICE INOPERATIVE

If the amber light goes out and the green light illuminate

2. Retain this thrust setting while in icing conditions.

If the amber light does not go out:

3. In icing conditions, do not reduce N1 below the Table AP-4).




4. Avoid or leave icing conditions.

Case 2

ENG 1, ENG 2 or ENG 3 amber and green lights on.

The engine anti-ice switches are on and both the amber andare on for one or more engines. It is quite rare for suappear. However, the aircraft manufacturer feels that this isin the eventuality any other indescribable indications migturn on the engine anti-icing switches. This indica

malfunctioning of the system and you must:

1. If in icing conditions, do not reduce N1 below theTable AP-4).


Since this is an unknown condition and impossiblebest to avoid or leave icing conditions.

1. No. 2 Engine Anti-ice Switch ..........................

Observe No. 2 Engine Amber Light.................

As you cycle the switch off, carefully waNormally, the amber light should flash a extinguish. If this is observed, turn the anti-icshould observe that the amber light comes on assume the system is operating, but only w




malfunction.

If the amber light does not come on and then go ou

2. In icing conditions, do not reduce N1 beloTable AP-4).


The No. 2 engine anti-icing system must btherefore, you must leave or avoid icing condit

ENGINE ANTI-ICE OVERPRESSURE

ENG 1, ENG 2 or ENG 3 flashing amber light.

The associated engine anti-ice switch is on and the aon one or more engines. This indicates that the preice system of the corresponding engines is too hig

ENGINE ANTI-ICE UNWANTED OPERATION

Case 1ENG 1 or ENG 3 amber light flashing.




The associated engine anti-ice switches are off and an This may mean that some bleed air may still be going to surfaces even though the engine anti-ice switch is off.

1. Associated Engine Anti-ice Switch ............................

If the green engine anti-ice light does not come on, occurred in the indicating system.

Flight may be continued with:

2. Associated Engine Anti-ice Switch.............................

If the green engine light comes on, it means the operationicing system is out of sequence. High-pressure bleed air mto the engine surfaces even when the engine anti-ice swi

If the TAT is above +10°C, the following procedure must inate, or to limit, the flow of high-pressure, high-temperaengine anti-ice surfaces. Make sure the associated anti-iperform the following:

3. Associated Engine Power Thrust................................

1. No. 2 Engine Anti-ice Switch...........................

If the green engine anti-ice light does not comoccurred in the indicating system.

The flight may be continued with:





If the green No. 2 engine light comes on, it means thanti-icing system is out of sequence. High-pressusupplied to the engine air intake and S-duct surfaanti-ice switch is off.

If the TAT is above +10°C, the following proceeliminate, or to limit, the flow of high-pressure, hto the No. 2 engine anti-ice surfaces:


The amber light is now on steady.

4. Isolation Valve Knob ........................................

Move the rotary knob on the overhead panel amber ISOL light should come on. This separthat the No. 1 and No. 3 engines feed one sfeeds the other side of the bleed-air manifold.

5. PRV 2 Switch ...................................................

Moving this switch to off should keep any Nhigh-temperature airflow from entering the bside of the manifold from which the S-duct an



After turning off the PRV 3 switch, carefully light, and note any changes in its actions.

If the BLEED OVHT light starts blinking aPRV 3 switch off and continue the flight in tthe faulty system.

OR

If the BLEED OVHT light starts blinking and




If the BLEED OVHT light starts blinking andconditions:

2. Associated Engines Blinking Light..................

Move the associated power lever to idle to lowand pressure in the manifold. Set N1 of other eengine inoperative conditions (see Table AP-3

anti-icing.

NOTE

If you are not in icing conditions, and toing to reduce to idle on an engine, an cedure is possible.

1. Turn the isolation valve knob to is

2. Set the passenger air-conditioning voff if the overheat is from No. 2 en

3. Set the crew air-conditioning valvif the overheat is from the No. 1 or

4. Move the COND control lever, lcopilot’s right console, to the tied

For the aircraft with AUXITROL EL 124 box (modification

• After a flight with a BLEED OVHT warning lighcockpit, it is mandatory to check the LEDs in thon EL 124 box before switching off electrical po

For the aircraft with AUXITROL EL 137 box (modificati

• After a flight with a BLEED OVHT warning lighcockpit, it is mandatory to check the LEDs in the r




EL 137 box. This information remains available afttrical power, which makes maintenance of the ble

Modification M1905 is applied in production since aircraft by changing boxes.

ECU OVERHEAT

This light indicates failure of the cold-air

1. Passenger Temperature Controller........ MANUAL/M

Move the passenger temperature control toward

minimum of 40% of the dial. This will decrease theair that is directed through the turbocooling unit comp

2. Passenger Air-Conditioning Valve Switch (If Require

If it is necessary to further limit the amount of air dthe turbocooling unit, move the passenger air-conditi

the off position. The air demand will quickly diminlight should go out.

ECU

OVHT

5. Crew Temperature Controller ............... MANU

Move the crew temperature control toward the40% of the dial. This again will decrease thedirected through the turbocooling unit. Permaking a larger demand than the passenger sys

6. Crew Air-Conditioning Valve Switch (If Requi




If it is necessary to further limit the amount othe turbocooling unit, move the crew air-condoff position. The air demand will quickly dimshould go out.

If the ECU OVHT light is still on:

7. Airspeed.................................................... LES

Reduce the airspeed of the aircraft to less thThis will allow the turbofan to operate andbottom aft of the empennage to open to prthrough the heat exchangers.

NOTEIf you are not in icing conditions, the and PRV 3 switches may be turned off

If the overheat warning persists:

8. Passenger Air-Conditioning Valve Switch (If R

If you are unable to control the overheat con

CABIN AIR-CONDITIONING UNIT OVERHEA(ON AIRCRAFT WITHOUT TURBOCOOLING

UNIT ANTI-ICING EMERGENCY CONTROL)High air temperature occurs along with the light.

COND’G

OVHT




This light indicates overheating in one of the passenger or cducts. Check the valve positions on both the passenger ancontrol panels.

1. Temperature Controllers..............................................

Move the mode selector switch on each panel from to turn off the automatic regulation system. Move thcontrol switch to the cold position.

If the light does not go out or if the temperature does no

2. Passenger Air-Conditioning Valve Switch..................

Move the passenger air-conditioning valve switch

eliminate some of the excess temperature comiconditioning ducts. However, in order to provide air cthe aircraft, you must perform the following procedur

3. COND Control Lever .................................................

Move the COND control lever to the tied position

systems together. This will provide circulation of airconditioning unit, which will supply the entire interio

CABIN AIR CONDITIONING OVERHEAEQUIPPED WITH ANTI-ICING EMERGE

(SB-131)

High air temperature occurs along

This light indicates overheating in one of the cabin ducts. Check the valve positions on both the passen

COND’G

OVHT




control panels.

1. Temperature Controllers...................................

Move the mode selector switch on each paneto turn off the automatic regulation system. Mcontrol switch to the COLD position.

Since anti-icing on the turbine is no lothe amber ECU A/I pushbutton can onhigh altitude (above 35,000 feet) whconcentration is practically close to z

If altitude is below 35,000 feet:

See note below.

If the temperature is NOT decreasing:

2. ECU A/I Pushbutton ........................................

The turbocooling fan is no longer deiced. Chethe ECU A/I pushbutton.

CAUTION

OR

If the temperature is decreasing:

3. Temperature Controllers .............................................

The temperature decrease confirms the source of malfunctioning turbofan anti-icing valve, which has result of the activation of the ECU A/I pushbutton.




At the beginning of descent:

4. Pressure Norm/Emergency Switch.............................

This will simultaneously close the passenger air-codrive the crew temperature control valve to full hposition indication on the control panel.

5. ECU A/I Pushbutton ...................................................

Check for extinguishing of light.

If temperature is too high and if not previously accompl

6. Crew Switch ................................................................

7. COND Control Lever .................................................

NOTE

For operation below 35,000 feet with the

COND’G OVHT light on, complete only thlisted items:

NOSE CONE OVERHEAT

This light indicates that the nose cone temperatur

1. Nose Control Lever (If Installed) .....................

NOSE

CONE OVHT




Check that the nose conditioning lever, locconsole, is in the normal, or forward, positionused for inflight ventilation of the electronic nose cone.

2. Unnecessary Avionics ......................................

Check maintenance panel MINELCO indicatequipment.

EFIS and MFD symbol generators are locaterequire ventilation. However, certain radios, aare also located in this area. This additional from aircraft to aircraft. It is recommendecomponents contained in the nose cone be m

aircraft in the event that an electrical isolation b

PRESSURIZATION

IMPROPER CABIN VERTICAL SPEED

This is an erratic indication on the cabin verticalcabin pressure should be maintaining a stable diffsteps are necessary to ensure that the controls and s

2. PRV 2 and 3 Switches.................................................

3. Bleed-Air Crew and Passenger Switches....................

Ensure that the air-conditioning valve switches are inif that does not seem to correct the problem, move thThis will bypass the automatic opening circuit and ato full open. Check that the PRV 2 and PRV 3 switchto ensure a proper air supply to the bleed-air manifoland pressurization purposes.




If normal operation is not restored:

4. Pressurization UP–DN Knob.................... 1- TO 2-O

At this point, it is assumed the automatic pressube malfunctioning. By moving the manual pressu

1- to 2-o’clock position, you command a level cabinpneumatic outflow valve in preparation for manual pYou may notice a slight fluctuation in cabin verperforming this step. This is normal, as the automacompensate for the manual knob being out of its com

If you do not move the knob to the 1- or 2-o’clo

pressurization surge may result when selecting manua

5. Auto/Manual Pressure Selector Switch.......................

Move the two-position switch, located to the rpressurization knob, from the AUTO to the MAN pautomatic pressurization control and allows man

pressure through the use of the manual pressurization

6 Pressurization UP DN Knob ADJ

TOO HIGH CABIN PRESSURE

The indication for this situation is seen on the cab

tor for altitude and differential pressure. An abnorcan cause pressurization surges if the overpressureit reaches its preset relief values.

1. Cabin Pressure Controller................................

Move the PROG–FL–LDG switch on the cab

FL iti Y l t fli ht l l




FL position. You can now select a flight levelflying, which increases cabin altitude and keeriding on the maximum differential limits.

2. Select a higher flight level.

If the cabin pressure does not decrease, the automa

be inoperative. Take the following steps:

3. Pressurization UP–DN Knob............................

In the event the cabin pressure problem is cpressurization system, this step prepares themanual operation. This position of the knob altitude command on the pneumatic outflow va

4. AUTO/MANUAL Pressure Selector Switch....

This is done to disable the automatic (electrenable manual (pneumatic) control of the press


Moving the knob clockwise from the 1- or 2-

If flight requirements so dictate maintaining altitude, cycsenger switches off to on periodically to maintain a cabithan 8,000 feet or a cabin differential pressure no greater9 psi. Descend, as required.

TOO HIGH CABIN ALTITUDEOR SLOW DEPRESSURIZATION

CABIN




The aural warning occurs along with the light.

The red CABIN light will illuminate accompanied by theing as heard from the cockpit speaker system. This indicatitude has climbed higher than 10,000 feet. This is a loss o

should be dealt with in a timely manner. If climbing, the climuntil the problem is corrected. A descent may even be in otitude, immediate donning of the oxygen masks is mandat

1. Bleed-Air Crew and Passenger Conditioning.............

1A. PRV 2 and PVR3 Switches.........................................

Check that the crew and passenger air-conditioning valthat air is coming into the aircraft. PRV 2 and 3 switchthese actions have not been taken, then air will not be acabin. It is suggested that the crew and passenger aswitches be placed to the on position if it is thought thais not working.

2. BAG Switch................................................................

When placing the baggage switch to ISOL you shut


In the event the cabin pressure problem is cpressurization system, this step prepares themanual operation. This position of the knob altitude command on the pneumatic outflow va

5. AUTO/MAN Pressure Selector Switch............

This is done to disable the automatic (electrbl l ( ti ) t l f th




This is done to disable the automatic (electrenable manual (pneumatic) control of the press

6. Pressurization UP–DN Knob ...........................

Use the manual knob to control cabin precounterclockwise, from the 1- or 2-o’clock p

down signal to the pneumatic outflow valve. counterclockwise position closes the pneucommands an approximate 1,500-foot-per-micabin, thereby increasing pressure differential.

If cabin pressure cannot be restored:

7. Isolation Valve Knob ........................................

This action separates the bleed-air manifold. Ton. This is an attempt to see if there is a bleemanifold or the other. If there is a leak, it willbleed air will be available from the other sipressurization.

If cabin pressure is restored:

8 Cycle the crew and passenger air conditionin



10. Microphone Selector........................................

Set the audio panels on each side console tomike communications in the cockpit.

11. No Smoking Sign.............................................

Smoking and oxygen are not a good mix in anover the loud speaker system might be appropr

12 P O M k




12. Passenger Oxygen Masks ................................

13. If necessary, execute an emergency descenminimum safe altitude, as required.

DOOR UNLOCKED INDICATIONCase 1

No aural warning occurs.

The red cabin light illuminates without any aural wawere heard, it is another abnormal procedure for t

This indicates that either the cabin access door forward toilet service door (for those aircraft so eqand latched.

If on the ground, proceed as follows:

1. Cabin Entrance Door........................................

CABIN

2. Cabin Entrance Door .................................................

Visually check the alignment of the index marks oposition of the two crankpins in their tracks on themicroswitches for cabin door warning indication are of each crankpin track.

If the door is, or seems to be, improperly locked:

3. Cabin Differential Pressure ................................ RED




Be cautious in the method of reducing cabin pressurthat the way to reduce cabin differential is to begin athe normal automatic reduction of pressurization diffsudden rash reduction of pressure by dumping might


Case 2

This light indicates that either the baggage compartment rear compartment door, is not fully closed and latched.

Access to the baggage compartment is not permitted.

If the suspect door is the baggage compartmeaccess to the baggage compartment must be

WARNING

REAR

DOORS

Case 3

This light should normally be out. This light comethe access door to the baggage compartment. Entrment is permitted in flight to a maximum flight le

If the baggage compartment is not in use, visually

of the baggage compartment access door.

BAG

ACCESS




gg g p

Case 4

This light illuminates if the baggage compartmentnot fully open. The baggage compartment may baggage access door is closed. Check the BAG selecbleed-air panel.

OXYGEN—NO AUTOMATIC DEPLOYM

If the oxygen masks do not automatically deploy w11,500 ±750 feet:

1. Oxygen Controller ............................................

This emergency manual setting for the control

if the automatic system fails.

2 Passenger Masks

BAG

ISOL

WINDSHIELD

CRACK OR BUBBLES FORMIf a windshield gets fissured or cracked, the flight can bdanger. Pressurization integrity should be maintained.

1. Limit Airspeed....................................................... 23

By limiting airspeed, the adverse effects of heavy a

window are lessened.




2. Cabin Differential Pressure........................................

Maintain this pressure differential as a maximumcracking due to a high-pressure differential.

3. Associated Windshield Heat Switch ...........................

HEAT SYSTEM INOPERATIVE

This light illuminates when there is an electrical short left or right regulator probe. Heating regulation is automto the operating regulator. System performance should n

1. Pilot and Copilot Windshield Heat Switches ..............

Pilot and copilot windshield heat switches should be sposition.

XFR

AUTOMATIC FLIGHT CONTROL

AFCS OUT-OF-TRIM CONDITIONCase 1

A message will also appear on the ID 802: PITCH M

DOWN. This indicates that there is a permanent loa

MISTRIM




p

1. Hold the control wheel firmly.

Before disconnecting the autopilot, grasp afirmly to overcome any out-of-trim conditionautopilot is disconnected.

2. Autopilot ..........................................................

Disengage the autopilot by pressing the disenbottom aft portion of the control wheel. Tdisconnected by pressing the go-around buttoor the emergency elevator trim switch or by

breakers. Be ready for any unusual aircradisconnect is initiated. The AP light will coshow disengagement. The AP light can be eautopilot disconnect button once again.

3. Manually trim the aircraft.

In other words, hand fly the aircraft and retrim

4 Try to reengage the autopilot

MACH TRIM INOPERATIVE

This indicates a malfunction of the Mach trim device.

Do not exceed .80 Mach unless the autopilot is engaged

ANGLE-OF-ATTACK STALL

MACH

TRIM




ANGLE OF ATTACK STALLPROBE HEATING FAILURE

Illumination of either light indicates a failure of the anticorresponding angle-of-attack stall probe.

1. Avoid icing conditions.

ANGLE-OF-ATTACK PROBE—HEAT SYSTEM INOPERATIVE

This indicates fai lure of the anti-icing system of tindicator probe.

The angle-of-attack indicator system must not be used i

AOA PROBE

L. AOA R. AOAOR

FAILURE OF APU FIRE PROTECTION

Illumination of this light indicates a defect in monitoring circuit.

1. APU..................................................................

EFIS

FAULT




EFIS

NOTE

I n t he e ve n t o f a f a i l u r e o f e i t happroach will be flown by the pilot CRTs operational.

NOTE

In the event of the failure of the EHSIcraft so equipped with the multifunctioMFD controller may be selected to theif required.

EITHER EADI CRT FAILURE

The CRT display on the EADI will go blank, or thdifficult to interpret.

On the associated EFIS reversion controller:

1. EADI ON–OFF Dimmer Knob .......................

EITHER EHSI CRT FAILURE

The CRT display on the EHSI will go blank or the color

difficult to interpret.


1. EHSI ON–OFF Dimmer Knob ..................................

Information that was formerly displayed on the

transferred and will appear in a composite form on th




If a multifunction display is installed and if desired:

2. Place the mode selector switch in the EHSI positioninoperative CRT.

The EHSI display will now appear on the MFD. TEADI will disappear, leaving the full EADI display.

SIMULTANEOUS FAILURE OF EADIAND EHSI CRTs ON THE SAME SIDE

Both CRT displays go blank, SG flags appear on the EH

color is altered and difficult to interpret. This indicates respective symbol generator for that EFIS system.


1. SG Pushbutton............................................................

This will allow the good EFIS information that is ppilot’s side to be transferred to the side that wmalfunction. White XSG annunciators will illumin

SUCCESSIVE FAILURE OF EADIAND EHSI CRTs ON THE SAME SIDE

In this situation, one CRT goes blank, soon folgoing blank.

1. EADI/EHSI ON–OFF Dimmer Knobs............

Move both knobs to the counterclockwise posi

2. Place the mode selector switch in the EHSI pinoperative CRT




inoperative CRT.

3. Use the standby horizon.

4. The pilot whose CRTs are operating flies the a

LOSS OF ASCB CONTROL

XDATA flag appears on both EADIs.

Items that are lost:

• Cross-side data

• Glide-slope, localizer, and radio altimeter

NOTE

If the aircraft is equipped with a third IRbe used in the backup mode for both I2 simultaneously. The associated displIRS annunciator on all four CRTs.


NOTE

If the aircraft is equipped with a third IRS, IR

be used in the backup mode for both IRS 1 a2 simultaneously. The associated display is aIRS annunciator on all four CRTs.

IRS ATTITUDE COMPARISON ANNUNCIATWITHOUT A HEADING COMPARISON ANN

IRS comparison annunciators appear on both EADIs andinvalid message appears on the ID 802.



g pp

1. Standby Horizon .........................................................

Compare the standby attitude indicator with the infothe EADI.

2. Faulty IRS ...................................................................

Troubleshoot the electronic instrument systems tfaulty IRS.

When the faulty IRS is identified, on the EFIS reversion

the faulty side:

3. IRS Pushbutton...........................................................

To cross over the system, to put the operative side the IRS pushbutton, and observe that an XIRS flag bad IRS.

IRS HEADING COMPARISON ANNUNCIATI

Follow either Procedure A or B below:

Procedure AIf the faulty IRS is in the NAV mode and the posion the reversion controller for the faulty side IRS

1. IRS Pushbutton.................................................

To cross over the system, to put the operativethe IRS pushbutton, and observe that an XIRS




OR

Procedure B

If the faulty side IRS is in the NAV mode, but

incorrect, on the mode select unit (MSU) of the fa

2. Selector.............................................................

Move the MSU knob from NAV to ATT. Thcomputer and allows the IRS to become a basi

3. FMS of Faulty Side IRS........................... ENT

To enter the correct heading for the FMS to trthe POS SENSORS page on the FMS. Whenselect the status page of the IRS that is in ATspace provided on this page. The heading information provided by the other FMS or

compass if straight and level and with the wind

IAS DISPARITY ANNUNCIATION

IAS comparison annunciation appears in each EADI.

It is possible that a DADC DATA INVALID message will as well.

1. Standby Airspeed Indicator.........................................

Cross-check the information given on the standby ai

the information given on each EADI. The EADI nearly the same, information as the standby airspconsidered having the good DADC information




considered having the good DADC information.

After the faulty DADC has been identified, on the facontroller:

2. IAS/Mach Pushbutton .................................................

This action will allow the crossover of IAS and Mathe good side. The XIAS annunciators in each EADI

LOC OR GS DISPARITY ANNUNCIATION

LOC or GS comparison annunciation appears in each EAbelow 1,200 feet AGL.

1. Faulty ILS ...................................................................

Check each navigation radio for proper identificatiotry to determine which radio is malfunctioning. Turreceiver off and:

2 Proper ILS Radio for Each Side

EMERGENCY PROCE

CONTENTS

GENERAL............................................................

ENGINE FIRE......................................................

INTERNAL ENGINE FIRE ON THE GROUND

APU FIRE.............................................................




BAGGAGE COMPARTMENT FIRE..................

MAIN WHEEL WELL OVERHEAT...................

AIR-CONDITIONING SMOKE..........................

ELECTRICAL SMOKE OR FIRE.......................SMOKE REMOVAL ............................................

INADVERTENT THRUST REVERSER

DEPLOYMENT IN FLIGHT...............................

TWO ENGINES INOPERATIVE—

APPROACH AND LANDING ............................Preparation ..................................................

Approach.....................................................

When Committed for Landing ....................

After Touchdown ........................................

TWO ENGINES INOPERATIVE—GO-AROUN

On the Go Around

Approach...............................................................

Just Before Touchdown.........................................

After the Airplane Has Come to a Stop ................

DITCHING.....................................................................

Preparation ............................................................

Approach—Parallel to the Major Swell ................

Just Before Touchdown.........................................

After Touchdown ..................................................




ILLUSTRATIONS

Figure Title

EP-1 Fire Panel and Warning Lights ........

EP-2 Fire Panel and Warning Lights—FIR

EP-3 Fire Panel and Warning Lights—FIR

EP-4 Bleed-Air Panel ..............................

EP-5 Inflight Airstart Envelope................EP-6 Hydraulic Panel ..............................




EMERGENCY PROCE

GENERALWhere these emergency checklist procedures diffe

Manual , the Airplane Fl ight Manual takes preced

Compliance with the order prescribed for applicis recommended.

Where more than one phase for a procedure is spe




• specifies immediate action to be need for reference to the checklist.

• Phase 2 items shall be completed only afteaccomplished by checklist.

• Phase 3 items shall be accomplished as so

Aural warnings shall be identified before being siaural warnings will enable better coordination duriemergency procedures.

The Airplane Operating Manual prescribes the follaural warning sounds without the presence of a FI

• Test the detection system by activating the

• If a light does not come on, the fire

• If all the lights come on, the aural w

In all cases, carefully observe all other indicationfirm either a malfunction or a false alarm

FIRE

FIRE

Phase 1

ENGINE FIRE

NOTE

The following procedure must be followed, wor not the FAULT light is on.

Phase 1

Af i i l id if i h i ff d d il i

FIRE + AURAL WARNING




After positively identifying the engine affected and silencin

1. Power Lever of Engine Concerned..............................

Moving the power lever cutoff closes the fuel suppl

fuel control.

2. Fuel Shutoff Switch of Engine Concerned .................

Raise the guard on the fuel shutoff switch (Figure switch up to electrically close the fuel shutoff valve. is located in the crossfeed manifold downstream from

Carefully observe that the amber TRANS light, loshutoff switch, comes on and then goes out after actuated. This light will illuminate during the moveorder to confirm positioning of the valve with the pos

3. Airspeed ...........................................................

The airspeed must be reduced to ensure that twhen discharged to the engine, stays within thsiphoned overboard.

4. Fire-Extinguisher DISCH Switch....................

Break the safety wire holding the fire-extingWithout pulling out on the switch, carefu

position 1. This is best done by placing onlyand moving it upward from the bottom.





Selecting position 1 discharges one fire-extiengine or two fire-extinguisher bottles to the ce

If the f ire warning persists:

5. Fire-Extinguisher DISCH Switch....................

Position 2 is a lever-locked position, and the out and moved upward to select this position.

Selecting position 2 discharges one alternate

lateral engine and two alternate fire extinguisher

Phase 2

Proceed with phase 2 items after phase 1 itemthe checklist.

CAUTION

7. GEN Switch ................................................................

The generator switch should be placed off to remove from the electrical circuitry.

8. Engine Anti-ice ...........................................................


9. Bus-Tie Switch ...........................................................

With the No. 2 engine shut down, generator powl i l b i l T h N 2 b




electrical buses is lost. To prevent the No. 2 batterytie the bus to supply the right-side bus from thgenerators. Check the generator volts and amps to ebeen tied. A higher-than-normal bus load may remaining generators. This higher load may be cause

being charged after tying the buses. Continue to monthe No. 2 battery is fully charged.

10. Standby Hydraulic Pump ...........................................

Consider moving the standby hydraulic pump switchdriven hydraulic pump may not have sufficient windm

operation of the No. 2 hydraulic system components. Iposition, the standby pump cycles only to support airb

If in icing conditions, operate the No. 2 engiice even with the engine shut down. The S-dcontinue to be anti-iced through bleed air s

from the main manifold by the No. 1 and Ngine.

CAUTION

If a fire breaks out during a starting attempt, or wgine, the engine must be rotated, with the power lefire is out.

1. Power Lever......................................................

Move the power lever to cutoff to shut off the f

2. Fuel Shutoff Switch .........................................

Move the fuel shutoff switch (Figure EP-2) umanifold for the engine affected.





This arms the start circuit for engine motoring.

4. Start Button ......................................................

Hold the respective engine start button depreThis rotates the engine without the presence oevacuate any residual fuel, causing the fire to e

5. Crew and Passenger Bleed Switches................

This is done to keep fumes from entering the c

6. Evacuation......................................................

Initiate the evacuation when crew and passenbrigade is also advised.

APU FIRESee Figure EP-2 for location of the FIRE APU and FAU


DISCH

2

1

0

FAULT

FUELSHUTOFF

TRANS

DISCH

2

1

0

FAULT

FUELSHUTOFF

TRANS

DISCH

2

1

0

FAULT

FUELSHUTOFF

TRAN



NOTEThe following procedure must be followed, wor not the FAULT light is on.

NOTE

If a fire or an overheat is detected and indica

APU stops automatically through closurefuel shutoff valve and removal of the el

APU FIRE + AURAL WARNING

FAULT

FIRE 1

FAULT

FIRE 2

FAULT

FIRE3

Figure EP-2. Fire Panel and Warning Lights—

NOTE

Wait for ten seconds to allow the APUcomplete stop before discharging thguisher. This will ensure that the distays within the APU compartment to infectiveness of the agent.

4. APU Extinguisher Switch ................................

There is only one position on the APU exting

bottle to the APU compartment. This same bothe baggage compartment.




If the fire warning persists:

5. No. 2 Engine Fuel Shutoff Switch...................

The fuel supply to the APU is provided throsystem. If the fire persisted, shutdown of the further attempt to isolate the cause of the fire.then extinguishing of the TRANS light on the

6. No. 2 Engine Power Lever ...............................

The No. 2 engine must be shut down as well tby firefighting personnel. This is the continensure engine shutdown.

7. Booster 2 Switch ..............................................

As it is no longer needed, the boost pump mus

8. No. 2 GEN Switch............................................

BAGGAGE COMPARTMENT FIRE

See Figure EP-3 for location of FIRE BAG COMP light

Phase 1

1. BAG Switch.................................................................

Move the BAG switch from NORM to the HEXAT

FIRE

BAG COMP + AURAL WARNING




Move the BAG switch from NORM to the HEXAT shuts off the hot bleed-air supply to the baggage cthe isolation valve between the cabin and baggageopen, the pressure between the two compartments allowing the baggage access door to be opened for fir

DISCH

2

1

0

FAULT

FUELSHUTOFF

TRANS

FIRE 1

DISCH

2

1

0

FAULT

FUELSHUTOFF

TRANS

FIRE 2

DISCH

2

1

0

FAULT

FUELSHUTOFF

TRANS

FIRE3

Fi EP 3 Fi P l d W i Li ht FI

The lavatory door must be properly c

opening the baggage compartment dosmoke and fumes from entering the pas

If access to the baggage compartment is ncompartment has depressurized, the aircrafaircraft is in the takeoff or landing phase of flpilots is required in the cockpit.

1. BAG Switch.................................................

This will close the isolation valve betwe

CAUTION




This will close the isolation valve betwecompartment in order to keep fumes fromConfirmation of valve closing is obtained BAG ISOL light and ISOL light indicatioIt may serve a secondary purpose by starv

fire existing in the compartment. Your jsituation and applying the proper decisionvery important.

2. BAG COMP Extinguisher Switch ..............

Break the safety wire on the baggage comswitch and move the switch to position 1.baggage compartment.

Phase 2


The aircraft must be landed at the nearest suand repair.



4. COND Control Lever .......................................

On the copilot’s right console, move the ca

control lever aft. This interconnects the ductinsystems. Clean conditioned air will now beconditioning system by the passenger air-condair from only the No. 2 engine.

OR

Troubleshoot the system to determine which causing the smoke. The air-conditioning sysback to normal operations, provided the defe3% below that of the other two engines. Sou




gyour choice in this emergency. It might be pr1 above.

If smoke persists:

5. Crew Bleed-Air Switch....................................

The passenger bleed-air system becomes suspcoming into the aircraft before proceeding furtthe crew bleed-air switch to ON.

6. Passenger Bleed-Air Switch.............................

If the smoke decreases or disappears, confaulty bleed-air switch isolated. The smoke pNo. 2 engine.

If the smoke persists, the cooling unit is suspect. P

7. Crew Temperature Controller..........................

If the smoke persists:

8. Crew Bleed-Air Switch...............................................

The only remaining procedure is to turn off the entirthe air-conditioning system. The aircraft will now byet may still be filled with smoke.

9. Descend to 14,000 feet or below or to the minimum s

The following procedure must not be applied i

CAUTION




are present in the cabin or cockpit.

At or below 14,000 feet:

10. Pressurization Dump Switch.......................................

Lower the guard on the dump switch, and set the position. This action opens the outflow valves apressure at a rapid rate down to the cabin altitud14,000 ±500 feet. Smoke should be evacuated fromdump operation.

At or below and indicated airspeed of 215 knots:

11. LH Direct-Vision Window ..........................................

If required to help eliminate smoke from the cockpvision window may be opened.

Phase 3

ELECTRICAL SMOKE OR FIRE

Phase 1

Smoke and Unusual Odors

1. Crew Oxygen Masks and Smoke Goggles .......

The crew oxygen masks must be put on immedsmoke. Be careful of eye glasses when puttpersonal injury and to avoid a poor fit of the m

mask must fit tightly around the face to ensurthe mask and enter the breathing pasNORMAL–100% oxygen selector button on th100% (pure oxygen) position. Remove the




00% (pu e o yge ) pos t o . e ove t estorage area, and place them over the eyes witvision during the smoke/fire emergency.


For hot mike communications within the cselected and the mask and CPIT buttons on panels must be pushed on. To communicate the proper radio microphone button on the aMIC button on the top inside of either control w

3. No Smoking Sign.............................................

Only if there are no flames in the cabin:

4. Passenger Oxygen Controllerand Passenger Masks........................................

On the copilot’s right console, move the selecoxygen panel clockwise to OVERRIDE and

Phase 2

1. Crew Air Gaspers........................................................

This action is designed to increase the airflow in thsmoke evacuation as the checklist continues.

If the origin of the fire or smoke is evident:

2. Suspected Equipment .................................................

If equipment has been malfunctioning, or if smokcoming from a specific electrical component, turn ofis the apparent cause of smoke or fire.




pp

OR

If the origin of the fire or smoke is not evident, and permit a total electrical shutdown:

2. GEN 1, BAT 1, BAT 2, GEN 2, and GEN 3 Switches

If inflight situational conditions will permit such aBAT and GEN switches to shut off all electrical paircraft. After the fire/smoke has stopped, attempt to

fire/smoke by turning on the batteries, one at a time, causes the fire/smoke to recur. When the origin is kno

3. BUS TIED Switch.......................................................

Confirm BUS TIED light out on the master warning p

4. BAT Switch (Affected Side)........................................

Move the BAT 2 and GEN 2 control switcheaction eliminates the electrical power supplybuses. The amber BAT 2 and GEN 2 lights o

panel will illuminate.

3. Bus-Tied Switch...............................................

4. BUS TIED Light ..............................................

Ensure that this normal flight configuration,

DC bus from the left DC bus, is maintained. Tthe troubleshooting procedure be effective.

5 Passenger Temperature Controller




5. Passenger Temperature Controller ...................

Set the passenger temperature controller to maswitches to the full cold position of the indicat

If the smoke or fire persists:

6. AUTO–MAN Pressurization Selector Switch..

Since the left DC buses are no longer suppwhich includes control of automatic pressuriza

be selected to control the cabin pressure. Adjucontrol knob, as required, to control the cabin p

7. BAT 2 and GEN 2 Switches ............................

This action is taken to restore electrical powebuses, prior to isolating the left DC electrical

smoke or fire was not caused by a component o

If fire or smoke is not visibly verif ied removed:


The aircraft must be landed at the nearest suitable and repair.

If smoke persists:

11. Crew Air Gaspers........................................................

This action is designed to increase the airflow in thsmoke evacuation as the checklist continues.




12. Descend to 14,000 feet or below or to the minimum s

The following procedure must not be applied iare present in the cabin or cockpit.


13. Pressurization Manual UP–DN Control ......................

Positioning of the manual knob control will ensuredifferential and cabin depressurization.

At or below an indicated airspeed of 215 knots:

14. LH Direct-Vision Window ..........................................

If required to help eliminate smoke from the cockp

CAUTION

SMOKE REMOVAL

Phase 1

1. Crew Oxygen Masks and Smoke Goggles .......

The crew oxygen masks must be put on immsmoke. Be careful of eye glasses when puttpersonal injury and to avoid a poor fit of the mmask must fit tightly around the face to ensurthe mask and enter the breathing pasNORMAL–100% oxygen selector button on th100% (pure oxygen) position. Remove the gogand place them over the eyes with a tight fit tothe smoke/fire emergency




the smoke/fire emergency.


For hot mike communications within the cselected, and the mask and CPIT buttons on panels must be pushed on. To communicate the proper radio microphone button on the aMIC button on the top inside of either control w

3. No Smoking Sign.............................................

If there are no flames in the cabin:

4. Passenger Oxygen Controllerand Passenger Masks........................................

On the copilot’s right console, move the selec

oxygen panel clockwise to OVERRIDE, andPASS ON indicator is fully open. This po



Phase 3

1. Descend to 10,000 feet or to the minimroute flown.

This is the maximum altitude for sustained flunpressurized aircraft.

If the smoke persists or if the fire is not visibly ve


The aircraft must be landed at the nearest suand repair.




INADVERTENT THRUST REVER

DEPLOYMENT IN FLIGHT

The aircraft may experience a pitch-down momenmal noise and buffeting.

The REV UNLOCK light means the thrust reversesupposed to be fully stowed. The clamshell doors not be fully stowed and locked.

The TRANSIT light means the clamshell door cla

The DEPLOYED light means the synchronizing l h ll d h d h d l d i i

REV

UNLOCK

TRANSITAND POSSIBLY DEPLOYEDAND/OR

3. Airspeed.................................................................. 23

This will additionally reduce the airloads on the clam

Phase 2

If the thrust reverser stows:

1. Continue the flight with the NORM/STOW switch in

Do not actuate the thrust reverser handle. Leave the Nin STOW to maintain a constant retraction signal won the stow side of the actuator.




If the thrust reverser remains deployed:

2. Thrust Reverser NORM/STOW Switch ......................

Keep the thrust reverser NORM/STOW switchthe emergency stow signal and hydraulic pressure onthe actuator.


Land at the nearest suitable airport for inspection and

NOTE

The drag associated from an idling No. 2 with the thrust reverser deployed, adverselythe performance characteristics of the aircr

recommended that the engine be shut downproach and landing.

TWO-ENGINES INOPERATIVE—APPROACH AND LANDING

PREPARATION

Reduce the aircraft weight as much as practical. Iaircraft exceeds the maximum limits for landing wing distance, burn off fuel, if practical.

Determine the weight limitation for enroute climfound in the Flight Manual, Section 5, subsection

1. Fuel Supply to Remaining Engine....................

Check that the booster pumps and booster c




Check that the booster pumps and booster cproper configuration to ensure an uninterroperating engine. A reasonable wing fuel balaconditions permit. Group 2 fuel should be u

and-balance maintenance of the aircraft center

2. Bus-Tied Switch ...............................................

Tie the left and right DC buses together togenerator supplies the electrical power requireshould be on.

3. Limit Generator Load ......................................

Generator limitations of 300 amps, 350 ampsobserved. Check the buses for proper load dgenerator. Turn off nonessential electrical equi

4. Crew and Passenger Bleed-Air Switches .........

6. ST–BY Hydraulic Pump Switch.................................

The standby hydraulic pump is now available to back

driven hydraulic pump, should No. 2 system pressure

7. Avoid icing conditions.

The one remaining engine may not be able to supplfor anti-icing purposes without jeopardizing the poweflight of the aircraft.

APPROACH

1. Flaps + Slat Handle.....................................................




p

When in the traffic pattern, in preparation for the fithe slats and flaps to the first detent, provided aircrareserve permit.

2. Emergency Slats Switch (If Necessary andInoperative Engines are No. 1 and No. 3)...................

The No. 2 hydraulic system is available to provide outboard slats at VFE (200 KIAS).

3. Make the decision to land or go around at not less ground level.

The aircraft should be in a position to land at this poA go-around is not recommended when below 1,00level. See the Two Engines Inoperative Go-Around ch

WHEN COMMITTED FOR LANDING

1. Landing Gear....................................................

NOTE

If the No. 1 or No. 3 engine is the opethe gear may be extended using the norm

If the No. 2 engine is the operating engine , it is necgency extension as follows:

2. Normal Gear Handle……………………………

CAUTION




The landing gear handle must be main

3. EMERGENCY GEAR PULL Handle …………

The electrical sequence is bypassed . The No. 1 syststraight to the gear uplock boxes and actuators.

If all three green gear down lights illuminate anlight (red) is not illuminated, the landing gear is dotuate any landing gear controls.

If at least one green gear light does not illuminatdle light is flashing , apply the FREE FALL EXTE

Free Fall Extension Procedure

4. Airspeed............................................................

CAUTION

NOTE

Illumination of the green gear down light mmore than 30 seconds with full rudder defle


6. RH Main Manual Gear Release Handle .....................

Gently apply up to full rudder to the left while accmax) until the left green gear down light is illumin

level with appropriate aileron input.

NOTE

Illumination of the green gear down light m




Illumination of the green gear down light mmore than 30 seconds with full rudder defle

Gently come back to neutral rudder .

7. Nose Gear Manual Release Handle............................

Accelerate until illumination of the green gear lighKIAS max).

NOTEFree fall extension of the three landing getake approximatelytwo minutes to complete

8. Flaps/Slats (As Required) .........................................

At some altitude below 1,000 feet above ground level,

is down and locked, extend the flaps to 20°, if perform

NOTE

With No. 1 and No. 3 engines inoperatislats using the emergency system. Ex

manually as described above.

AFTER TOUCHDOWN

1. Thrust Reverser (If Available)..........................

The normal deceleration devices, such as airbshould be applied, if available. The thrust revehigher speeds of the landing rollout.

If the No. 2 brake system must be used, use m




braking is not available. Use the thrust reverseeffective braking speeds. This is necessary wheels and the possibility of tire failure.

If No. 2 braking system is used, increase the above by 50%, due to the lack of antiskid bra40° flaps for landing.

TWO ENGINES INOPERATIVE—

The decision to land or go around muor above 1,000 feet above ground leveloss associated with this go-around prproximately 500 feet.

ON THE GO AROUND

CAUTION

The landing gear cannot be retracted if No

No. 3 engines are inoperative.

3. Accelerate the aircraft while in descent on normal slo

At VREF + 25 knots:

4. Flaps + Slat Handle.....................................................

As soon as the airspeed reaches the 40° flap VREF +slats and flaps incrementally to clean.

5. Accelerate to and maintain the enroute climb speed

CAUTION




pthe aircraft.

As soon as the airspeed of 1.43 velocity of stall in this achieved, begin to climb the aircraft to a safe aenroute climb speed until the assigned, or a safe, altit

The altitude loss associated with this go-arou

cedure is approximately 500 feet.

ALL ENGINES INOPERATIVE

Phase 1

1. Communications .........................................................

CAUTION



Phase 2

If no engine(s) can be started:

1. Prepare for a forced landing or a ditching.

Refer to the Forced Landing or Ditching checklists fothe manual.

2. ST-BY Hydraulic Pump (ForEmergency Flaps + Slats Extension) ..........................

Remember to use this pump carefully, as its powbatteries will be great. Turn it on only when you aremergency slats and the flaps to the desired configuroff after the desired configuration has been achieved




off after the desired configuration has been achieved.

3. Extend the flaps/slats at VFE (200 KIAS) using the em

if necessary.

If a forced landing is anticipated:

4. Extend the landing gear.

At VLO (190 KIAS) if possible, and if necessary, pe

emergency (freefall) extension of the landing gear.

The landing gear should not be extended if ing is anticipated.

CAUTION

LOSS OF BOTH HYDRAULIC SY

Hydraulic indicators show a pressure drop (Figure

PUMP 1 PUMP 2

PUMP 3AIL

FEEL

PI

FEAND POSSIBLY AND




Phase 1

1. Autopilot and Yaw Damper .............................

Figure EP-6. Hydraulic Pa

Phase 2

1. Avoid high-pitch attitudes and zones of air turbulence

Very slight movements of the control surfaces argentle bank and pitch changes. If at all possible, avoiturbulence and high wind conditions.

LANDING PREPARATION

1. Flaps + Slats Handle ...................................................

Since there is no hydraulic pressure to the systems, tmust not be actuated.




2. Landing Gear ..............................................................

With no hydraulic pressure available to lower the geamanual extension of the gear is required.

3. Normal Landing Gear Handle .....................................

4. Emergency Hydraulic Gear Control ...........................

Push in on the latching mechanism to release the locagainst the panel, and pull the handle aft to its mecha

5. Airspeed .............................................................. 160

Maintain an airspeed of 160 to 190 knots to faclowering. Extend the main landing gear first.

NOTE

Illumination of the green gear down ligmore than 30 seconds with full rudde


7. RH Main Manual Gear Release Handle ..........

Gently apply up to full rudder to the left whmax) until the left green gear down light is i

level with appropriate aileron input.

NOTE

Illumination of the green gear down ligmore than 30 seconds with full rudder




more than 30 seconds with full rudder


8. Nose Gear Manual Release Handle .................

Accelerate until illumination of the green gearKIAS max).

NOTE

Free fall extension of the three landtake approximately two minutes to co

Do not actuate the landing gear contror mechanical, once the landing gear

locked.

CAUTION

• 40° Flaps + Slats .................. VREF + 10

10. Vertical Speed ............................................. APPROX

If possible, do not exceed a 300-foot-per-minute ratapproach for landing.

The landing distance will be twice the nominal40° flaps + slats landing distance.

AFTER TOUCHDOWN

CAUTION




1. Full Reverse Thrust .....................................................

In the event of a total hydraulic failure of the main sybe accumulator hydraulic pressure available for at leathe thrust reverser.

2. Parking Brake..................................... INTERMEDIA

If the #2 P. BK light is not flashing, there should bstored in the parking brake accumulator circuit for uof the parking brake. Flashing of the #2 P. BK ligmore brake application is possible. Be careful when dbrake handle to the first detent. Be sure the aircraft ipossible down the runway before pulling the brakelimit the number of pulls on the brake handle to one,of the stored hydraulic pressure.

CAUTION

Pitch attitude is controlled essentiallyizontal stabilizer.

Loss of hydraulic power causes an upwof both ailerons, which produces a slighment that is easily counteracted with If the aircraft is not perfectly trimmedhydraulic power is lost, it is difficulwings level while accelerating to 260case, it is best to decelerate, recentethen reaccelerate.

The rudder is the most maneuverablface when hydraulic power is lost. Thbe turned by induced roll using the ru

At speeds below 200 knots, the aircra




p ,ability improves.

Approach is made at a shallow angleclean configuration at VREF + 30 knot

LOSS OF ALL THREE GENERAT

NOTE

Simultaneous illumination of these thdicates that the three engine generators connected from the main DC buses. Tswitches may be tripped off.

1 B V l

GEN 1 GEN 2 GEN 3AND

If the generators have tripped, and if the batteries are thetrical power:

3. Battery Load......................................... SHED AS M

Turn off nonessential electrical equipment to conMonitor the voltmeters and ammeters while athe generators.

To limit the battery load, accomplish the foin the sequence shown:

a. Booster Pump Switches (3)....................................

b Windshield Heat (3)

CAUTION




b. Windshield Heat (3) ...............................................

In icing conditions, it is imperative to keep the followin

• Pilot Windshield Heating System

• Engine and Wing Anti-icing System

4. Pilot and Copilot Pitot Heater .....................................

Do not switch off the standby pitot heat.

5. Right Avionics Master.................................................

6. Lavatory and Galley Master Circuit Breakers ............

7. Unnecessary Lights.....................................................

A h l h f il d

10. Power Lever (of Engine Concerned) ................

11. Engine at Idle Setting.......................................

The reset attempt should be deferred until the idle power to enhance the reset attempt.

12. GEN Switch (of Generator Concerned)............

There are two alternatives to follow. Follow alt

Carefully observe the voltmeter and ammeter w

Alternative 1




If the voltmeter deviates to maximum attempting t

1. GEN Switch (of Generator Concerned) ...........

Turn and leave the GEN switch off should an o

2. Bus-Tied Switch ...............................................

Before tying the buses, check the ammeter anthe buses together unless a short is detected onthat the bus-tied light is on. If a short is detectbuses. The battery supplying that bus wilprobably should be turned off.

3. Reduce the electrical load to the minimum req

Switch off as many systems as possible to m


Land at the nearest suitable airport for inspection and

Alternative 2

If voltmeter is within the green band:

1. Power Lever (of Generator Concerned).....................

2. Bus-Tied Switch ..........................................................

Before tying the buses, check the ammeter and voltmnormal values. Check that the bus-tied light is on.




3. Turn on previously shed items.

4. Monitor voltmeter and ammeter indications.

After load-shedding, the batteries in good condition

• 40 minutes of operation with an average load otery in nonicing conditions.

• 20 minutes of operation with an average load otery in icing conditions.

NOTE

The following items are recommended foshedding to help prolong battery life even lo

Cabin Entrance Light ......................................

Do not pull the fire warning circuit bre

following left-panel circuit breakers:

• Navigation section:

• AFCS 1, (4)

• RAD ALT 1, ALT 1, FMS 1, C

• IRS 3 BAT, IRS 3

• R/T WR, MFD/WRD

• SG 3

• GPWS

• Radio section:

• HF I (2)

CAUTION




• HF I (2)

• FLITE FONE

• ADF 1

Pull the following center-panel circui

• Lights, miscellaneous sect ion:

• RH CABIN READING

• LH CABIN READING

If in icing conditions, do not pull the ers for anti-ice of the engines and wing.do not pull the circuit breaker for the

shield heat.

CAUTION

In Icing Conditions

The load corresponds to a total output from the batteries othat ensures 25 minutes of operation with two 23-Ah batte

of their capacity.

During Approach

During the last five minutes of flight, the load is increabe considered that two 23-Ah batteries, charged at 75% ovide sufficient reserve power in the following condition

• In nonicing conditions, for 57 minutes of operationutes during approach

or




• In icing conditions, for 24 minutes of operation,

utes during approach

RAPID DEPRESSURIZATION

Phase 1

1. Crew Oxygen Masks ..................................................

The crew oxygen masks must be put on immediatesmoke. Be careful of eye glasses when putting onpersonal injury or to avoid a poor fit of the maskThe mask must fit tightly around the face to enspenetrate the mask and enter the breathing passaNORMAL-100% oxygen selector button on the fro

the 100% position.

from their storage bins at each passenger pooxygen when the passenger pulls the mask dow

5. Emergency Descent..........................................

Simultaneously, as the airbrake is deployed anto idle, begin the descent by rolling the aircraftwill expedite entry into the descent without advthe aircraft. The initial descent angle shounosedown pitch attitude. After the 20° nosedowremove all bank, hold the 20° pitch down unti

then adjust pitch to approximately 10° of nomaintain the aircraft at MMO /VMO in the descen

EMERGENCY DESCENT




This procedure assumes normal structof the aircraft. If structural integrity islimit the airspeed to the lowest practiavoid high maneuvering loads.

1. Autopilot ..........................................................

Disengage the autopilot, and hand-fly thprocedure. Disengagement is achieved by prethe control wheel. Check for the AP light oagain to clear the disengagement message on th

2. Power Levers ....................................................

It is advisable to turn on the airstart ignitioi fl b f di

CAUTION

then adjust pitch to approximately 10° of nosedownmaintain the aircraft at MMO /VMO in the descent.

5. Airspeed (Smooth Air)................................................

Lead the desired level-off altitude by at least 2,000 fedescent to a maximum of 2,000 feet per minute. Lealtitude at an altitude which is 10% of the rate of 14,000 feet, the assigned altitude, or to the minimumroute flown.

6. Transponder ................................................................

Advise ATC of your emergency and set the transpto ensure identification, priority handling, and a safethis emergency.




FORCED LANDINGPREPARATION

1. Communications Transmission............................. TR

On the most recently used communications radio, tthe monitoring communications agency. If you havecommunications agency on an assigned frequency, owhat the correct frequency is, change to 121.5 Mdistress message on that frequency.

2. Transponder ................................................................

Unless directed otherwise, set the transponderemergency distress code.



gear cannot be extended, it is desirable that as manextended to help cushion the landing. A belly landing

12. Flaps + Slats..............................................................

If possible, the maximum flap setting should be usedslow as possible for the landing.

13. Approach Speed ..........................................................

With the flaps/slats set at 40° FLAPS + SLATS, flythe charted VREF speed.

JUST BEFORE TOUCHDOWN






If possible, control the rate of descent to a maximminute to minimize impact forces.

15. Fuel Shutoff Switches (3) ...........................................

On the FIRE PANEL, raise all three guards and rswitches up to shut off the fuel supply to the engi

fuel tanks.

16. GEN Switches (3) .......................................................

Move all three generator control switches to off to electrical power supply to the buses.

17. BAT Switches (2) ........................................................



2. Transponder ................................................................

Unless directed otherwise, set the transponderemergency distress code.

3. Passenger Emergency Briefing....................................

If carried, use the cabin attendant to prepare theemergency. If a cabin attendant is not carried, use thethe passengers on their roles during this emergency. Thave already been prebriefed, prior to departureemergency procedures and the location and use of em

4. Life Jackets ............................................................. D

Direct the passengers and crew to remove the life jacput them on Life jackets are not to be inflated




put them on. Life jackets are not to be inflatedpassengers evacuate the aircraft.

5. Fasten Belts/No Smoking Signs .................................

All cabin and cockpit occupants must have their seatand, where applicable, have shoulder harnesses plocked. Smoking is prohibited throughout the aircraft

6. AFT CABIN ISOL Light ............................................

On aircraft with the midcabin separation door, this dthe AFT CABIN ISOL light must be out.

7. Cockpit Jump Seat (If Possible) ............. UNOCCUP

APPROACH—PARALLEL TO THE MAJ

9. Wing Anti-ice...................................................

This is one of many procedures designed to cloof the aircraft before ditching. This, and otheprovides airtight systems and subsystems for echaracteristics of the aircraft.

10. Crew and Passenger Bleed Switches................

This is done to further isolate the cabin from eto prevent pressurization of the cabin.

11. BLEED AIR BAG Switch................................

This will isolate the cabin and baggage co




This will isolate the cabin and baggage cobaggage compartment heating valve and the c

BAG ISOL light on.

12. HP 1, PRV 2, and PRV 3 Switches ..................

This is done to further isolate engine bleed-air

13. Pressurization Switch.......................................

On the emergency pressurization control panDUMP switch, and move the switch down tThis has a dual purpose:

• To release any residual pressure in the copening of the emergency exits after di

• To release any residual pressure in the c

16. Approach Speed..........................................................

With the slats/flaps set at 40° FLAPS + SLATS, flythe charted VREF speed. Because the gear is not down

warning horn will sound and cannot be silenced unleA and the audio warning B circuit breakers are pulled

JUST BEFORE TOUCHDOWN

17. Rate of Descent ........................................... APPROX

If possible, control the rate of descent to a maximminute to minimize impact forces.

18. Fuel Shutoff Switches (3) ...........................................

On the FIRE PANEL, raise all three guards and ri h h ff h f l l h i




switches up to shut off the fuel supply to the engi

fuel tanks.

19. GEN Switches .............................................................

Move all three generator control switches to off to electrical power supply to the buses.

20. BAT Switches (2) ........................................................

This removes the remaining electrical power supply to prevent electrical sparking and fire. The aircrapowered except for any emergency battery poweshould be turned off to prevent any fire after ditching

21. Contact the water on the crest, parallel to the swelattitude of 12° to 15 5°

23. Cockpit Jump Seat ...........................................

If it has not already been stowed, stow it toevacuation.

24. Emergency Exit................................................

Open the emergency exit located over the riaircraft. Deploy the life line to the anchor poithe gravity-fueling cap. Inflate the life rafts onaircraft. The life raft may be secured to the rinwing to keep it in place until all occupantpassengers should not inflate their life vestswindow or door.

Selection of the proper emergency exit, either wing or the main cabin door, should depend o

l i i




total situation.

25. Main Cabin Door .............................................

An alternate means of evacuating the aircrafdoor, dependent upon the attitude of the airalternative for evacuation is the pilot’s direct-viused as a last resort if all other exits fail to open

26. Evacuate the aircraft.

LIMITATIONS AND SPEC

CONTENTS

LIMITATIONS......................................................

General........................................................

Weight—Structural.....................................Weight—Performance ................................

Center of Gravity ........................................

Loading.......................................................

OPERATING LIMITATIONS




OPERATING LIMITATIONS..............................

Kind of Operation.......................................

Altitude .......................................................

Maneuvering Flight Load Factors................

Takeoff and Landing...................................

Minimum Flightcrew...................................

Maximum Number of Passengers...............

Airbrakes.......................................................

Airspeed......................................................

SYSTEM LIMITATIONS .....................................

Fuel System ................................................

Hydraulic ..............................................................

Electrical...............................................................

INSTRUMENT MARKINGS........................................

Airspeed Indicator Markings................................

Engine Instrument Markings ................................

Miscellaneous Instrument Markings ....................

F




ILLUSTRATION

Figure Title

LIM-1 Temperature and Altitude Limits ....




LIMITATIONS AND SPEC

LIMITATIONSGENERAL

The limitations presented in this chapter focus pricapabilities of the aircraft. Specific system limitatidividual systems’ chapters with the exception of inare presented in this chapter. Refer to the FAA-ap

limitations listings.

WEIGHT—STRUCTURAL

Maximum ramp ...........................................................

-5AR 45 700 lb or




5AR .....................................................45,700 lb or

-5BR ..........................................................................

Maximum takeoff ........................................................

-5AR .....................................................45,500 lb or

-5BR ..........................................................................

Maximum landing .......................................................

Maximum zero fuel .....................................................

-5AR .....................................................28,200 lb or

-5BR ..........................................................................

Minimum flight weight .............................................

• Brake energy

• Climb gradients

Landing

The landing weight is limited by the following:

• Approach and landing climb gradients

• Available landing field length

CENTER OF GRAVITY

General

The center-of-gravity limits are expressed in percent odynamic chord). The landing gear position has no effe

F




y ) g g p

gravity. Refer to the center-of-gravity limits chart insection of the AFM .

Datum

Datum is 25% of MAC; it is marked on the aircraft exteriofuselage station (FS) 420.43 inches (10,679 mm). FS 0 i

the aircraft nose cone.

Mean Aerodynamic Chord

Length is 113.69 inches (2,887.7 mm).

Zero percent MAC is at FS + 392 inches (9,957 mm).

OPERATING LIMITATIONS

KIND OF OPERATION

This aircraft is certified in the transport category lowing kinds of operations when the appropriate inrequired by the authorities and/or operating regulaproved and are in operable condition :

• Day and night VFR, if permitted by the reguwhich the aircraft is flying

• IFR and automatic approaches to category I

• Extended overwater

• Icing conditions

• The overflight of polar regions is limited to




less than 85°

• The overflight of polar regions is authorizeFMS computer software is identified 9102

• Flight in the former USSR airspace:

° The aircraft is not allowed to fly on route

secondary radars operating in UVD mo

° For aircraft without M1846: the aircraequipped with VOR/DME . The VOR/Done hour and 20 minutes on routes whwidth, and two hours and 40 minutes on(5.4 NM) width.

° For aircraft with M1846 ( equipped with

RVSM

Reduced vertical separation minimum (RVSM) requiremeaircraft complies with SB F900-186.

In addition to SB F900-186, specific approval from the ity is needed prior to RVSM operation .

NOTE

In normal operation for RVSM areas, select the coupled side.

Minimum equipment list for RVSM operations is provid

RNP 10

In accordance with FAA Order 8400.12 A, paragraph 12 B, Rrequirements are met provided aircraft is equipped with d

F




q p q pp

• FMS NZ2000 software 4.1 or later and either of th

° GPS

° IRS (6.2 hours after last aligment or 5.7 hours a

OR

• FMS FMZ 800 series and the following mode:

° IRS (6.2 hours after last aligment or 5.7 hours a

OR

• FMS NZ 920 and the following mode :


B-RNAV

Basic RNAV(B-RNAV) airworthiness requirementis equipped with:

• FMS HONEYWELL FMZ 800 or FMZ 920or DGRAD is present on FMS CDU, and eiigation mode:

° GPS type HG 2021 GB/GD

° DME/DME

° VOR/DME

° IRS (two hour time limit after last IRS

NOTE

• When GPS remains the unique mea



When GPS remains the unique mea

navigation source (GPS stand-alonintegrity monitoring (RAIM) prediis mandatory before B-RNAV ope

• GPS stand-alone not authorized for and FMZ 920.

• At least one VOR/DME must be avasource (DC 820) on PFD.

P-RNAV

Precision RNAV (P-RNAV) airworthiness requiremJAA TGL 10 provided aircraft is equipped with:

• FMS HONEYWELL FMZ 920 or FMZ 200DGRAD warning on FMS CDU and either

NOTE

Select FMS approach procedure must not beally modified.

US terminal and enroute area navigation (RNAV) operatioAC–100 airworthiness requirements are met provided aircr

• FMS HONEYWELL FMZ 920 or FMZ 2000 ope

° GPS

° DME/DME

° VOR/DME

° Navigation mode without any DR or DGRANOTAM navaid entered in the FMS NOTAM

R NAV airworthiness approval has not accounted for datab

F




patibility.

RNP flight operations are subject to GPS satellite availacoverage for the selected route. Navigation based on DMEupdating modes is permitted but may be restricted by theformance of the applicable ground navaid. Crew should ground navaids that are not to be used for navigation.

ALTITUDE

Maximum operating altitude is 51,000 feet (Figur


51

34

50

40

30

20

44

ALTITUDE (x 1,000 ft)

ISA



10

0

TAKE-OFF AND

LANDING

-75

-80

-54

-60 -40 -20 0

STATIC AIR TEMPERATURE

Figure LIM-1. Temperature and Al

Demonstrated crosswind ......................................................

Tailwind component at takeoff:

Aircraft fitted with tires for 210 mph:

Pressure altitude of airport from –1000 to 10,000

Pressure alt itude of airport more than 10,000 f t ....

Aircraf t fitted with 6 t ires approved for 225 mph .....

Tailwind component at landing:

Aircraft fitted with tires approved for 210 and/or 22

Ambient temperature ......................................... Refer to taltitude lim

“Limitations”

Runway surface .......................................................... Paved

F




MINIMUM FLIGHTCREWThe minimum flightcrew is one pilot and one copilot.

MAXIMUM NUMBER OF PASSENGERS

The maximum number of passengers is 19.

AIRBRAKES

Airbrakes must not be extended in flight within 300 feet

AIRSPEED



Do not intentionally fly the aircraft slower tinitial stall warning onset.

Maximum Landing Gear Operating Speed (

VLO ..........................................................................................

MLO ..........................................................................................

VLO /MLO is the maximum speed at which it is safe to e

landing gear.

Maximum Landing Gear Extended Speed (V

VLE ...........................................................................................

MLE ..........................................................................................

CAUTION

F




VLE is the maximum speed at which the aircraft can be slanding gear extended and locked.

Minimum Control Speed (VMCA)

VMCA (A/C with TFE 731-5AR-1C) ...................................

VMCA (A/C with TFE 731-5BR-1C)...................................Demonstrated Crosswind .....................................................

Miscellaneous Limit Speeds

Windshield wiper operating.................................................

Direct vision window opening

SYSTEM LIMITATIONS

FUEL SYSTEM

The total usable fuel quantity is 2,845 U.S. gallons pounds (8,648 kg) at a density of 6.7 pounds per U

This total quantity is distributed as shown in the f

Fuel used must conform to the following specificat


LITERS

Left wing and left centerwing tanks 3,422 2

Right wing and right centerwing tanks 3,422 2

Front and rear fuselage tanks 3,925 3

Pressure fueling maximum feed pressure: 50 psi/3.5 bars/350



g pis representative of the fuel definition on the date:

DesignationALLIEDSIGNAL

Specification

KEROSENE

WIDE CUT

EMS 53111

EMS 53112

EMS 53113

Equivalence(for info.)

ASTM D 1655CAN 2-3.23

ASTM D 1655CAN 2-3.23DERD 2494DERD 2453

MIL-T-83133AIR 3405CAIR 3405 C

ASTM D 1655CAN 2-3.22

MIL-T-5624

JET AJET A

JET A1JET A1AVTUR

AVTUR/FSii

JP8––

JET BJET B

JP4

Freezinpoint(°C)

– 40

– 47

– 50

– 50

POWERPLANT

Garrett TFE 731-5AR-1C Engine

Thrust Ratings (Uninstalled, Sea Level, ISA):Takeoff ............................................................................... 4,

Maximum continuous ...................................................... 4,

Thrust Setting

The engine low-pressure rotor speed N1

is used as the thru

• Takeoff thrust (5-minute time limit)

Maximum Engine Rotor Speeds—N1 and N2

F


CONDITION OF USE N1



Maximum Interstage Turbine Temperature: ITT

Takeoff—Maximum continuous 100%*

Transient (5 seconds maximum allowable) 103%

* 100% N1 = 21,000 rpm * 101% N2 = 29,989 rpm

Starting, ground/air Normal

Takeoff (with increased thrust) Normal (5 minutes maxim

Takeoff (without increased thrust) Normal (5 minutes maxim

Takeoff Transient Normal (5 seconds maxim

Starting Time

Fuel Control Computers

The engine fuel control computers must be operat

Thrust Reverser

The thrust reverser is approved for ground-use on

Garrett TFE 731-5BR-1C Engine

Th R i (U i ll d S L l IS


Groundstart and starter-assist airstart— from 10% N2 speed to light-off

Windmilling airstart— from windmilling N2 speed to 60% N2

Groundstart— from light-off to idle



Thrust Ratings (Uninstalled, Sea Level, ISTakeoff .........................................................................

Maximum continuous ................................................

Thrust Setting

The engine low-pressure rotor speed N1 is used as th

• Takeoff thrust (5-minute time limit)

Maximum Engine Rotor Speeds—N1 and N

CONDITION OF USE

Generator Load:

To 43,000 feet ........................................................................

Above 43,000 feet ................................................................

One minute transient ............................................................

Starting Time

F l C t l C t

F


Groundstart and starter-assist airstart— from 10% N2 speed to light-off 10 seco

Windmilling airstart— from windmilling N2 speed to 60% N2 45 seco

Groundstart— from light-off to idle 60 seco



Fuel Control ComputersThe engine fuel control computers must be operative for

Thrust Reverser

The thrust reverser is approved for ground-use only.

LUBRICATION SYSTEM

Approved Oils

Type II oils are Aeroshell/Royco Turbine Oil 500 andExxon/Esso 2380 Turbo Oil, Mobil Jet Oil II and Mobiwith AlliedSignal Engines EMS 53110 type II.

Oil Temperature

SYSTEMSAUXILIARY POWER UNIT (APU)GARRETT GTCP 36-150 (F)

The APU must be operated on the ground only.

Operation of the APU with passengers in the cabin


OPERATIONAL LIMITS

From sea level to 30,000 feet Above 30,000 feet

Transient all altitudes 149°C m

Minimum for exceeding idle power



Operation of the APU with passengers in the cabinitoring is not authorized.

Maximum N1 speed....................................................

NOTE

T he du r a t i on o f ope r a t i on on a(679°C/732°C to 1,255°F/1,350°F) mu

as possible

EXHAUST GAS TEMPERATURE

Starting Between 870° and 985°C (1,

less than 10

Stabilized 679°C (1

AUTOMATIC PILOT (SPERRY DFZ 800)

The autopilot must not be engaged for takeoff or landing

The autopilot is certified to the minimum height as follo

Mininum height radio altimeter operative ........................

Mininum height radio altimeter inoperative......................

Minimum decis ion height ....................................................

Minimum height for autopilot

operation, except during approach .....................................Minimum height for use during an FMS approach ...........

On aircraaft equipped with FMS computer s

9004 before use of the APP mode the modb di d b f APP d

CAUTION

F




9004, before use of the APP mode, the modmust be disengaged before APP mode engag

ANTI-ICE

Icing Conditions

Icing conditions exist when the OAT on the ground TAT in flight is 10°C or below, and visible moisture in (such as clouds, fog with visibility of one mile or less, raice crystals).

Icing conditions also exist when the OAT on the ground anor below when operating on ramps, taxiways or runways ice, standing water or slush may be ingested by the engi

gines nacelles or engine sensor probes

Conclusion

During climb and cruise, the pneumatic anti-ice sy

• Below +10°C (50°F) TAT and above –40°C

and

• In vis ible moisture.

If both of these conditions are not met, the ant i- i

Wing Anti-iceThe wing anti-ice system must not be used with tocess of +10°C. It must not be used on ground exceconducted in accordance with Maintenance Manu

BAGGAGE COMPARTMENT

The lavatory baggage compartment door must be




The lavatory-baggage compartment door must be any operation above 41,000 feet.

CABIN PRESSURIZATION

Maximum differential pressure ............................ 9

(pre

HYDRAULIC

Hydraulic fluid approved for use must conform to M(NATO codes H515 for H520).

After SB F900-94-1:

Amber light (WARM) at or above............................

Red light (HOT) at or above .....................................

NOTE

Any popped CB can be reset only if it is lesamps strictly rated.

INSTRUMENT MARKINGS

AIRSPEED INDICATOR MARKINGS

F




ENGINE INSTRUMENT MARKINGS

Instrument Color Codes

Maximum operating limit ....................................................

Precautionary range .......................................... Amber or

Normal operating range.......................................................

N1 RPM

TFE 731-5BR-1C ........................ 968° to 996°

Red line:

TFE 731-5AR-1C ..................... 952° and 974°TFE 731-5BR-1C ..................... 978° and 996°

Blue test point:

TFE 731-5AR-1C .................................. 1,000°

TFE 731-5BR-1C .................................. 1,040°

N2 RPM

Green arc................................................ 48% to 100

Red trapezoid ...................................... 100% to 103

Blue test point .................................................... 106

Oil T d P




Oil Temperature and Pressure

Temperature

Green arc (sea level to FL 300)........... 30° to 127°

Yellow arc (above FL 300) ................ 127° to 140°

Pressure

Red l ine (minimum at idle) .............................. 25 p

Yellow arc (idle range) ........................... 25 to 38 p

Green arc (normal operating range)...... 38 to 46 p

Yellow arc (transient

[maximum < three minutes]) 46 to 55 p

MISCELLANEOUSINSTRUMENT MARKINGS

Fuel Quantity

Left and Right Quantity Indicators

Yellow arc ............................................... 0 to 1,000 lb

Center Quantity Indicator

Aircraft with electric transfer valve XTK2:

Yellow arc ....................................... 0 to 1,000 lb

Yellow arc ............................... 2,200 to 2,400 lb

Green arc ................................. 3,100 to 3,500 lb

Green arc ................................. 4,300 to 4,500 lb

Aircraft without electric transfer valve XTK2:

Y ll 0 1 000 lb

F




Yellow arc ....................................... 0 to 1,000 lb

Cabin Pressure

Cabin Rate of Climb

Green arc .................................... –495 to +715 ft/min

Cabin Altitude

Yellow arc ...................................... 8,000 to 10,000 ft

Red arc ......................................... 10,000 to 50,000 ft

Pressure

Green range ................................... 2,800 to 3 ,200 p

Electrical System

Battery Temperature

Green range .......................................... 100° to 120°

Yellow range ........................................ 120° to 150°

Red range.............................................. 150° to 190°

Aircraft above SNs 132 and aircraft below 132 witSB-94 applied:

Yellow range ................................ 120° to 160°

Red range ...................................... 160° to 190°




DC Voltmeter

Green arc ............................................... 24 to 26 vol

Green arc ........................................ 28.2 to 28.8 vol

Maximum vol tage .......................................... 32 vol

Ammeter—Maximum Load

Yellow triangle ............................................ 250 amp

Red line......................... Approximately at 300 amp

Flight Control Trim

Aileron

White range ..................................................... 0 to 5%

Rudder

White range................................................. –5 to +5%

Stabilizer

Green range ............... –4° 30' aft to –7° 30' forward

Oxygen

Red arc ...................................................... 0 to 200 psi

Yellow arc ............................................ 250 to 700 psi

White arc ........................................... 700 to 2,000 psi

F




pYellow arc...................................... 2,000 to 2,200 psi

APU

N1 RPM

Green arc................................................ 95% to 105%Yellow arc............................................ 105% to 110%

Red line ............................................................... 110%

T5

Temperature



ILLUSTRATIONS

Figure Title

MAP-1 Takeoff Flight Path ..............................

MAP-2 Typical ILS Approach—Three Engines

MAP-3 Typical Nonprecision Approach—Thre

MAP-4 Typical Circling Approach—Three Eng

MAP-5 Missed Approach—Three Engines ......

MAP-6 Typical ILS Approach—One Engine InMAP-7 Typical Nonprecision Approach—

One Engine Inoperative........................

MAP-8 Circling Approach—One Engine Inope

MAP-9 Missed Approach—One Engine Inoper




MANEUVERS AND PRO

AIRWORK MANEUVERSTAKEOFF WITH ENGINE FAILURE AFT

If an engine failure occurs after V1, during the takeoff using the rudder to maintain directional contr(Figure MAP-1). Perform the following procedure

1. Rotate at the charted VR

speed to the approp20/7 + slats.

2. Maintain the charted V2 speed throughoutments of climb. If a speed slightly higher is attained, accept that speed and climb out opitch attitude in order to lower the speed.

3. When a positive rate of climb is establishe

altimeter and vertical velocity indicators, 4 At the altitude selected to retract the flap




4. At the altitude selected to retract the flap400 feet above the airport elevation or 35 fewhichever is higher, begin acceleration to

5. Retract the slats and flaps, accelerate to the(1.43 VS), reduce the power on the operatingtinuous thrust, and climb at 1.43 VS to 1,5

elevation.6. Engine shutdown procedures may commenc

climb if the situation allows.

AIRWORK—STEEP TURNS

Steep turns may be accomplished at altitudes of 5,0

d hi h t i ltit d f 18 000 f t D

F OR T RAI NI N G P URP O SE S

ONL

Y

M

AP - 2

SLATS /FLAPS—TAKEOFF POSITION

LANDING GEAR—UP

TAKEOFF GRADIENT 2.7% MIN.

SPEED—V2

SLATS /FLAPS—TAKEOFF POSITIONLANDING GEAR—DOWN

TAKEOFF GRADIENT .3% MIN.

SPEED—V2

GROUND ROLL

REF 0

TAKEOFF THRUST

(5 MINUTES MAX)

RETRACTION IF

ACCELERATELANDIN

LANDING

GEAR—UP

2 N D

S E G M E N T

1 S T

S E G

M E N

T

35 FT



35 FT

V1

VR /V

2

Figure M

turn. This type of recovery allows the nose of thehorizon with a positive G condition, thus avoiding G forces.

Nosedown maneuvers will be given by placing tbank with the nose down no more than 30° beloshould be smooth, reducing power to idle, levelnoseup elevator to ease the nose to the horizon. Tas required.

Proper evaluation of instrument presentations isVerify attitude indicator presentat ions with the cop

zon. If the attitude indicator is inoperative, evaluatealtimeter, and vertical velocity indicators.

COORDINATION MANEUVER

The objective of this maneuver is to sharpen the cooof the pilot while flying Dassault Aircraft and Lev

neuver is especially helpful for pilots undergoing iaircraft or simulator. It is also useful for pilots un




pto assist in the application of coordination skills tosimulator.

For initial training pilots, this maneuver is perforing ride in the Level C simulator; accomplishmrecommended but is optional dependent upon indi

This maneuver is not a specific requirement of thethe aircraft manufacturer. It is designed for trainingperience, has proven to be most effective in enhansimulator controllability of pilots in training.

This maneuver is normally accomplished after theprior to the stall series. It is accomplished at the sf h ll i I i d f d f


configuration. While in the landing configuratiengine instruments to preclude exceeding limitatnecessary to retract one notch of flaps to prevel imi t a t i ons . Af t e r t he l and ing conf igura t iincrementally retract gear, flaps, and leading-edgeceeding turn reversal until the clean configurestablished. Maintain assigned altitude, 160 kno15° angle of bank left and right on the 60° arc.

5. The maneuver is completed after reaching the cleaone subsequent turn reversal.

Performance criteria are:

• Angle of bank ...........................................................

• Airspeed ....................................................................

• Altitude......................................................................

• Power limits ..............................................................

APPROACH-TO-STALL SERIES



General

The approach-to-stall series should be performed at an alfeet above ground level and below flight level 180. To appditions and satisfy FAA requirements, the stall series isthe clean, takeoff, and landing configurations. These proped for training use in the aircraft and Level C simulat

Emphasis must be placed on proper approach-to-stall recopositive aircraft control, and proper recovery procedure

Proper recognition of an approaching stall regime canobservation of any one of the following indications:

Clean Configuration

This maneuver is performed while turning in a coAirspeed at start of maneuver is 190 knots. MaintLevel C simulator, this may be performed at normto present realism.

1. Power setting—50% N1

2. Trim—Do not use stabilizer trim below an above expected stall indication speed.

3. Stall indication normally occurs at a speed

4. Recovery—At the first indication of stall:a. Maximum power—Smoothly advance

power setting. Copilot monitors limita

b. Simultaneously and smoothly roll thelowers the stall speed.

c. Smoothly maintain pitch only enoughor buffet and minimize altitude excurssigned altitude be maintained throughabrupt control movements to preclude




abrupt control movements to preclude

d. Accelerate to and hold 190 knots airsp

Takeoff Configuration

This maneuver is performed while turning in a copreferably in a direction opposite from the one usver. Airspeed at start of maneuver is 190 knots. MIn a Level C simulator, this may be performed altitudes to present realism.

1. Landing gear—Up

2. Flaps/leading-edge devices—Set flaps 20°


Landing Configuration

This maneuver is performed while holding a constant hstart of maneuver is 190 knots. Maintain assigned altitudthe aircraft and setting power.

1. Landing gear—Extend in normal landing sequen

2. Flaps/leading-edge devices—Set flaps +40° + sl

3. Power setting—After landing configuration is es

4. After the power is set, allow the aircraft/simulatospeed. Descend on VREF to a simulated decision

500 feet below the originally assigned altitude. Wbitrary DH, level the aircraft/simulator at this aadjusting the power, allow the aircraft/simulatoindication of a stall. In a Level C simulator this main the traffic pattern while on final instrument afinal approach fix.

5. Trim—Do not use stabilizer trim below an airspe

above expected stall indication speed.6. Stall indication normally occurs at VREF –20 kno



7. Recovery—At the first indication of stall:

a. Maximum power—Smoothly advance throttlesetting. Copilot monitors limitations.

b. Simultaneously retract flaps one notch to 20

c. Smoothly maintain pitch only enough to stoor buffet and minimize altitude excursions.

d. As a irspeed increases to VREF, execute norcedures: Climb on VREF to the original altituing gear upon a positive indication of climb,and leading-edge devices in accordance withprocedures. VREF may be exceeded in the cliexceeding a 20° angle. In a Level C simu

i d h d

EMERGENCY DESCENT

This maneuver affords training in the recommended the highest practical rate of descent possible during ing from an uncontrollable fire, sudden loss of pr

situation dictating an immediate and rapid descen

This procedure assumes structural integrity of the apect, reduce the rate of descent and avoid high loa

The following procedure is prescribed by the AFM

1. Disengage the autopilot.2. Retard power levers to idle .

3. Extend the airbrakes to position 2.

4. Descent airspeed: MMO /VMO, smooth air c

5. ATC Transponder, MAYDAY code—Squaw

6. Ensure that landing gear is up.

7. To start the descent, the initiation of a 45°to the descent without adversely affecting




to the descent without adversely affectingcraft. The initial descent angle should be 2ADI. After the 20° nosedown is achieved, sary. Contnue in a 20° nosedown pitch angle At this time, a reduction to a nosedown pitsure maintenance of VMO /MMO throughou

off altitude.8. At 1,000 feet above the level-off altitude

angle to achieve a vertical velocity of notminute. Retract the airbrakes, if necessary, off altitude. Begin level-off to assigned altitvertical velocity in the descent.

INSTRUMENT APPROACH PROF


4. Prior to crossing high station outbound, slow the in the clean configuration. If using the flight directhe heading mode should be selected on the flight duntil the aircraft is aligned within 90° of the inbaircraft is proceeding inbound for course interce

5. At high station outbound, set the slats/flaps to Sthe Approach checklist. Slow the aircraft to 150ver through the procedure turn as prescribed on

6. When inbound from procedure turn and establishcept (normally a 45° intercept heading), select ththe flight director control panel (if flight direc

proach). Intercept the final approach course, andthe final approach fix or glide-slope intercept poaltitude.

7. When the glideslope is alive and reaches 1 dot lowgear down, and call for the Before Landing chec

8. When intercepting glideslope, extend the flaps tofinal approach speed of VREF plus the wind correc

wind plus the full gust, not to exceed 20 knots).slope with elevator and airspeed with minor powpilot making the approach will handle the throttl



pilot making the approach will handle the throttl

9. Cross-check instruments at 1,000, 500, and 100 feeimums. The pilot making the approach should remthroughout the final approach. When a visual cestablished and a safe landing can be made, as vthe pilot not flying the approach, continue to thelanding.


WING — CLEAN

SPEED — 190 KTS WING — SLATS +20°

SPEED — 150 KTS

(APPROACH CHECKLIST)

LEGEND

IN THE FOLLOWING APPROACH FLOW CHARTS, THE APPROACH HA

AREAS. FIRST THE MANEUVERING AREA WHEREIN THE SPEED FOR

THE PILOT TO USE STANDARD RATE TURNS FOR MANEUVERING W

OND THE NONMANEUVERING OR FINAL APPROACH AREA WHICH IS

STABILIZED IN THE APPROACH OR LINED UP WITH THE RUNWAY, SOQUIRE ONLY ASMALL BANK ANGLE.

THE MANEUVERING IS DEPICTED ON CHARTS A

THE NONMANEUVERING OR FINAL APPROACH I

SHADED AREA.



WING — SLATS

SPEED — VREF

+ W

AT THE FINAL FIX*


NONPRECISION APPROACH—THREE ENG

Figure MAP-3 illustrates the Nonprecision Approach—

1. Review the approach chart for the approved proce

minimums. Conduct the nonprecision approachthese charted procedures and as prescribed by tagency.

2. The Descent checklist should be accomplished papproach. Set all instruments for the approach, approach briefing has been completed prior to bapproach.

3. Tune and identify the VOR and/or ADF radios tthe approach. Heading and course values will be instruments in preparation for the approach. Epointers have been properly selected.

4. Prior to crossing high station outbound, the aircrto 190 knots in the clean configuration.




g p p

6. At the completion of procedure turn and prior tofix (FAF), select the landing gear down, and cLanding checklist. Maintain the airspeed at 150 FAF at this speed with all checklists accomplish

7. At the FAF, start the timing; if the missed approabegin a descent to the minimum descent altitudeto exceed 1,000 feet per minute and at a speed of 15is reached, fly the aircraft at MDA until visual cway is made or the missed approach point (MAP)contact is achieved and landing is assured, extendslow the aircraft to the VREF speed plus the winding. If not visual prior to reaching the designate

bli h d i d h d


WING — CLEAN

SPEED — 190 KTSWING — SLATS +20°

SPEED — 150 KTS




WING — S

SPEED — V

LANDING ASSURELINED UP WITH RU

START TIMING..........................

SINK RATE — 1,000 F.P.M.

SPEED — 150 KTS

AT THE FINAL FIX

MINIMUMS


CIRCLING APPROACH—THREE ENGINES

Figure MAP-4 illustrates the Circling Approach—Three

Although the Falcon 900 is considered a Category C airc

inition as obtained from the TERPS Manual, it is recommD minimums be used for circling to an airport. This affoand .6 of a mile greater maneuvering airspace.

The typical three-engine ILS and nonprecision approacexcept that circling approaches are flown to published circslats/flaps set at S +20° and at an airspeed of 150 knotscircle for landing, and on the downwind leg, select the and complete the Before Landing checklist. Maintain 15normal VFR landing pattern procedures.

When on final approach, at or above 500 AGL, and slow to Vwind corrections.




WING — CLEAN

SPEED — 190 KTS

RECOMMENDED U

WING — SLAT

SPEED — 15

(APPROACH CHEC

START TIMING..........................


SPEED — 150 KT

S

AT THE FINAL FIX



(AT MINIMU

SPEED

15 SEC

90°


MISSED APPROACH—THREE ENGINES

Figure MAP-5 illustrates the Missed Approach—Threecision height is reached on an ILS approach, or when reacproach point as specified for a nonprecision approach, and

the airport or landing runway cannot be made, execute a maround).

Follow the missed approach instructions as depicted onstrument approach plate used for approach.

Simultaneously apply maximum thrust to the engines, rthe go-around 14° pitch attitude, airbrakes zero, and retrWhen a positive rate of climb is indicated on both the altical velocity indicators, retract the landing gear. At 400 fairport elevation and at VREF +25 knots, retract the slats/power to normal climb thrust. Accelerate to the normal ewhile climbing to the missed-approach or assigned altitu



NORMMAXIMUM THRUST(5 MINUTES MAX)

SLATS—+40° PITCH—14° /FLAPS— +20°

POSITIVE RATE—GEAR UP

CLIMB GRADIENT—3.2%

SPEED—VREF

+25 KT

FLAPS + SLATS—AFTER

TAKEOFF CHECKLIST

SPEED—2+ 10 KT

–0

3. Tune and identify the VOR and ADF radioapproach. Heading and course values willinstruments in preparation for the approapointers have been properly selected.

4. Prior to crossing high station outbound, sloin the clean configuration.

5. At high station outbound, set the slats/flapthe Approach checklist. Slow the aircraft ver through the procedure turn as prescrib

6. When inbound from the procedure turn, oand approaching the glideslope at the 1 dot

landing gear down, and call for the Before Lthe airspeed at VREF +5 knots + the wind 40° of flaps when runway is in sight is optiouse should be judiciously considered in liweather conditions. A stabilized approach fmay be the best choice.




WING — CLEAN

SPEED — 190 KTS WING — SLATS +20°

SPEED — 150 KTS



NONPRECISION APPROACH—ONE ENGINE INOPERATIVE

Figure MAP-7 illustrates the Nonprecision Approach—One

1. Review the approach chart for the approved proceminimums. Conduct the nonprecision approachthese charted procedures and as prescribed by the

2. The Descent checklist should be accomplished papproach. Set all instruments for the approach, approach briefing has been completed prior toapproach.

3. Tune and identify the VOR and/or ADF radios tthe approach. Heading and course values will be instruments in preparation for the approach. Ensers have been properly selected.

4. Prior to crossing high station outbound, slow the in the clean configuration.




6. After completion of the procedure turn, and juapproach fix, select the gear down, and completechecklist. Cross the FAF at 150 knots.

7. At the FAF, start approach timing; descend at a rate

feet per minute, and maintain a speed of 150 kndescent altitude. When MDA is reached, fly the avisual contact with the runway is made or the miis reached.

8. When landing is assured, slow the aircraft to VREFcorrection for landing. An option exists to select landing is assured. This option is based upon the


WING — CLEAN

SPEED — 190 KTSWING — SLATS +20°

SPEED — 150 KTS(APPROACH CHECKLIST)

START TIMING

AT THE FINAL FIX



WING — S

SPEED — VREF

ORWING — S

LANDING ASSURELINED UP WITH RU

START TIMING.......................... SINK RATE — 1,000 F.P.M.SPEED — 150 KTS

MINIMUMS


CIRCLING APPROACH—ONE ENGINE INO

Figure MAP-8 illustrates the Circling Approach—One E

Although the Falcon 900 is considered a Category C aircr

TERPS Manual, it is recommended that Category D micircling to an airport. This affords a higher ceiling and amaneuvering airspace.

The typical two-engine ILS and nonprecision approach profthat circling approaches are flown to published circlislats/flaps set to S +20° and at an airspeed of 150 knestablished on downwind leg for landing. At this pointgear, and accomplish the Before Landing checklist. Mai150 knots until established on final approach. When aligproach to the runway, begin to slow the airspeed to VREFcorrection, if any. Again, the option to extend the flaps Captain.




WING — CLEAN

SPEED — 190 KTS

RECOMMENDE

WING — SLATS

SPEED — 150

(APPROACH CHECKL

START TIMING..........................


SPEED — 150 KTS

AT THE FINAL FIX



(AT MINIMUMS)............

SPEED — 150 KTS

15 SEC

90°


MISSED APPROACH—ONE ENGINE INOPE

When decision height is reached on an ILS approach, omissed approach point as specified for a nonprecision acontact with the airport or landing runway cannot be mad

approach (go-around). Figure MAP-9 illustrates a Misengine inoperative.

Follow the missed approach instructions as depicted instrument approach plate used for the approach.

Simultaneously apply maximum thrust to the operating eaircraft to a 13° pitch attitude. The flaps are assumed seton the one-engine inoperative approach profile. When a pis indicated on both the altimeter and the vertical velocitthe landing gear. Maintain the charted VREF speed throuuntil a minimum altitude of 400 feet above airport elevthis altitude, accelerate to VREF +25 knots, and retract the slto 1.43 VS, clean configuration, reduce the engine powetinuous, and climb the aircraft to the prescribed missed-



MAMAXIMUM THRUST(5 MINUTES MAX)

SL ATS— + 2 0 PITCH—13° /FLAPS—+20°POSITIVE RATE—GEAR UP

SPEED—VREF

+5 KT

SPEED—VREF

+25 KT

FLAPS + SLATS—AFTER

TAKEOFF CHECKLIST

SPEED—1+ 10 KT

–0


WEIGHT AND BAL

CONTENTS

INTRODUCTION................................................

WEIGHT AND BALANCE.................................

Definitions....................................................




WEIGHT AND BAL

INTRODUCTIONIt is the responsibility of the airplane operator tois properly loaded. At the time of delivery, the mnecessary weight and balance data to compute indsequent changes in airplane weight and balance arairplane owner and/or operator. Information in weight and balance definitions.

The second portion of this chapter covers perfordefinitions and supplies acceptable performance g

WEIGHT AND BALANCE

DEFINITIONS

Maximum Gross Weight—The maximum grossplane has been certified in compliance with the Fe



Payload—Weight of passengers, baggage, and carand usable fuel)

Center of Gravity (CG)—The point at which the

sidered to be concentrated

Arm (or Moment Arm)—The horizontal distanaxis from the datum to the point where a force is apin inches, aft of the datum is plus (+), and forward

Moment—The product of a weight or force and its


Example

Although seemingly complex at times, all weight and bahandled by use of the following moment equation.

(1) Moment = Weight X Arm

This equation is the basic equation used to find the cention of an airplane and/or its components. By rearrangemto the following forms,

(2) Weight = Moment, and (3) Arm = Moment

Arm Weight

with any two known values, the third value can be found

In the airplane weight and balance problem, the moment eqtimes in calculating moments for each individual item. Wmoments have been totaled, the charts provided in the Pof this chapter should be used to determine CG.




PERFORMANC

DEFINITIONSTakeoff Path—Extends from a standing start to a takeoff surface

Takeoff Distance—The horizon distance along start to a point 35 feet above the takeoff surface foor 115% of all engines operating to a point 35 feet(this includes a legal clearway)

Accelerate Stop—The distance required to accelercraft to a full stop, assuming that one engine failedseconds (this includes a legal stopway)

Balanced Field—When the takeoff distance is eqdistance

V1—The speed at which, if an engine failure occu

• Reach 35 feet above the takeoff surface, or



• Come to a full stop on the takeoff surface

VR—The speed at which rotation is initiated; attain35 feet

V2—The takeoff safety speed selected by the manufclimb gradient is attained

VFR—The minimum speed to initiate flap retracti

VMIN—The minimum speed at which the takeof


VMIN — Replaces VMCG

Landing Distance—The horizontal runway surface nethreshold from 50 feet at VREF, maintaining a steady 3°surface, and come to a full stop using brakes, and airbra

Landing Field Length—Landing distance multiplied b

ACCEPTABLE PERFORMANCE GUISteep turns—±100 feet, ±5° bank, ±10 knots, ±10° hea

Approach to stall—Recognize perceptible stall/warninrecover at first indication, striving floss.

Holding—±100 feet, ±10 knots

IFR approaches—Initial: ±100 feet, ±10 knots

Final: –0 +10 knotsDH/MDA: –0 +50 feet, +5 knots ± 1

Circling—Not to exceed 30° bank MDA –0 feet +100 f



Circling Not to exceed 30 bank, MDA 0 feet +100 f

Missed approach—DH/MDA: –0 feet (except in instaenvironment is in sight)

Engine failure—V1: V2 KIAS, runway heading, –0 +10

Clean climb: VFS KIAS, –0 +10 knots

In flight: Shutdown/restart ±20° headiKIAS on driftdown

Landings—Traffic pattern: ±10 knots, altitude ±100 fee

CREW RESOURCE MAN

CONTENTS

CREW CONCEPT BRIEFING GUIDE ..............

Introduction ....................................................

Common Terms ..............................................

Pretakeoff Briefing (IFR/VFR) ......................

Crew Coordination Approach Sequence ........

ALTITUDE CALLOUTS ....................................

Enroute ..........................................................

Approach—Precision ....................................

Approach—Nonprecision ..............................

Significant Deviation Callouts........................




ILLUSTRATIONS

Figure Title

CRM-1 Situational Awareness in the Cockpi

CRM-2 Command and Leadership ..............

CRM-3 Communication Process ................

CRM-4 Decision-Making Process ..............




CREW RESOURCE MAN(CRM)

CREW CONCEPT BRIEFING GUI

INTRODUCTION

Experience has shown that adherence to SOPs helpscrew cockpit situational awareness and will allow ato be attained. Our objective is for standards to be a

and then adhered to, such that maximum crew perfoprocedures are not intended to supercede any indirather are examples of good operating practices.

COMMON TERMS

PIC Pilot in Command

Designated by the company for flights reqResponsible for conduct and safety of thflying and pilot not flying duties.




F Pilot Flying

Controls the aircraft with respect to assi

tude, airspeed, etc., during normal anAccomplishes other tasks as directed by

N Pilot Not Flying

Maintains ATC communications, copies checklists and other tasks as directed by

F


CAPTAININDIVIDUAL

S /A

COPILOTINDIVIDUAL

S /A

IT's U

CLUES TO IDENTIFYING:• Loss of Situational Awareness• Links in the Error Chain

M A N

O P E

R A T I O N A L 1. FAILURE TO MEET TARGETS

2. UNDOCUMENTED PROCEDURE 3. DEPARTURE FROM SOP 4. VIOLATING MINIMUMS OR LIMITATIONS

5. NO ONE "FLYING AIRPLANE" 6. NO ONE "LOOKING OUT WINDOW"

7. COMMUNICATIONS

8. AMBIGUITY

9 UNRESOLVED DISCREPANCIES

Re

2

2(S

GROUPS /A



H U M 9. UNRESOLVED DISCREPANCIES

10. PREOCCUPATION OR DISTRACTION

11. CONFUSION OR EMPTY FEELING

12.

CRM-1. Situational Awareness in the C

LEADERSHIP STYLES

PRETAKEOFF BRIEFING (IFR/VFR)

NOTE

The following briefing is to be com

item 1 of the pretakeoff checklist. The paccomplish the briefing.

1. Review the departure procedure (route ansignificant terrain features, etc.)

2. Review anything out of the ordinary

3. Review required callouts, unless standard cin which case a request for "Standard Ca

4. Review the procedures to be used in case of

5. As a final item, ask if there are any quest

6. State that the pretakeoff briefing is comp

CREW COORDINATION APPROACH S

NOTE




The following crew coordination approshould be completed as early as possibltiating an IFR approach. These items are

during the “APPROACH (IN RANGE)

F—Requests the pilot not flying to obtain destinacommunication duties to the pilot flying may ment of this task.

N—Advises the pilot of current destination weathercial information pertinent to the destination

F


THINK:• Solicit feedba• Listen• Focus not pe• Mainta the go• Verify

outcom achiev

ADVOCACY: to increase others' S/A

• State Position

• Suggest Solutions• Be Persistent and Focused

• Listen Carefully

— REMEMBER—Questions enhance communication flow.

INTERNALBARRIERS

EXTERNALBARRIERS

INTERNALBARRIERS

OPERATIONAL

GOAL

RECEIVENEED SEND

FEEDBACK

INQUIRY: to increas

• Decide What, Who

• Ask Clear, Concis• Draw Conclusions

Valid Information• Keep an Open Min



Don't give in to the temptation to ask questions when Advoca

Use of Advocacy or inquiry should raise a "red fla

HINTS:

• Identify the prob

CRM-3. Communication Process

F—Requests the pilot not flying to perform the ap

N—Accomplishes the approach setup and advisestified and course set

F—Transfers control of the aircraft to the pilot not fcontrol, heading ____________ , altitude _____structions. (Communications duties should bpilot not flying at this point.)

N—Responds, “I have control, heading ___________

F—Advises, "Approach briefing."

F—At the completion of the approach briefing,“Approach briefing complete.”

F—Advises, “I have control, heading ___________

N—Confirms “You have control, heading ________

F—“Before Landing checklist.”




N—“Before Landing checklist complete.”

NOTE

The above sequence should be compthe FAF.

NOTE

During the above sequence, the terms

F


ALTITUDE CALLOUTS

ENROUTE

1,000 Feet Prior to Level-Off

N

State altitude leaving and assigned “ROGlevel-off altitude

“100 above/below” “LEV

APPROACH—PRECISION

N

At 1,000 feet above minimums

“1,000 feet above minimums” “DH _

At 500 feet above minimums

“500 feet above minimums” “NO F



500 feet above minimums NO F

At 100 feet above minimums

“100 feet above minimums” “APPR

MIN

At decision height (DH)



F


SIGNIFICANT DEVIATION CALLOUTS

N

IAS ± 10 KIAS

“VREF ± ______” “CORREC

Heading ±10° enroute, 5° on approa

“Heading ________ degrees left/right “CORREC

Altitude ±100 feet enroute, +50/–0 feet on fina

“Altitude _________ high/low” “CORREC

CDI left or right one dot

“Left/right of course ________ dot” “CORREC

RMI course left or right ±5°



“Left/right of course ________ degrees” “CORREC

Vertical descent speed greater than 1,000 fpm on final

“Sink rate _________” “CORREC

Bank in excess of 30°

“Bank ________ degrees” “CORREC


ILLUSTRATIONS

Figure Title

SR-1 DC Distribution Buses ....................

SR-2 Normal Flight Configuration..........SR-3 Battery Bus Circuits—Typical.........

SR-4 Reservoir Fluid Level Indication .....

SR-5 No. 1 and No. 2 Hydraulic Systems

SR-6 Standby Pump Pressurizing No. 2 S

SR-7 Standby Pump Pressurizing No. 1 S

SR-8 Landing Gear Retraction .................

SR-9 Landing Gear Extension .................

SR-10 Gear Emergency Hydraulic Extensio

SR-11 Position and Warning Indications—Hydraulic and Gravity Extensions...

SR-12 Displays and Warnings duringLanding Gear Extension .................



SR-13 Normal Braking with Antiskid—Airwithout SB F-900-42 .......................

SR-14 Normal Braking with Antiskid—Aircraft with SB F-900-42..............

SR-15 Emergency Brake Operation ..........

SR-16 Emergency/Parking Brakes .............

SR-17 Leading Edge Slats .........................


SR-25 Fuel Distribution ......................................

SR-26 Crossfeed X-BP1 3—Pump 1 Inope

SR-27 Crossfeed X-BP1 3—Pump 3 Inope

SR-28 Crossfeed X-BP1 2 and 32—Normal Configuration ........................

SR-29 Crossfeed X-BP1 2 and 32—Pump 1 or 3 Inoperative ....................

SR-30 Crossfeed X-BP1 2 and 3 2—Pumps 2 Inoperative ................................

SR-31 Refueling System Controls and Indicators

SR-32 Pressure Refueling....................................

SR-33 Gravity Distribution..................................

SR-34 Bleed-Air Operation—Anti-icing Off ......

SR-35 Bleed-Air Operation—Anti-icing On ......SR-36 Distribution System Normal Operation—

Flight (Heating) ........................................

SR-37 Distribution System Normal Operation—



SR 37 Distribution System Normal OperationFlight (Cooling) ........................................

SR-38 Bleed-Air System—Power Lever to TakeoSR-39 No. 3 Nacelle Anti-icing Operation..........

SR-40 No. 1 Nacelle and Ram-Air InletAnti-icing Operation ................................

SR-41 No. 2 Nacelle and S-Duct Anti-icing Oper


TABLES

Table Title

SR-1 Electrical Power Sources ................

SR-2 Start-Assist Logic ..........................SR-3 Sequential Start-Assist Logic ........

SR-4 No. 1 and No. 3 Engine Nacelle AntLogic System ...................................




LEFT CIRCUIT-BREAKER PANE

PRIMARY BUS A1

DESIGNATION UTILIZATION

FIRE AUDIO WARN A Audio warning RADIO WARNING

WARN LIGHTS A Warning panel PANEL

WARN LIGHTS A Light test EX D/N

EXTING 1 Fire NAVIGA- TION

DETECT 1 Fire

LIGHTS WARN A-B

Table SR-1. ELECTRICAL POWER



NAVIGA- IRS

1 BAT IRS

1 battery TION

IRS 1 IRS 1

TEMP PROBE Probe heating

HRZN ST BY Standby


LEFT CIRCUIT-BREAKER PANEL (Cont)

PRIMARY BUS A2

DESDESIGNATION UTILIZATION

FIRE EXTING 3 Fire RADIO HF 1 WARNING CON

DETECT 3 Fire PHON

BAG COMP Fire

BAT TEMP Battery NAVIGA- AFCS

temperature TION CM

AFCS

ADV

BLOWER LH Ventilation RAD A

TEST WARN A-B FMS

CDU

NAVIGA- IRS 3 IRS 3 SG 3*TION

SAT C

Table SR-1. ELECTRICAL POWER SOURC



TION

IRS 3 BAT (Option) IRS 3 battery MFD/

ANNUNC LH Radio nav R/T W lighting

AFCS 1 AP Pilot FGC GPWS

AFCS 1 YD Pilot FGC

S


CENTER CIRCUIT-BREAKER PA

PRIMARY BUS A1


LIGHTS ANTICOL FIN Anticollision HYDR MISC lights

LH EXT External LIGHT lighting

CKPT LH Lighting READING

NAV Navigation lights

STROBE Strobe ANTI- lights ICE COND’G

CENTER Lighting

INSTR LH Instrument lighting

INV (115V/60 Hz) Option

ENGINES IGNTR AUTO Starting

or 115-VAC master

Table SR-1. ELECTRICAL POWER SO



ENG FAIL 2 Takeoff warning

N2 1 Indicators N1 ITT 2 Turbine temp

CMPTR Engine Computer

IGNTR 1 Start


CENTER CIRCUIT-BREAKER PANEL (Cont

PRIMARY BUS A2

DESDESIGNATION UTILIZATION

LIGHTS BELTS NO Passenger ANTI- ENGIMISC SMK’G signs ICE COND’G DV W

ENTRY Entrance lighting

SHIELD Glareshield PRV 3 lighting

DRAIN HEAT Drain CAB T anti-icing CON

LANDING LH Lights STBY

WIPE

ENGINES N2 3 Indicators

CMPTR 3 Engine FLT FLAP computer CONTROL IND




IGNTR 3 Starting TRIM TRIM

OIL 3 Engine STICK control

FUEL 2 Fire SHUT OFF

FUEL STBY BOOST 2 Fuel




CENTER CIRCUIT-BREAKER PANEL (Cont)

PRIMARY BUS B2

DESIGDESIGNATION UTILIZATION

LIGHTS GALLEY MASTER Galley FUEL BOOSTMISC

LANDING RH Lights GAGES

ANTICOL BELLY Anticollision XBP1-2 lights

RH EXT LIGHT Right external PRESS

lights FUELIN

INSTR RH Instrument lighting

PEDESTAL Instrument HYDR NOSE lighting

AFT CABIN Cabin INDIRECT lighting

LH CABIN Reading ANTI- CKPT T READING lights ICE CONT COND’G BAG P

VALANCE OR




ENGINES APU APU WSHLD RH

N1 ITT 3 Turbine RH PIT temperature

FUEL 1 Fire RH STA SHUT OFF HEAT

REVERSE Engine 2CONTROL reverser


RIGHT CIRCUIT-BREAKER PANE

PRIMARY BUS B1


FIRE AUDIO WARN B Audio warning RADIO WARNING

WARN LIGHT B Warning panel PANEL

WARN LIGHTS B Light testing EX D/N

EXTING 2 Fire

DETECT 2 Fire NAVIGA- TION

APU Fire

Table SR-1. ELECTRICAL POWER SO



APU Fire

NAVIGA- IRS 2 BAT IRS 2 battery TION


RIGHT CIRCUIT-BREAKER PANEL (Cont)

PRIMARY BUS B2

DESIDESIGNATION UTILIZATION

MISC NOSE FAN Ventilation RADIO HF 2 CON

BLOWER RH Ventilation VHF 2*

CREW SEATS Crew seats SELCA

EMERG LIGHTS Emergency light batteries

NAVIGA- AFCS 2 TION CMP

RADIO HF 2 PWR HF 2 AFCS 2 ADVI

PUBLIC Public FMS 2* ADDRESS address




ICS RH Intercom CDU 2*

OMEG

NAVIGA- AFCS 2 AP Copilot FGC FLIGHTTION

AFCS 2 YD Copilot FGC

VOICE R

RAD ALT


ONL Y

X XX

X X X XX X X X

X X X X X

X X XX X X XX X X X

XX X X XX

OFF

OFF

S T A R T G E N

ON

OFF

ON

OFF

ON

OFF

ON

BATLIGHTS

STARTASSIST

MAINBUS

POWER

ON ON ON ON ON ON ON OFF ON OOFF

G3 A

G3G3G3 AG1G2

G1

G3G2

G3

B1

G1

G2

G1

G2

B2

B1

B2

APU

APU

2

3

1

Table SR



S R- 9

S R-

1 0

F

OR T RAI NI N G P URP O SE S

ONL Y

X X

X X X XX X X X

X X X X

X XX X X XX X X X

XX X X XX

OFF

OFF

S T A R T G E N

ONOFF

ON

OFF

ON

OFF

ON

BATLIGHTS

STARTASSIST

MAINBUSPOWER

ON ON ON ON ON ON ON OFF ONOFF

G3

G3G3G3 G1G2

G1

G3G2

G3

B1

G1

G2

G1

G2

B2

B1

B2

APU

APU

2

3

1

Table SR-3. SE




ONL Y

BUS A1 BUS A2

LEFT CIRCUBREAKER

PANEL

(1

TOPO

CONTAC

LEFT MAIN BUS

MAROTA

1 3 0 A

1 3 0 A

1 5 0 A

1 5 0 A

1 5 0 A

8 0 A L

B U S

T I E D

S T A NDB Y HY DRA UL I C

P UMP A 5

P I L OT F R O

NT

WI ND S HI E L DA 3

GA L L E Y 2 B A R

A 6

GA L L E Y 1 B A R

A 4

OFF

L AVMASTER

0

2025

30V –

BAT

GEN



S R- 1 1

PANELAVIONICMASTER

Figure S




ONL Y

FUABO(SS

SINGLE-POINTREFUELING

BAT 2

BAT 1

1. BAT 1 MAKE-AND-BREAK 2. GEN 1 LINE CONTACTOR 3. GEN 3 LINE CONTACTOR 4. BUS-TIE RELAY 5. GEN 2 LINE CONTACTOR 6. APU LINE CONTACTOR 7. BAT 2 MAKE-AND-BREAK 8. BAT 1 LINE CONTACTOR 9. ENG 1 START RELAY10. ENG 3 START RELAY11. APU LINE CONTACTOR12. ENG 2 START RELAY13. APU START RELAY14. BAT 2 TO START BUS CONTACTOR

B

A

T

T

E

R

Y

B

U

S

MAIN DC BOX COMPONENTS



S R- 1 3

15. CURRENT LIMITERS

16. BAT 1 CONTROL CB

17. BAT 2 CONTROL CB * GEN 1 AND GEN 3 LINE SWITCHES

** GEN 2 AND APU GEN LINE SWITCHES

Figure SR-3


VOLUME(LITERS)

LEVELINDICATOR

NO. 1 RESERVOIR NO. 2 RES

90°C (194°F)

90°C (194°F)

9090°C (194°F)20°C (68°F)

20°C (68°F)

2020°C (68°F)–40°C (–40°F)

–40°C (-40°F)

–40

–40°C (–40°F)

9

8

7

6

5

4

3

2

1

0

U N P R E S S U R I Z E D

P R E S S U R I Z E D

U N P R E S S U R I Z E D





O

NL Y

P1

LEGENDNO. 1 SYSTEM PRESSURE

NO. 2 SYSTEM PRESSURERETURN

SUPPLY FLUID

NITROGEN

ELECTRICAL

EXTERNAL HYDRAULIC POWER

QUICK DISCONNECT

FILTER WITH SELF-SEALING VALVE

PRESSURESWITCH

PRS

NO. 1 SYSTEMACCUMULATOR

SYSTEMPRESSURE

TRANSMITTER

ACCUMULATOR

PRESSURE-HOLDINVALVE

NO. 1 RESERVOIR



S R- 1 5

FILTER WITH SELF SEALING VALVEAND CLOGGING INDICATOR

CHECK VALVE

RELIEF VALVE

Figure SR-5. No. 1

S R- 1 6

F

OR T RAI NI N G P URP O SE S

ONL Y

P1

LEGENDAUXILIARY PRESSURE

RETURN

SUPPLY FLUID

NITROGEN

ELECTRICAL


QUICK DISCONNECT

FILTER WITH SELF-SEALING VALVEAND CLOGGING INDICATOR

CHECK VALVE

RELIEF VALVE



Figure SR-6. Stan

F OR T RAI NI

N G P URP O SE S

O

NL Y

P1

LEGENDAUXILIARY PRESSURE

RETURN

SUPPLY FLUID

NITROGEN

ELECTRICAL


QUICK DISCONNECT




S R- 1 7

AND CLOGGING INDICATOR

CHECK VALVE

RELIEF VALVE

Figure SR-7. Stan

S R- 1 8

F

OR T RAI NI N G P U

RP O SE S

ONL Y

NORMAL/EMERGENCYGEAR SELECTOR

VALVE

LANDING GEARELECTRIC

SELECTOR VALVE

RETRACTION

AUTOMATICWHEEL

BRAKING

GEARUPLOCK

BOX GEARUPLOCK

UNIT

DOORUPLOCK

BOX

MAIN GEARBRACING

CYLINDER

LEFT MAIN

RETRACTION

RETRACTION

NOSE GEARACTUATOR

RETRACTION

OPENING

DOORACTUATOR

EMERGENCY NORMAL



Figure SR


VALVE


SELECTOR VALVE

RETRACTION

GEARUPLOCK

BOX GEARUPLOCK

UNIT

DOOR

MAIN GEARBRACING

CYLINDER

EXTENSION

EXTENS

ION

NOSE GEARACTUATOR

EXTENSION

OPENING

DOORACTUATOR



UPLOCKBOX

LEFT MAIN

Figure SR

S R- 2

0

F

OR T RAI NI N G P U

RP O SE S

ONL Y


VALVE


SELECTOR VALVE

GEARUPLOCK

BOX GEARUPLOCK

UNIT

DOORUPLOCK

BOX

MAIN GEARBRACING

CYLINDER

LEFT MAIN

EXTENSION

EXTENSION

NOSE GEARACTUATOR

EXTENSION

OPENING

DOORACTUATOR

EMERGENCY NORMAL



Figure SR-10. G

F OR T RAI NI

N G P URP O SE S

O

NL Y

LANDING GEARNOSE MAIN

NOSE GEAREMERGENCY

CONTROL ME

*WHEN THE SLAT/FLAP CONTROL IS SET TO 40°,THE WARNING VOICE CANNOT BE SILENCED

WITH ALL GEAR NOT LOCKED DOWN.



S R- 2 1

Figure SR-11. Position and Warning I

S R- 2

2

F

OR T RAI NI N G P U

RP O SE S

ONL Y

NORMAL GEAREXTENSIONCONTROL

BEFORE GEAR

EXTENS

ION

*WHEN THE SLAT/FLAP CONTROLTHE WARNING VOICE CANNOT B

LANDING GEAR

NOSE MAIN

HORNSILENC

*

*

*

Fi SR 12 Di l



Figure SR-12. Displays a

F OR T RAI NI

N G P URP O SE S

O

NL Y

PUMP 1TEST

ASKID

1 / 4

HYDR

#1 ON

PUMP 3

L R

HYDRAULIC PANEL

ANTISKIDSERVOVALVES

EMERGENCYCHAMBER

NORMALCHAMBER

NBVA

COPPILOT PEDALS

FLEXIBLE CABLES

AUTOMATICBRAKING

NORMAL CONTROLHANDLE

GEAREXTENDED

BRAKES

EWHEELHOMETERS

NORMAL SELECTOR VALVE(DEENERGIZED OPEN)

WHEELMETER

ERATOR



S R- 2 3 Figure SR-13. Normal Brak

S R- 2

4

F

OR T RAI NI N G P U

RP O SE S

ONL Y

DOUB

BRAKIPC BOA

NO. 1 OR

NO. 3 ENGINEFULL POWER

ST BYPUMP

ANTISKID

GEARNORMAL CONTROL

(2GA)

GEAR DOWN BRAKESOL

#1 ON

#2 OFF#1 OFF

+28 VDCBUS A1

+28 VDCBUS A2

PBR

DISTRIBU

PARK BRAKE HANDL

LEGENDNORMAL PRESSURE

EMERGENCY PRESSURERETURN

NO. 1 SYSTEM SUPPLY

NO. 2 SYSTEM SUPPLY

ELECTRICAL

GROUND

Figure SR-14. Normal Bra



Figure SR 14. Normal Bra


O

NL Y

BRAKE

ASKID

1 / 4

1 / 2

0

#1

#2

#1

ON

OFF

OFF

ANTISKIDELECTRONIC

CONTROLBOX

HY

NORMALCHAMBER

PILOT PEDALS

FLEXI

AUTOMATIC

BRAKING


GEAREXTENDED

HYDRAULICPANEL

+

+

#2 P BK

NO. 1SYSTEM

NORMAL SELECTOR VALVE(ENERGIZED CLOSED)



S R- 2 5

HY

Figure SR-1

S R- 2

6

F

OR T RAI NI N G P U

RP O SE S

ONL Y

Figure SR-1

T

#1

BRAKE

ASKID

1 / 4

1 / 2

0

#1

#2

#1

ON

OFF

OFF

ANTIS

KIDELECTRONICCONTROL

BOX

HYD

SHVA

PILOT PEDALS

FLEXIBL

NORMAL CONTROL

HANDLE

GEAREXTENDED

HYDRAULIC

PANEL

+

+

NO. 1SYSTEM

#2 P BK



Figure SR 1


SLAT/FLA

S

POSITINDICAT

CENTERCIRCUIT-BREAKER PANEL

WARN

XTK 2OPEN

AILZERO

AP

NOSECONE OVHT

REARDOORS

L PITOT

HOTBAT



21CF10

SLATINDIC

LH AUTOSLAT

CLEA

7° FLAPS + SLA

20° FLAPS + SLAT

40° FLAPS + SLA

5

S R- 2

8

F

OR T RAI NI N G P U

RP O SE S

ONL Y

SLATS /FLAPCONTROL BOX

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

LEGENDNO. 1 PRESS

ADC 1IAS LESS THAN

265 KT

AOALESS

THAN 23°LEFT ANGLE-OF-ATTACK

VANE

INBOARDVALVEBOX

INBOS

LSOLESELE

VA

EXTEND

LEFT OUTBOARD

R E R E



Figure SR


CONTROL

1

AIRPLANEANGLE OF

ATTACKAOA11°

AIRPLANE AOA

SLATPOSITI

CLEAN

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

CLEAN

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

A

CROSS-SOUTBOA

CROSS-SECTION INBOARD SLAT

A

CROSS-SOUTBOA

CROSS-SECTION INBOARD SLAT

INCREASING

AOA = 11°

IGN

DECREASING

AOA = 11°

IGN

1

2

2



INITIALCONFIGURATION (INCREASING)

AIRPLANE IN CLEANCONFIGURATION,SLAT/FLAP HANDLEIN CLEAN

• EXTENSION OF OUTBOARD SLATS

• RED SLAT-MOVING LIGHT ON

• AUTOMATIC IGNITION, IGNLIGHT ON

• R

• G R

• A

• A




CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

A

CROSS-SECTION AINBOARD SLAT

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

A

CROS

OUTB


CONTROL7° FLAPS +SLATS

AIRPLANEAOA

SLATPOSITI

DECREASING

AOA = 23°

IGN

DECREASING

AOA = 16°5

4

5



AOA<16°5

INITIALCONFIGUR-

AIRPLANE ANGLE-OF-ATT

CROSOUTB

AOA≥16°5

AOA≥23°

1 2 3

S R- 3

2

F OR T RAI NI N G P U

RP O SE S

ONL Y

OUT OFCLEAN

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

LEGEND

NO. 1 PRESSURE

ADC 1IAS LESS THAN

265 KT

AOA

LESS


VANE

INBOARDVALVEBOX

INBOASLAT

SOLENSELECT

VALV

EXTEND

LEFT OUTBOARD

R E R E

AOA 11°



Figure SR-21. Automatic E

F OR T RAI NI N

G P URP O SE S

O

NL Y

OUT OFCLEAN

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

LEGEND

ADC 1IAS LESS THAN

265 KT

AOAABOVE

23°LEFT ANGLE-OF-ATTACK

VANE

INBOARDVALVEBOX

INBOARDSLAT

SOLENOIDSELECTOR

VALVE

LEFT OUTBOARD LEINBO

R E R E R

AOA 11°



S R- 3 3

NO. 1 PRESSURE

EXTEND

Figure SR-22. Automatic Retrac

S R- 3

4


RP O SE S

ONL Y

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS LEGEND

NO. 2 PRESSURE

ADC 1IAS LESS THAN

265 KT

AOALESS


VANE

INBOARDVALVEBOX

INBOARDSLATSOLENOIDSELECTOR

VALVE

EXTEND

EMERGENCYACTUATOR

LEFT OUTBOARDLE

INBO

R E R E R



Figure SR-23. Em


FUEL 1

LOFUEL 1

BLEEDOVHT

ECUOVHT

COND GOVHT

LOFUEL 2

REVUNLOCK

LOFUEL 3

FUEL 2

-2P BK CABIN

FUEL 3

FUELING

LEVEL LESS THAN 200 LB

CENTER WINGTANK

PROBE

SUMP DRAIN V

FRONT TAN

* NOT FEATURED ON AIRCRAFT

ELECTRIC TRANSFER VALVE X



G1

G2

≥ 159

G1

NO. 1 ENGINE

FILLERPORT

CENTERWING TANKS

NEGATIVEPRESSURERELIEF

VALVE

PROBE

INT

(NO

R1 CROSSFEEDVALVE

GROUP 1 INTER-CONNECTION

MANIFOLD

TRANSFER VALVE(ON SOME AIRCRAFT)

GROUP 1CROSSFEED

MANIFOLD

GROUP 2FUEL

MANIFOLD

FUELSHUTOFF

VALVE

Fi

S R- 3

6


RP O SE S

ONL Y



Figure


XTK

X.BP

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

BP1

R1

G1

G1

E

AUTOOPENCLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C

POSITION OFTHE CONTROLS

TANKINTERCONNEC

NOINTERCONN

LIGHTS

XTKX-BPBP1

BP3

NEUTRALOPENOFF

ON

XTKX-BP

OUTON

XTK

X.BP

BP1

LEVEL LEVEL

BOOSTER BOOSTER

BOOST PUMP 1 FAILUREFLOW

*



XTK 2

X.BP X.BPLEVEL

BP1

R1

G1

AUTOOPENCLOSED

BOOSTER

NORM

ST-BY

OFF

°C

*


XTK

X.BP

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

BP1

X

R1

G1

G1 G3

ENGINE 1

AUTOOPENCLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C


TANKINTERCONNECTION

NOINTERCONNECTION

LIGHTS

XTKX-BPBP1BP3

NEUTRALOPENONOFF

XTKX-BP

OUTON

XTK

X.BP

BP1

LEVEL LEVEL

BOOSTER BOOSTER

°C


*



XTK 2

X.BP X.BPLEVEL

R1

G1

AUTOOPENCLOSED

BOOSTER

NORM

ST-BY

OFF

*


XTK

X.BP

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

AUTOOPENCLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C

LEGEND

GROUP 1 B

GROUP 2 B

ENG

NOT FEATURED ON AIRCRAFT WITHOUTELECTRIC TRANSFER VALVE XTK 2*

*



G

GROUP 3 B


TANKINTERCONNECTION

NOINTERCONNEC

LIGHTS

BP1BP2

ONNORM

ELECTRIC TRANSFER VALVE XTK 2




XTK

X.BP

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

AUTOOPENCLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C

XTK

X.BPLEVEL LEVEL

BOOSTER BOOSTER

BOOST PUMPS 2 FAILURE—ENGINE 2 FE

LEGEND

GROUP 1 BOOST

GROUP 3 BOOST

E


LIGHTS

BP1BP2BP3

X-BP 1–2X-BP 2–3

ONOFFON

X-BPX-BP

ONOFF

OPENCLOSED

*

NOT FEATURED ON AIRCRAFT WITHOUT

ELECTRIC TRANSFER VALVE XTK 2*



XTK 2

X.BP X.BPLEVEL

AUTOOPENCLOSED

BOOSTER

NORM

ST-BY

OFF

°C

*

S R- 4 2

F O

R T RAI NI N G P U

RP O SE S

ONL Y

20

15

10

5

0

FUEL QTY

LBS X 1000

20

1510

5

0

FUEL QTY

LBSX 100

GROUNDIRECEPTAC

LEFT REFUELINGSWITCH

FUELINGOK LIGHT

RAUTO

TEST

REFUELING MODESELECTOR

(OPTIONAL)

STOPFUELING LIGHT

GROUP 1 INDICATION

PARTIAL REFUELINGSELECTOR

PRESSUR

FULL

LEFT

ON

FULL STOPFUELIN

FUELINOK

PARTIAL

OFF

PRESSURE FUELING PANEL

FULL FULLFULL

LEFT CENTER RIGHT

ON

FULL STOPFUELING

CLOSED

ON

OFF

VENTVALVE TEST

OPEN

FUELINGON

PARTIAL

OFF

ON

OFF

ON

OFF

Figure SR-31. Ref



g

F OR T RAI NI N

G P URP O SE S

ONL Y

S R


FULL FULLFULL

LEFT CENTER RIGHT

ON

FULL STOPFUELING

CLOSEDTEST

DEFUELING

ON

OFFVENT

VALVE TEST

OPEN

FUELINGOK

PARTIAL

OFF

ON

OFF

ON

OFF

G1

G1 END-OF-REFUELINGTHERMISTOR

PRESSURIZATIONINTERCONNECTION

CENTER WINGTANK

PRESSURE REFUELING

ELECTRICAL

LEGEND



R- 4 3

Figure S

S R- 4 4

F O

R T RAI NI N G P URP O SE S

ONL Y

G1

BOOSTERCROSSFEED

MANIFOLDCROSSFEEDVALVE

GRAVITYFUELING

VENTVALVE

GROUP 1 BOOST

LEGEND

GROUP 3 BOOST

BAG 1 COMP 1

BLEED AIR

FUEL FLOW

TEST

ENG INDIC

TEST

GRAVITY

FUELING

STABILIZER

RELAY TEST

COOLING

FANANTI-SKID

ISOL HEAT

ON

OFF

RESET

MAINTENANCE PANEL

F U E L

J E T A, J E T A1, J E T

B , J P 4 ,

J P 5

C A U T

I O N : PRESSU R I Z E D T A N

K

F O R F U E L I N

G I N S T R U C T I O NS SE E O P

E R G M A

N U A L

T O T A L A I R C R A F

T U S AB LE

QUA N T I T Y 2 8

4 5 U S

G A L

* NOT FEATUR ELECTRIC T

ELECTRICAL

Figure S



F OR T RAI NI N

G P URP O SE S

ONL Y

S R-

NO. 1ENGINE

MAINBLEE

WING ANTI-ICING SYSTEM

CREW AIR-CONDITIONINGVALVE

FUEL TANK PRESS

URIZATIONS

YS

TEM

PRESSURIZATION JET PUMP

PASSENGER CABINAIR-CONDITIONING VALVE

MAIN HPBLEED PORT

MAIN LPBLEED

PORTAIR INTAKEANTI-ICING

AIR INTAKEANTI-ICING

PRV3

AUXILBLEE

MAIN LPBLEED PORT

NO. 3ENGINE

LP3BLEED-AIR

TEMPERATURE PROBE

AUXILIARY HPBLEED PORTAUXILIARY LP

BLEED PORT

LP1 BLEED-AIRTEMPERATURE

PROBE

Fi SR 34



-4 5 Figure SR-34.

S R- 4 6

F O


ONL Y

NO. 1ENGINE

MAIN HPBLEED PORT


CREW AIR-CONDITIONINGSYSTEM

FUEL TANK PRESSURIZATION SYSTEM


PASSENGERAIR-CONDITIONING SYSTEM

MAIN HPBLEED PO

MAIN LPBLEEDPORT



PRV3

AUBL

MAIN LP

BLEED PORT

NO. 3ENGINE

LP3BLEED-AIR

TEMPERATURE PROBE

AUXILIARY HBLEED POR

AUXILIARY LPBLEED PORT


PROBE

BLEEDOVHT

Figure SR-35.



F OR T RAI NI N

G P URP O SE S

ONL Y

S R-

COPILOTGASPER

FLAPPERVALVE

PASSENGERDOOR

HEATING

TWO-WAYMANIFOLDS

PASSENGERGASPERS

EFISCOLD AIR

PILOTGASPER

CREWTEMPERATURESENSOR

FLAPPERVALVE

TWO-WAYDUCT

F

PILOTFOOTWARMERS

NOSE CONEISOLATION

VALVE(PRIOR TO SNs 69)

EFISBLOWERS

NOSECONE

BLOWER

DEFOGOUTLETS

EFIS VENTILATIONAIR OUTLETS

MANUALDISTRIBUTOR

CONTROL

AIR SOURC1. COOL2. BLEED

LEGENDHOT AIR

COOLED AIR

CONDITIONED AIR (HEATED)

RECIRCULATING AIR

AFT WINDOWVENTILATION DUCT(PRIOR TO SN 116)

GASPER PICK-OFF

THIRD CREW MEMBER

INTERCONNE

FLAP(SNs PRIOR

Figure SR-36 Distributio



4 7 Figure SR-36. Distributio

S R- 4 8

F O


ONL Y

COPILOTGASPER

FLAPPERVALVE

PASSENGERDOOR

HEATING

TWO-WAYMANIFOLDS

PASSENGERGASPERS

EFISCOLD AIR

PILOTGASPER

CREWTEMPERATURE

SENSOR

FLAPPERVALVE

TWO-WAYDUCT

PILOT

FOOTWARMERS

EFISBLOWERS

NOSECONE

BLOWER

DEFOG

OUTLETS


MANUALDISTRIBUTOR

CONTROL

AIR SOURCES:1. COOLED AIR F2. BLEED AIR TOCONDITIONED AIR (COOLED)

LEGENDHOT AIR

COLD AIR

AFT WINDOWVENTILATION DUCT

(PRIOR TO SN 116)

GASPER PICK-OFFTHIRD CREW MEMBER

INTECONN

FLA(SNs PRIOR

NOSE CONEISOLATION

VALVE(PRIOR TO SN 69)

Figure SR-37. Distributio



F OR T RAI NI N

G P URP O SE S

ONL Y

S R- 4

MAIBLEED P

WING ANTI-

CREW AIR-CONDITIO

PRESSURIZ

PRESSURIZA

AIR-CONDITIO

MAIN HPBLEED PORT

MAIN LPBLEED PORT

MAIN LP

BLEED PO

AIRINTAKE

ANTI-ICING

AIRINTAKE

ANTI-ICING

FLASHING IFCORRESPONDING

SWITCHOFF

ON IF:• APU BLEED VALVE OPEN IN FLIGHT• APU BLEED VALVE OPEN ON THE GROUND AND POWER LEVER AT A POSITION ABOVE 54˚ FCU

ON IFISOLATION

VALVECLOSED

WARNING PANEL

OVERHEAD PANEL

BLEEDOVHT

BLEEDAPU

ISOL

HEAT

APU

NORM

PRV3PRV2

LPLP

ISOLATION

PASSENGER CREW

BLEED AIR

BAG

AUTOONOFF

HP1

IS

OL

Figure SR-38 Blee



4 9 Figure SR 38. Blee

S R- 5 0

F O


ONL Y

ENG1/ENG 3 PT2TT2 AIR IN SWITCH SENSOR ANTI P

Off Not energized Cl

On Energized Op

On Energized Reg

On Energized Reg

On Energized Fully open (

Off Not energized Clo Off Not energized Cl

Off Not

Table SR-4. NO. 1 AND NO.




ENG 1 ENG 2 ENG

ANTI-ICE

T2 T2P T

PROBE

PRINTED CBOAR



NO. 3 ENGIN


PRESSURE-REGULATINGVALVE

T2 T2P T

PROBE

NO. 1 ENGINE

PRINTED CIRCUITBOARD

ECU HEATEXCHANGER

RAM-AIR INLETANTI-ICING

ENG 1 ENG 2 ENG 3

ANTI-ICE



HP

FLOWLIMITER

F OR T RAI NI N

G P URP O SE S

ON

L Y

S R- 5 3

LE

CENTPEDES

PICCOLO TUBE

AIR C

NO. 1 ENGINE

LP PRESSURESWITCHES

FLOWLIMITER

PRINTEDCIRCUITBOARD

INTAKE LIP

S-DUCT

NO. 2ENGINE

Figure SR-41 No 2



3Figure SR-41. No. 2

S R- 5 4

F O

R T RAI NI N G P UR

P O SE S

ONL Y

WIN

MAIN LP BLEEDAIR (LP1)

LEFT WING PRESSURESWITCH

RIGHT WING

PRESS

URESWITCH

TELESCOPICTUBES

PEDES

TAL

POWER LEVER SWITCHES

FIXEDWING ROOT

LEADING EDGE

PRINTED

CIRCUITBOARD

MANIFOLD

Figure SR-42. Wing



F OR T RAI NI N

G P URP O SE S

ON

L Y

S R- 5 5

LEGENDBLEED AIR

TURBOCOOLERCOMPRESSED AIR

TEMPERED AIRTURBOCOOLERTURBINE AIR

LP WATERSEPARATOR

ANTI-ICINGSENSOR

HP WATERSEPARATOR

CREW AIR-CONDITIONING

DUCT

CREWCOLD AIR

CABIN

COLD AIROR

SUCTION

COLD AIRPRESSURE-

REGULATINGVALVE

COLD AIROR

SUCTION

CABIN

PASSENGER AIR-CONDITIONING

DUCT

LOW-PRESSURE

AIR INLET

NO. 1 ENGINE PYLONRAM AIR INLET

WATERINJECTOR

MAIN HEATEXCHANGER

COMPRESSOR

TURBOCOOLE

TURBINE



5

Figure SR-43. Normal O

S R- 5 6

F O

R T RAI NI N G P UR

P O SE S

ONL Y

LEGENDBLEED AIR

TURBOCOOLERCOMPRESSED


LP WATER

SEPARATORANTI-ICING

DUCT SENSOR

HP WATERS

EPARATION


CIRCUIT

CREWCOLD AIR

CABIN

COLD AIROR

SUCTION

COLD AIRPRESSURE-

REGULATINGVALVE

CABIN


CIRCUIT

LOW-PRESSUREAIR INLET


WATERINJECTOR

MAIN HEATEXCHANGER

PRIMARY HEATEXCHANGER

ECU OVERHTEMPE

SCOMPRESSOR

TURBO-COOLER

TURBINE

COLD AIR

ORSUCTION

Figure SR-44.




P

SNs PRIOR TO 96

BAGGAGECOMPARTMENT

HEATVALVE

THERMAL SWITCH

DUCTCOCK-PIT

INTER-CONNECT

FLAP55%

FULLY HLIMITAT



COLDAIR

GENERATION

SENSOR

RECYCLEDAIR

PASSENGERSENSOR

CREWSENSOR

AMPLIFIER

CABINCONTROLS

PIT

S R- 5 8

F O

R T RAI NI N G P UR

P O SE S

ONL Y

Figure SR-46. Ai

LEGEND

BLEED AIR

RAM AIR

RECIRCULATED AIRWATER

INJECTOR

LOW-PRESSUREAIR INLET

PASSENGERAIR-CONDITIONING

CIRCUIT

PASSENGERCOLD AIR

COLD AIROR SUCTION

COLD AIROR SUCTION

COLD AIRPRESSURE-

REGULATINGVALVE

CABINCOLD AIR

CREW COLDAIR

CREW-AIR CONDITIONING

CIRCUIT

HP WATERSEPARATIO

ANTI-ICINGDUCT SENSOR

LP WATERSEPARATOR

NO. 1 ENGINE

PYLON RAM-AIR INLET

MAIN HEXCHA

CABINPRESSURIZED



S R- 6 0

F O

R T RAI NI N G P UR

P O SE S

ONL Y

PILOT VERTICALSPEED INDICATOR VMO /MMOTEST

PUSHBUTTONS

STANDBY

MACH SPEEDINDICATOR

* PILOT AND COPILOT ELECTRIC MACH SPEED INDIC

ADC 1 ADC 2

RESET

20°c SAT 25°C TAT 175 KTAS

L AFCS FAIL

R AFCS MASTER

AP ENGAGED

L AFCS R AFCS

BRT

NAVIGATION

ADC1

XXX2.5

BUS

A1

RIGHT CIRC

PILOTALTIMETER

STANDBYALTIMETER

PILOT MACHSPEED INDICATOR*

PILOT ID 802 ANNUNCIATORLEFT CIRCUIT-BREAKER PANEL

PILOT EADI

019

46

37

28

5

00 00

10

120

300

100

350

400

8060

6

4

5140

250160

200

1801000 FT PER

MIN

VERTICAL SPEED

UP

DOWN

14

1 4

6

2

2

5

0

5

019

46

37

28

5

ALT

28 720

BARO

100

300

80

350

40060

120

250

140

200

230160

180

6 6

IASKT

180

160

120

180

100

80

LNAV HDG VASL

ASEL

300020

10

20

10

0

00

GS

SPERRY

X IRS

Figure SR-48. Pito



MASTER WARNING S

TABLES

Table Title

MWS-1 Annunciator Illumination Causes ...MWS-2 Fire Panel Illumination Causes ......

MWS-3 Hydraulic Control and Indicator Pan

Illumination Causes .........................

MWS-4 Battery Temperature Indication

Illumination Causes .........................MWS-5 ENG 2 FAIL Illumination Causes ...

MWS-6 ITT Light Illumination Causes ......

MWS-7 Audio Warning Causes ...................

MWS-8 Audio Warning Testing ...................

MWS-9 Configuration Panel and Landing G

Control Handle Illumination Causes

MWS-10 Thrust Reverser Indicator Lights

Illumination Causes.........................

MWS-11 Overhead Panel Light Illumination C





Table MWS-1. ANNUNCIATOR ILLUMIN

ANNUNCIATOR CAUSE FOR ILL

CMPTR 1

FWD

DOORS

CMPTR 2

CMPTR 3

OIL 1

OIL 2

OIL 3

L. AOA

R. AOA

The control switch of the indicated

or MAN position.

The indicated computer has failed failure, internal failure, or incorrect

The light comes on if: (Specific ligh

M880A modification).

The main entrance door is not lockcompartment service door is not lowith this lavatory compartment).

The indicated heating systems are

Angle-of-attack heating has failed.

The oil pressure of the indicated en(1.72 bar).

Chips are detected in the indicated

Red light with M880A



L. PITOTThe indicated heating systems are

Pitot or static pressure probe heati


Table MWS-1. ANNUNCIATOR ILLUMINATION C

ANNUNCIATOR CAUSE FOR ILLUMIN

BUS TIED

HOT

BAT

AUTOSLATS

BAT 1

BAT 2

L. WHL

OVHT

R. WHL

OVHT

An overheat condition is detected in the lef

wheel well.

The indicated battery is not connected to tsystem through its make-and-break switch

The main left and right buses are tied or thcontactor remains closed.

The temperature of one of the batteries ex(65.5°C) for aircraft prior to SN 172 with S

The temperature of one of the batteries ex(71.1°C) for aircraft SN 172 and subsequeto SN 172 with SB-94 applied.

The HOT light located on the battery tempilluminated.

There is a discrepancy between the two slflight/ground contacts.

Th i di b t th t



There is a discrepancy between these twonose and left landing gear flight/ground cogear retraction.


Table MWS-1. ANNUNCIATOR ILLUMINAT


The pressure switch located on the

LP line indicates a pressure drop (4.6 psi [approximately 320 mb]).

The front-to-rear tank transfer valveopen.

The cabin baggage compartment a

A fuel level below 200 pounds is deor G3.

For aircraft SNs 1 to 11—A fuel levdetected in tank group G2 (or belo

LO

FUEL 1

LO

FUEL 3

LO

FUEL 2

XTK 2

OPEN

XTK 2

CLOSED

BAG

ACCESS

FUEL 1

FUEL 2

FUEL 3

The front-to-rear tank transfer valveclosed.

Aircraft with transfer valve XTK2.




detected in tank group G2 (or belopumps 2 are off).

For aircraft SNs 12 and subsequen

FUEL 2




The light comes on if:

The foldable door of the partition at frame when the “No smoking” passenger call sig

One of the two fuel vents is not closed.

The defueling/refueling valve is not closed

The refueling connector access door is no

The refueling control panel access door is

The GRAVITY FUELING switch is set to O

Bus B-2 has failed.

The DEFUELING switch is set to ON.

The vent valve control lever is raised.

The autopilot has failed or has automatica

The thrust reverser clamshell doors are noposition.

NOTEThe light normally illuminates during the th

phase.

AP

FUELING

REV

UNLOCK

AFT CABIN

ISOL

(option 25-21-01)



On aircraft incorporating M880C, when thiaudio warning sounds.


Table MWS-1. ANNUNCIATOR ILLUMINAT


An overheat (446°F or 230°C) is de

outlet of the turbocooling unit.

The turbofan bypass valve is not con the ground or flight with the land

Overheating is detected in one of tducts (air temperature higher than

Overheat is detected in the nose cor equal to 158°F [70°C]).

The APU bleed-air valve is not comswitch off or one of the power lever

The baggage compartment electric

this condition the baggage compar

Steady illumination: Hydraulic systeapplied to the brakes (pressure hig[approximately 18 bars]).

Flashing illumination: When the pa

pressure is between 1,305 to 1,102can be applied only once.

With audio warning:

#2 P BK

ECU

OVHT

COND’G

OVHT

NOSE

CONE OVHT

BLEED

APU

BAG ISOL



With audio warning: • Cabin altitude is greater tha

Without audio warning:

CABIN




The light illuminates and the audio warning

is on the ground, with at least one of the pobeyond 82° and one of the following mode

• The slat/flap control is in CLEAN.

• Flap deflection is higher than or equal

• The airbrakes are not retracted.

• The horizontal stabilizer is out of the arange between –4° 30' and –7° 30'.

• The slats are not extended.

• On aircraft incorporating M880C, the ppulled and the dual braking system is n

T/O

CONFIG

Table MWS-2. FIRE PANEL ILLUMINATION


Fire is detected in the indicated engine coFIRE 1

FIRE 2



FIRE 3


Table MWS-3. HYDRAULIC CONTROL APANEL ILLUMINATION C


The pressure of the indicated pum(approximately 103 bars).

The standby pump selector locatedis not in the normal flight position.

The standby pump cycle time is lo

Pressure supplying the left or righthigher than or equal to 232 psi (apincreasing in system No. 1.

The lights extinguish when the brapressure becomes lower than or edecreasing in system No. 1).

PUMP 1

PUMP 2

PUMP 3

ST BYPUMP

L R

Table MWS-4. BATTERY TEMPERATILLUMINATION CAUSE




The No. 1 and/or No. 2 battery ovewhen the battery internal temperat[48 9°C])

WARM


Table MWS-5. ENG 2 FAIL ILLUMINATION


The aircraft is on the ground and the No. 2

lever is at the 84° setting and the No. 2 engthan 85% N1.

The No. 2 engine S-duct access door is no

A second light is installed on the copilot insaircraft with M880B incorporated.

On aircraft with SB-131, this pushbutton ligwhen the turbine emergency anti-icing valv

ENG 2 FAIL

ECU

A/I

Table MWS-6. ITT LIGHT ILLUMINATION C


The light flashes if the associated engine Ion TFE-731-5AR-1C or 980°C on TFE-731

If the power increase function is used, the 974°C on TFE-731-5AR-1C or 996°C on T



The light indicates the correct operation ofcontrol relays for high altitude takeoff (apprand hot weather conditions (over 18 5°C)

PWRINC

F OR T RAI NI N G

P URP O SE S

ON

L Y

M

W S - 9

Table MWS-

WARNING

TYPE OF

SOUND

VMO /MMO Continuous varying sound with frequency varying

between 660 Hz and 3,330 Hz during a one-second period

Cabin pressure Warning voice CABIN

Fire Continuous two-pitch audiblbe 500-Hz tone for 150 ms and then 555 Hz for 150 ms

Stall Intermittent 1,660-Hz sound (beep beep)—on for 100 ms and off for 100 ms



MW S - 1 0

F OR T RAI NI N G P UR

P O SE S

ONL Y

Table MWS-7.

WARNING

TYPE OF

SOUND

Stall Intermittent 1,660-Hz sound (beep beep)—on for 100 ms and off for 100 ms

Altitude deviation Warning voice ALTITUDE

Horizontal Continuous clacker sound Stabilizer with pulse frequency at

in movement 12.5 Hz



F OR T RAI NI N G

P URP O SE S

ON

L Y

MW S - 1 1

Table MWS-7.

WARNING

TYPE OF

SOUND

Landing gear Warning voice GEAR

Decision height Warning voice MINIMUM

Autopilot (Aircraft Warning voice AUTOPILOT with M880C)



MW S - 1 2


P O SE S

ONL Y

Table MWS-7.

WARNING

TYPE OF

SOUND

Takeoff Configuration Warning voice NO TAKE-OFF

Red lights on warning 770 Hz gong for 0.4s. panel that do not have

their own audiowarnings and ENG 2

FAIL red lights (aircraft with M880C).




Table MWS-8. AUDIO WARNING


VMO /MMO Depressing ADC 1 or ADC 2 push

Cabin altitude Depressing the test pushbutton onaccompanied by illumination of the

panel

Fire Setting the warning panel LIGHTS (all the FIRE lights on the fire pane

Stall On the ground depressing the

STA

Horizontal Action on the pitch trim controlstabilizer inmovement

Landing gear Depressing the landing gear test p landing gear configuration panel

Takeoff On the ground, reproducing config cause the warning panel T/O CON




Table MWS-9. CONFIGURATION PANEL AND LACONTROL HANDLE ILLUMINATIO


The control handle is in the gearup positiogears are not yet uplocked.

The control handle is in the down positiongear is not fully downlocked.

The control handle is in the up position, sp160 knots, at least one of the power leverssetting, but not on STOP for aircraft incorp

at least one of the three gears is not down

Steady illuminationAt least one of the six airbrakes is not in thposition.

FlashingAfter automatic retraction of the airbrakes,handle is in the extended position while theretracted.

Steady illuminationAll the slats are extended.

FlashingOnly the outboard slats are extended.

During any movement of the slats or if anyfailed to either extend or retract.

AIR

BRAKE

SLATS




Table MWS-10. THRUST REVERSERLIGHTS ILLUMINATIO


The clamshell doors are in movemreverser lock latches are released.

The synchronizing bell crank contrreached the fully deployed position

TRANSIT

DEPLOYED




Table MWS-11. OVERHEAD PANEL LIGHTILLUMINATION CAUSES


Steady illumination: It is depressed to on.

Flashing: In the event of automatic shutdowflight/ground contact, ECU overheat, faultyregulation, or starting faults

The light illumination indicates low oil pres

temperature.

The APU generator is off the line.

The APU generator has excitation voltage.

switch is on.

OIL

GEN

IGN 1 IGN 2 IGN 3

MASTER

APU

APU

APU PANEL

DC SYSTEM PANEL

The igniter unit of the indicated engine is e

ENGINES PANEL

The side tank group interconnection valve

FUEL SYSTEM PANE

XTK



One of the 1–3 or 3–1 crossfeed valves is X.BP


Table MWS-11. OVERHEAD PANEL LILLUMINATION CAUS


Green light: The ENG 1, ENG 2 or position, and all the required condiare satisfied (temperature and presand pressure for the No. 1 or No. 3

Amber light (steady): The ENG 1, Ein the on position, and a pressure d

corresponding system.

Amber light (flashes): OverpressureNo. 3 engine, or overtemperature oin the No. 2 engine.

Green light: The WING switch is seconditions for proper operation are

Amber light (steady): The WING swpressure drop is detected in the sy

Amber light (flashes):Overtemperature is detected in the

NOT

On aircraft with wing-brake heatingthe green and amber lights are theset to WING or WING-BRK.

ANTI-ICING

WINDSHIELD

ENG 1 ENG 2 ENG 3

WING



Either pilot or copilot regulation circ

WINDSHIELD


Table MWS-11. OVERHEAD PANEL LIGHTILLUMINATION CAUSES (C


The FASTEN BELTS passenger ordinance

The EMERG LIGHTS selector is in the ONthe aircraft is electrically powered.

The no smoking ordinance sign is illuminat

FASTEN

BELTS

EMERG

LIGHTS

Switch illuminates when associated avionic

INTERIOR LIGHTS

PAN

AVIONIC MASTERSWITC

OFF

LH

AVIONICS

MASTER

OFF

RH

AVIONICS

MASTER




APPENDIX BAppendix B presents a color represthe annunciator lights in the aircraft

Please remove page APP-3, align to thAPP-1 and leave it open for ready r

annunciators are cited in the text.



FlightSafety international


FOR TRAINING PURPOSES ONLY

CMPTR 1

OIL 1

GEN 1

BAT 1

FUEL 1

LOFUEL 1

BLEEDOVHT

CMPTR 2

OIL 2

GEN 2

BUS TIED

FUEL 2

#2 P BK CABIN

LOFUEL 2

ECUOVHT

CMPTR 3

OIL 3

GEN 3

BAT 2

FUEL 3

FUELING

LOFUEL 3

COND'GOVHT

L. PITOT

AP

AILZERO

NOSECONE OVHT

AOAPROBE

HOTBAT

XTK 2OPEN

XTK 2CLOSED

REARDOORS

L AOA

MISTRIM

AILFEEL

BLEEDAPU

ST BYPITOT

L. WHLOVHT

AUTOSLATS

T/OCONFIG

R. AOA

R. PITOT

BAG ISOL

MACHTRIM

R. WHLOVHT

FLAPASYM

BAGACCESS

PITCHFEEL

BRIGHT TEST

FIRELIGHTS

DIM

ITT

°C

0

200

400 600

800ITT

°C

0

200

400 600

800ITT

°C

0

200

400 600

800

IRS

XFR

SGIASM

OFF

HSI ADI

BRT

COMPOS

TGT

GAIN MODEPRESET

GCR RCT

STBYOFF

TEST WXGMAP

RANGE TILT

2510 0

50 100 +15

_15

200300

REVUNLOCK

ENG 2 FAIL PWR

INC

RESET

L AFCS R AFCS

BRT20°c SAT 25°c TAT 175KTAS

L AFCS FAIL

R AFCS MASTER

AP ENGAGED

OFF

LHAVIONICS

MASTER

OFF

RHAVIONICS

MASTER

L RPUMP 1

PUMP 3

PUMP 2

ST BYPUMP

1/1

1/2

0

4

HYDRASKID

TEST

ON

OFF

OFFOFF

ON

AUTO

3

2

1

0

1/1

1/2

0

4

HYDR

QTY PSI X 1000QTY PSI X 1000

3

2

1

0

#1

#1

#2

BRAKE

ST-BYPUMP

GALLEY

ON OFF

PASS ON

HORN SIL

E BATTONERESET

DEPLOYED

TRANSIT

E BAT 1

ON O FF

AUTO

MAN

E BAT 2

ON OF F

AUTO

MAN

ATC 1

ATC 2

* *

* ON AIRCRAFT WITHOUT ELECTRONIC TRANSFER VALVE XTK2, THESE LIGHTS ARE NOT FEATURED.





FLAPS

4020

7

0UP

DN

AIRBRAKE

SLATS

MOVING

LANDING GEAR

TEST

LH NOSE RH

GEAR

PULL

UNLOCK

PUSH

CABINCALL SELCAL

FWD

AFT

LAV

ATNDT

HF 1

HF 2

VHF 1

VHF 2

VHF 3

SPARE

SPARE

1 2

OFFLESS 50° F

WARM HOT180

160

140

120F

OFF

180

160

140

120F

BAT. TEMP.

TEST

120

10080

MASTER

OIL

60

40

20

0

N% 1000

800

600

400

200

0

T5°C

0

20 25

30V –

BAT

GEN

A –100

100200

300

350

?0

+

_

APU

STOPSTART

GENG EN 1 B AT 1 G EN 3

NORMAL

EXT POWER

APU

FUEL SYSTEMENGINES

PITOT

CMPTR 1AUTO

MAN

OFF

IGN 1

CMPTR 2

AIR START

MOTOR-START STOP

CMPTR 3AUTO

MAN

GRD START

OFF

IGN 2 IGN 3

PRESS TO START

HOLD TO MOTOR

PILOT

NAV

LDG

ANTICOL LANDING WINGTAXI

PILOT

FAST

SLOW

OFF

NAV

NAV-LOGO

OFF

ALL

RED

OFF

ST-B Y CO PI LOT ENG 1 ENG 2 ENG 3

WIPER EXTERIOR LIGHTS

ANTI-ICE

IRS 1 HRZN IRS 2 IRS 3

GEN 1E BAT 2TEST

A

BOOSTERNORM

ST-BY

OFF

ENG 2ENG 1

XTK 2

LEVEL

°C

LEVE

XTK

X.BP

LEVELX.BP X.B

BOOSTER

*ON AIRCRAFT WITHOUT ELECTONIC TRANSFER VALUE XTK2, THE OVERHEAD PANEL DOES NOT FEATURE T*A/C <172



FALCON 900 PILOT TRAINING MANUA

Record of Revision No. 4.04

This is a revision of the Falcon 900 Pilot Training Manual, V

A solid vertical line in the margin indicates the content of the

changed. A vertical line adjacent to a blank space indicates

Any page affected by the revision is marked “Revision 4.04”If a page has “Revision 4.04” in the lower left or right cornergin, it is a page in which format only has been changed.

The changes made in this revision will be further explainedtraining course.

FlightSafetyinternational

COURSEWARE SUPPORT—HURST 8900 Trinity Blvd. Hurst, Texas 76053



FALPILOT TRAINING



Courses for the Falcon 900 and other Dassault aircraft are FlightSafety learning centers:

FlightSafety InternationalTeterboro Learning CenterTeterboro Airport 100 Moonachie Avenue

Moonachie, New Jersey 07074Phone: (201) 528-0100Toll-Free: (800) 827-8058Fax: (201) 528-0101

FlightSafety International



FlightSafety InternationalParis Learning Center

FOR TRAINING PURPO

NOTICE

The material contained in this training manual is basfrom the aircraft manufacturer’s Airplane Flight Manuaance Manuals. It is to be used for familiarization and trainin

At the time of printing, it contained then-current informbetween data provided herein and that in publicationsor the FAA, that of the manufacturer or the FAA shall take p

We at FlightSafety want you to have the best trainingsuggestions you might have for improving this manuatraining program.



CONTENTS

Chapter 1 AIRCRAFT G

Chapter 2 ELECTRICAL

Chapter 3 LIGHTING

Chapter 4 MASTER WA

Chapter 5 FUEL SYSTE

Chapter 6 AUXILIARY

Chapter 7 POWERPLAN

Chapter 8 FIRE PROTEC

Chapter 9 PNEUMATIC

Chapter 10 ICE AND RAI

Chapter 11 AIR CONDIT

Chapter 12 PRESSURIZA

Chapter 13 HYDRAULIC

Chapter 14 LANDING GE

Chapter 15 FLIGHT CON

Chapter 16 AVIONICS



Chapter 17 MISCELLAN

CHAPTER 1AIRCRAFT GENERA

CONTENTS

INTRODUCTION ......................................................................

GENERAL..................................................................................

STRUCTURES...........................................................................

General ................................................................................

Fuselage..............................................................................

Empennage..........................................................................

Wing....................................................................................

AIRPLANE SYSTEMS ............................................................

General................................................................................

Electrical Power Systems ...................................................

Lighting...............................................................................

Master Warning System.....................................................

Fuel System........................................................................

Auxiliary Power Unit..........................................................

Powerplant ..........................................................................

Fire Protection.....................................................................

Pneumatics ..........................................................................




Hydraulic Power Systems..............................................................

Landing Gear and Brakes ..............................................................

Flight Controls ...............................................................................

Avionics .........................................................................................

Oxygen System..............................................................................

LIMITATIONS ......................................................................................

General............................................................................................

Weight (Structural).........................................................................

Weight (Performance)....................................................................

Center of Gravity ...........................................................................

Loading ..........................................................................................

Operating .......................................................................................

Airbrakes........................................................................................

Airspeed .........................................................................................




ILLUSTRATIONSFigure Title

1-1 Falcon 900...................................................................

1-2 General Dimensions ...................................................

1-3 Composite Structures..................................................

1-4 Fuselage Sections........................................................

1-5 Nose Cone...................................................................

1-6 Cockpit Layout (Typical).............................................

1-7 Nose Wheel Well Ceiling Door ..................................

1-8 Cockpit Windows........................................................

1-9 Cabin Interior (Typical)...............................................

1-10 Main Entry Door .........................................................

1-11 Emergency Exit...........................................................

1-12 Cabin Windows..........................................................

1-13 Aft Fuselage ...............................................................

1-14 Baggage Compartment................................................

1-15 Empennage.................................................................

1-16 Wing ............................................................................





INTRODUCTIONThis training manual provides a description of the majorinstalled in the Falcon 900. This information is intended does not supersede, nor i s it meant to substitute for, anyor operating manuals. The material presented has been pdata. All subsequent changes in airplane appearance or ered during academic training and subsequent revisions

Chapter 1 covers the structural makeup of the airplane antems. It also contains operating limitations and a pictoria

Appendix B in this manual displays all light indicationsreference while reading this manual.

AIRCRA



The airplane is certificated in the transportcategory for two-pilot operation and is ap-proved for all-weather operation to a maximumaltitude of 51,000 feet. Extended overwater op-erations are authorized when the requiredequipment is installed and operational.

STRUCTURES

GENERAL

The structural design of the airplane (Figure 1-1) conforms to the fail-safe structural concept.

The structure meequirements and, thelife limit. The airpmonocoque construaluminum alloy, stKevlar, fabric, andondary structure. Tfuselage, wings, ansion on the fuselagwindows. The winfuel tanks; fuselagpart of the fuselageshown in Figure 1areas of the airplanstructural materials





63 FT 5 IN.

(19.33 m)

25 FT 11 IN.

(7.90 m)

66 F(20

25 FT 11 IN.

(7.90 m)

66 F

(20

WITHOUT SATCOM ANT

WITH SATCOM ANTENN

TECHNICAL CHARACTERISTICS

OVERALL LENGTH 20,21 m 66 FT 4 IN.

OVERALL HEIGHT (EMPTY WEIGHT)

BEFORE SB F900-176 7,55 m 24 FT 9 IN.

AFTER SB F900-176 7,67 m 25 FT 2 IN.

SPAN 19 33 m 63 FT 5 IN

PASSENGER CABIN VOL

COCKPIT VOLUME

MAIN ENTRANCE DOOR

FLOOR HEIGHT ABOVE G

(MAXIMUM HEIGHT)




METAIRCRA

BAGGAGECOMPARTMENT

DOOR SKIN(METAL ON

AIRCRAFT SN 80>)

LEGENDKEVLAR

CARBON

KEVLAR AND CARBON

METAL ONAIRCRAFT SN 157>



FUSELAGE

General

The fuselage is of all-metal, semi-monocoqueconstruction with circu lar bulkheads. It is di-vided into three major sections (Figure 1-4).The nose section extends the length of theradome to the forward cockpit bulkhead. Thecenter section extends from the forward cock-pit bulkhead to the baggage compartment par-tition and includes the cockpit, passenger cabin, lavatory, wing attach points, and frontand rear fuel tanks. The aft fuselage sectionincludes the baggage compartment and the

rear structure, which bears the empennage,the aft equipment compartment, the APU, andthe three engines. The baggage compartmentis pressurized and accessible in flight.

Nose ConeWhen unlatched by five latches, thized nose cone can be slid forward anit can then be swung upwards for inccess and locked open by the action opensating rod (Figure 1-5). The nhouses radar, avionics, and other eq

Center Section

General

The center section is pressurized andthe cockpit, passenger cabin, and

(See Figure 1-4). The cockpit seats tThe main entry door is located on thof the airplane immediately aft of pit. The emergency exit is located onside of the passenger cabin over The lavatory is normally located passenger cabin.


g t in

Revision 3 FOR TRAINING PURPOSES ONLY

Figure 1-4. Fuselage Sections

NOSECONE

CENTERSECTION

AFTFUSELAGE

COCK-

PIT

PASSENGER

CABIN

LAV BAG

COMP

AFT

EQUIPCOMP

APCOMPAR

NO. 2 ECOMPARWITH TH

REVER

EMERGENCYEXIT (RIGHT SIDE)

MAINENTRYDOOR FUEL TANKS

BAGGAGEDOORLEGEND

PRESSURIZED AREA

EMPENNAGE




Figure 1-8. Cockpit Window

Figure 1-7. Nose Wheel Well Ceilin



Cabin Windows

Twenty-four stretched acrylic windows areinstalled in the cabin (Figure 1-12). The eighthwindow aft on the right side is installed in theemergency exit.

Lavatory

The rear lavatory iof the passenger cwith outside drain (rinet on the left sidtank. The tank is fiis drained through


Figure 1-11



Aft FuselageGeneral

The aft fuselage section (See Figures 1-4 and 1-13) includes the baggage compartment and therear structure which bears the empennage, the aftequipment compartment, the APU, and the threeengines. The pressurized baggage compartmentis located in the forward part of the aft section

and is accessible in flight. The unpressurized aftequipment compartment is located immediatelyaft of the baggage compartment and houses theNo. 2 engine electronic fuel computer andhydraulic and air-conditioning components.Access to the aft equipment compartment isthrough a door with an attached step ladder onthe underside of the airplane. The door isconnected to the REAR DOORS (open) warning

light. The auxiliary power unit is located in afire-proof compartment under the No. 2 engineair intake.

Baggage Compartment

Access to the pressurized baggage compart-ment (Figure 1-14) is through the door lo-cated in the aft partition of the lavatory and the

exterior door on the left side of the airplane.The exterior door closes electrically and has

an integral ltacts a micrstep, allowinThe door haopening is aREAR DOO

The exteriorlocated insid

of the door. NF ORE CLOSTEP. The dtated to unlo

When tmotor r

the mots i t i on iLOCKEbutton cess panrectly f

EMPENN

The empennaizontal and v




stabilizer is mounted

away from airflow No. 1 and No. 3 engtical and horizontal spars and stressed-sking edge of the horiable for pitch trim, operated jackscrewthrough normal trim

WING

The Falcon 900 winble-sweep tapered don the fuselage. It haft spars sandwichand lower load-bearcontrols attached to

• One inboard aedge slat

• Three airbrakface

• Two flaps on

• One aileron

The wing box struc

gral (wet) fuel tankof the box supports forward spar suppoing-edge slats.

AIRPLANE

GENERALThe following is amajor airplane sysDetailed descriptiosystems are containchapters of this trai




Two batteries provide normal electtem backup power and a power sourternal engine starting. A ground r

provides for electrical supply fromnal DC ground power unit. Equipmrequires stabilized 115-VAC, 400-VAC, 400-Hz power is provided bywithin each component that require Aor by inverters installed as custome

LIGHTING

The Falcon 900 has standard navigticollision, wingtip strobe, landinglights. A wing ice inspection light ison each side of the fuselage. Interioincludes cockpit, cabin, lavatory,compartment and nose cone lightinglighting includes general illuminspecific lighting for instruments and ing. Cabin lighting provides illumiwarning signs and specific area illufor passenger safety and convenien

MASTER WARNING SYST

The Falcon 900 warning system proving to the crew of airplane equipmfunctions, indications of unsafe conditions which require immedition, and indications that a particulis in operation. Aural warning is aldraw attention to selected situationster warning panel is mounted in theter instrument panel.

FUEL SYSTEM

During normal operations, the Falcosystem consists of three separate suthe left, center, and right. Each snormally supplies fuel to its respegine. Interconnect and crossfeed vafuel transfer between tanks and enfeed from any tank in the event of fuance or boost pump failure. The airpbe pressure- or gravity-refuele d.

Fuel storage consists of three tank g

left wing and center wing, the forwarfuselage, and the right wing and ce


g t in


Figure 1-15. Empennage



tanks. Total usable fuel capacity is 19,065

pounds; an additional 119 pounds is unusable.Low-pressure fuel is supplied to the engine-driven fuel pumps by combined pressurizationand boost pumps. In addition, the submergedboost pumps supply fuel for crossfeed opera-tions, tank-to-tank transfer, and motive-flowfuel for fuel transfer. Fuel quantity is monitoredby a DC-operated capacitance system. Single-point pressure refueling may be accomplished

for full or partial tanks. Gravity refueling portsare located in the left and right wings. Drainvalves are provided to check for fuel contam-ination. All tanks are automatically pressurizedby low-pressure bleed air as soon as either theNo. 1 or the No. 2 engine is started.

AUXILIARY POWER UNIT

A Garrett auxiliary power unit is located in afire-proof compartment in the tail cone beneaththe No. 2 engine air intake. It is certificated forground use only. The APU provides DC elec-trical power when the engine generators arenot on the line, charges the airplane batterieswhen the unit is operating, and may be used toassist engine starting to prolong battery life. The

APU provides reduced starting time, startingwhen the battery is low, and satisfactory coldweather starting. In addition, the APU suppliesbleed air to the airplane’s environmental sys-tem for ground heating and cooling.

POWERPLANT

Thrust is supplied by two aft-fuselage, pylon-

mounted engines and one centerline-mountedengine inside the tail cone. The engines aremanufactured by Garrett Turbine Company atPhoenix, Arizona, a division of the Allied-Signal Aerospace Company. The engines aredesignated TFE731-5AR-1C or TFE731-5BR-1C and are forward fan two spool (low and

efficient fuel sched

an electrohydromecdigital engine electfuel computer automnomical and precisethe entire spectrumrequirements. High-is extracted from tsurization, air conother airplane syste

is equipped with a t

FIRE PROTEC

Fire/overheat can bgines, auxiliary powgear wheel wells. Thas an optical smok

warning is displayetinguisher panel andaudible warning ishorn. Wheel well oof a hot brake) is lights only. Five fixthe United States) ftrolled by the extingcated on the cockp

For the No. 1 and Ntle is fired. For theguisher bott les arbecause of the greabottle is for use inpartment or APU atinguishers are loccabin, respectively

PNEUMATICSHigh- and low-preextracted from eachused to maintain thpit temperature andengine is capable of




shutoff valves. High-pressure bleed air is con-

trolled by valves activated from the cockpit.Bleed sys tem opera t ion i s moni to redby sensors that cause illumination of lights onthe master warning panel.

ICE AND RAIN PROTECTION

The Falcon 900 is approved for flight into

known icing conditions when the requiredequipment is installed and operational. Thewing leading-edge slats and the No. 2 engineS-duct are anti-iced by engine bleed air sup-plied from the engine low-pressure and high-pressure bleed ports. Each engine nacelle isanti-iced from the related engine high-pressurebleed-air port. An electropneumatic valve ei-ther isolates or supplies the system. The three

front windshield panels are heated by 28 VDCfrom two separate buses: A3 for the pilot’s win-dow and the left half of the center windowand B3 for the copilot’s window and the righthalf of the center window. The side windowsare electrically heated by two bus systemsseparate from the windshield heating: bus A2supplies power to heat the forward side win-dows, and bus B1 supplies heating power for

both aft windows. The pitot-static, angle-of-attack, and temperature probes are electri-cally heated from buses A1 and B2. The pilot’sand copilot’s windshield wipers are poweredfrom buses A2 and B2, respectively.

AIR CONDITIONING

Air supply for air conditioning is taken from

the low- and high-pressure bleeds of each en-gine. The air passes through a ram-ai r heat ex-changer and then an environmental controlunit (ECU). Precooling of engine bleed air isaccomplished in the ram-air heat exchanger ei-ther by ram air when in flight or by the turbofanwhen on the ground or during slow airspeeds

move moistu

can be usedwhen on the be used whcontrol of temof malfuncti

PRESSU

The cabin isthrough the matic modecabin altitudtude of 51,00maintained uferential preautomaticallvalves instal

tory above tBoth valves operate togecabin and agovern the emosphere. Sthe outflow preset valuepressure safe

trol of pressof malfuncti

HYDRAU

The airplanesystems whian auxiliarydraul ic f lu

(NATO codeMain hydraself-regulatsponding enpump delive2,987 ±50 pmotor drive




actuation of the landing gear, slats, flaps,

wheel brakes, emergency and parking brakes,a i rbrakes, nosewheel s teer ing, a i leron,elevator , rudder servoactuators, ai leronArthur Q unit and elevator Arthur unit, andthrust reverser.

LANDING GEAR AND BRAKES

The Falcon 900 has retractable tricycle landinggear consisting of a dual-wheel main gear anda dual-wheel, steerable nose gear. The landinggear, wheel brakes, and nosewheel steering areall actuated by hydraulic pressure. When re-tracted, all gears are completely enclosed bydoors. The nose gear will turn 60° either sideof center and features an antishimmy system.The nose gear is self-centering after lift-off.

The antiskid system is available with the nor-mal braking system only; it modulates brakepressure from touchdown to low speed to min-imize braking distances. The hydraulic brakesare used for normal and emergency braking andfor parking.

FLIGHT CONTROLS

Primary flight controls include ailerons, rud-der , and elevators. All are hydraulical lyboosted but can be manually actuated in theevent all hydraulic systems fail. Aileron andrudder trim is hydraulic, initiated by electricmotors. All gust damping is hydraulic, but isnot dependent on hydraulic pressure, and willoperate automatically if a hydraulic systemfails. A hydraulically actuated, autopilot servo-

controlled yaw damper dampens yaw oscilla-tions. The movable horizontal stabilizer istrimmed electrically. The secondary flightcontrols consist of trailing-edge flaps, lead-ing-edge slats, and airbrakes. They are elec-trically controlled and hydraulically actuated.

The pitot-static sy

probes, six static aiperature sensor, whp u t e r s , s e l e c t e d associated pilot and

The standard automcontrol system incluautopilot and flighterence system, and t

flight instrument syfunction can be usetopilot with the pilsatisfy the flight digrammed, or the auautomatically steeflight director commyaw damper systemthe autopilot and ma

out the autopilot en

The Sperry flight comprehensive comuse of multiple naviand blends them insition, using the betype of sensor.

Falcon 900 commucludes interphoneground, and naviga

OXYGEN SYS

The Falcon 900 oxygen for crew and pas

pressure steel cylentrance door under oxygen is availablethrough a quick-donregulator and microxygen is available aaltitudes or manually




LIMITATIONSGENERAL

The limitations presented in this chapter focusprimarily on the operational capabilities of the airplane. Specific system limitations areprovided in the individual systems’ chapterswith the exception of instrument markings

and placards which are presented in this chap-ter. Refer to the FAA-approved AFM for com-plete limitations listings.

WEIGHT (STRUCTURAL)

Maximum ramp:

5AR .................... 45,700 lb (20,730 kg) or

46,700 lb (21,183 kg)5BR ........................ 46,700 lb (21,183 kg)

Maximum takeoff:

5AR .................... 45,500 lb (20,639 kg) or46,500 lb (21,092 kg)

5BR ........................ 46,500 lb (21,092 kg)

Maximum landing .... 42,000 lb (19,051 kg)Maximum zero fuel:

5AR .................... 28,200 lb (12,800 kg) or30,870 lb (14,002 kg)

5BR ........................ 30,870 lb (14,002 kg)

Minimumflight weight.............. 20,700 lb (9,390 kg)

WEIGHT (PERFORMANCE)

General

The approved maximum weights indicated

• Takeo

• Brake

• Climb

Landing

The landing

• Appro

• Availa

CENTER

General

The center-opercent of MAlanding gear

of gravity. Rechart in the “

Datum

Datum is 25%plane exteristation (FS) 0 is the forw

Mean AeLength is 11

Zero perce(9,957 mm)

LOADINGThe airplanwith the Cethe “L imi tInformationweight and




OPERATING

Types of OperationThis airplane is certificated in the transportcategory and is eligible for the followingkinds of operation when the appropriateequipment and instruments require d by theairworthiness and/or operating regulationsare installed and approved and in operable condition:

• Day and night VFR if permitted by theflight regulations of the country overwhich the airplane is flying

• I F R a n d a u t om a t ic a pp r o ac h e s t oCategory I and II weather minimums

• Extended overwater operation

• Flight into icing conditions

Altitude

Maximum operating altitude is 51,000 feet.

Flight ManeuveringLoad Factors

Flaps retracted ..................... +2.53 to –1 gFlaps extended ......................... +2.0 to 0 g

These load factors l imit the permissiblebank angles in turns and the severity of pullupmaneuvers.

Takeoff and Landing

Weights............. See Weight (Performance)in Limitations, this chapter.

Airport pressure

altitude .................. –1,000 to +14,000 feet

Runway slope ................................. ±2.5%

Pressure alti

>10,000 feetAirplanes fittedapproved for 22

Tailwind componen

Airplanes fittedtires approved fand/or 225 mph

Ambient temperatu

Runway surface ....

Enroute

Ambient temperatu

One or two enginesinoperative ...... Re

Minimum FligThe minimum flighcopilot.

AIRBRAKES

Airbrakes must not b300 feet AGL.

AIRSPEED

GeneralUnless otherwise sare expressed in teInstrument error is





CHAPTER 2ELECTRICAL POWER SY

CONTENTS

INTRODUCTION......................................................................GENERAL .................................................................................

DC POWER SYSTEM...............................................................

General................................................................................

Distribution Buses...............................................................

Protection............................................................................

Indication and Warning.......................................................

Batteries ..............................................................................

APU Generator ...................................................................

Generators...........................................................................

Ground Power Unit.............................................................

DC POWER DISTRIBUTION...................................................

General................................................................................

Priority Distribution............................................................

Avionic Masters..................................................................

ELECTRICAL POWER SOURCES..........................................

LIMITATIONS..........................................................................




ILLUSTRATIONS

Figure Title

2-1 DC Power System Components and Locations...........

2-2 DC Distribution Buses................................................

2-3 Circuit-Breaker Panels................................................

2-4 Overhead Switch Panel ...............................................

2-5 Warning Panel .............................................................

2-6 Battery Installation and Ventilation.............................

2-7 Battery Temperature Monitoring System....................

2-7A Battery Temperature Indicator (A/C ≥172).................

2-8 Batteries Installed and Connected, All Switches Off..

2-9 Battery 1 Switch On, All Other Switches Off.............

2-10 Battery 2 Switch On, All Other Switches Off.............

2-11 Both Battery Switches On, Bus Tie Open...................

2-12 Battery 2 Switch On, Bus-Tie Switch Closed.............

2-13 Battery Bus Circuits (Typical) ....................................

2-14 APU Generator Installation.........................................

2-15 APU Generator Operation...........................................

2-16 Generator Cooling Airflow.........................................

2-17 Generator 3 Operation.................................................

2-18 Generators 2 and 3 Operation .....................................

2-19 Normal Flight Configuration ......................................




TABLE

Table Title

2-1 Electrical Power Sources .....................................................




INTRODUCTION

This chapter describes the electrical power systems of the Falcoused for voltage, amperage, and tolerance are for illustrative pube obtained from the manuals and publications issued by, or onturer, the certification agency of the country of origin, and the F

GENERAL

The Falcon 900 uses DC power for control,operation, and indication of the various systemsinstalled in the airplane.

The Falcon 9conventional26-volt requi

ELECTRICAL POW




DC POWER SYSTEMGENERAL

The DC power system of the Falcon 900 is aconventional 24- to 28.5-volt system, which hasfour power sources—batteries, engine-drivenstarter-generators, and an APU-driven starter-generator for ground operation only. Provisions

are also incorporated to provide DC power froma ground power unit (GPU).

The batteries provide the basic source of DCpower to the entire distribution system, surgedamping for the generators, as well as powerfor starting the engines. The batteries are alsocapable of an emergency in-flight source of power for a limited period if the engine-driven

generators fail.

Any engine-driven generator is capable of pow-ering the entire DC system, providing batterycharging, and, through a logic system, assist-ing engine starting.

The APU-driven sta

ble of powering the tion to charging theengine starting, whground.

An approved GPU mperiods to power thmaintenance and serbe used for engine st

to charge the batterbattery charging systaircraft.

Figure 2-1 depicts theof the DC power syste

DISTRIBUTIO

DC power distributiobus system consistinas follows: battery bus, right main busand bus B2 (Figure 2

OVERHEADPANEL

CIRCUIT-BREAKERPANEL

STARTER-GENERATORS

STARTER-

GROUNDRECEPTACLE

BA



R ev i s i on 3

2 - 3

F OR T RAI NI N G P URP O S E S

ONL Y

BUS A1 BUS A2

START BUS

225

BUS

CENTERCIRCUIT-BREAKER

PANEL

LEFT CIRCUIT-BREAKER

PANEL

(14P)

(4PA)

TO EXTPOWER

CONTACTOR

FLIGNOR

LEFT MAIN BUS

MAIN BUS-TIEROTARY SELECTOR

1 3 0 A

1 3 0 A

1 5 0 A

1 5 0 A

1 5 0 A

8

0 A

T

IEDL B U S

T I E D

S T A NDB Y HY DRA U

L I C

P UMP A 5

P I L OT F R ONT

WI ND S HI E L DA 3

GA L L E Y 2 B A R

A 6

GA L L E Y 1 B A R

A 4

AVIONICMASTER

OFF

L AVMASTER

0

2025

30V –

BAT

GEN

Figure 2-2. DC Dis



2 - 4

F O

R T RAI NI N G P URP O

S E S

ONL Y

Figure 2-3. Circuit




The DC distribution system normally operates

as independent left and right systems. These twosystems may be interconnected, when and if required, by a selective bus-tie system. The bustie must be selected closed for APU and enginestarting. The bus tie automatically closes whena GPU is connected and selected for operation.This will be described in detail later in thischapter.

PROTECTION

Circuit protection is provided by conventionaltrip-free circuit breakers (Figure 2-3) locatedabove the overhead panel. Feeder cables are pro-tected by current limiters. Special anticrush

sheathing is

which can ptemperature o

INDICAT

Voltmeters anpanel (Figurethe DC pow

directly connThe ammetetrolled by anselector switLights on thAppendix B)the system.




Figure 2-5. Warning Panel

BAT 1

NOTE:

ON AIRCRAFT WITHOUT TRANSFER VALVE XTK 2, LIGHTS “XTK 2 OPEN” AND “XTK 2 CL




BATTERIES

GeneralTwo 26-volt, 23-ampere-hour, steel-cased,nickel-cadmium batteries are mounted near themain electrical box (Figure 2-6) in the rear com-partment. Batteries with a capacity of 36 amperehours are also available as an option. Eachbattery contains 20 cells. The batteries are

connected by standard quick-disconnectadapters.

VentilationThe batteries are ventilated on the ground andin flight. An electrical battery blower (Figure2-6) provides ventilation on the ground. Theblower forces air through the battery cases,

which exhausts into a Y-shaped duct. The duct

protrudes thr

protruding sesure in the vcontinuous ve

On the grounwhen the BAselector switc

MonitorinAircraft <1

Battery tempby probes intemperature ted to a dual-located on thinstrument p

degrees up

ILLUMINATESAT 120°F

LIGHT

LIG150OR

OFF

180

160140

120

°F1 2

BAT.

WARM HOT

LESS 50°F

TEMP.

OFF

180

160140

120

°F

TEST

A/C <172



green, amber, and red. The green range extends

from 100 to 120°F, the amber range from 120 to150°F, (SNs 132 and subsequent to 160°F) and thered range from 150 to 190°F (SNs 132 and subse-quent from 160°F to 190°F). The dial face has twolights, one amber and one red. The amber lightcomes on when a battery temperature exceeds120°F. The red light comes on when a battery tem-perature exceeds 150°F (SNs prior to 172 withoutSB-94) or 160°F (SNs 172 and subsequent).

This red light operates in conjunction with theHOT BAT light on the warning panel (AppendixB). The instrument may be tested by a push-button switch labeled “TEST” near the gage.When this switch is pushed and held, the needlesmove up scale, and the amber and red lightscome on as the needles pass through 120 and150°F, respectively.

The word “OFF” is printed at the top of eachscale. If the sensors on the battery are dis-connected, the affected scale needle will go tothe OFF position. The battery-temperature-sensing system includes an extended-rangeselector push-button switch, labeled “LESS50°F,” located near the gage. When battery tem-perature is low, pushing this button adds 50°F to

the battery temperature indication, so subtract50°F from the indicated test reading to obtainthe actual battery temperature.

Aircraft ≥172Temperature level is sensed by a thermistorinstalled at a corner of each battery between two

elements. The thermistors are connected to abattery temperature indicator on the copilotinstrument panel (Figure 2-7A). This thermistorcontrols a digital temperature indicator graduat-ed from 32° to 200°F.

When the relevant battery temperature is <32°F

A HOT light illumina

vant battery temperatcoupled to the HOTpanel.

A TEST test button lallows testing of the tor. If the test is corTEST GOOD otherw

In case of thermistocuit (thermistor disdisplays ERR.

ControlThe batteries are coposition magnetic sw

“BAT 2” on the ove2-4). These switchposition when the aswitch opens due to ebatteries. The switchfor the make-and-brethe on (up) position,the trip has cleared.

When both battery sboth batteries are inoutput is supplied dibattery bus through for each battery. Batfor certain essential scribed later under DC


BHOT

8

8



2



2- 1 0


ONL Y

NORM

EXT POWER

OIL

APU STOP

START

IRS 1HRZNIRS 2IRS 3

GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

A B

GEN 3

GEN

TO

AMMETER

DC SYSTEM

OVERHEAD PANEL

B AT 1 B US

GEN 1 GE

WARNINGPANEL

REVERSE

GEN 1 BAT 1 GEN 3 GEN 2 BAT 2 APU

NORM

EXT POWER

OIL

APU STOP

START

IRS 1HRZNIRS 2IRS 3

GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

A B

GEN 3

GEN

G1

GENERATOR 1

TO

AMMETER

G2

GENERATOR 2

TO

AMMETER

G3

GENERATOR 3

START

CONTACTOR

TO AMMETER

G

START CONTACTOR

BATTERY 1

CONTACTOR

BATTERY 1

START

CONTACTOR

GPU CONTACTOR

APU

TO

AMMETER

TO

AMMETER

FIRE

EXTINGUISHERS

COCKPIT

DOME LIGHT

TO

AMMETER

GPU RECEPTACLE

BATTERY

BUS

START

CONTACTOR

BATTERY 2

DC SYSTEM

OVERHEAD PANEL

BATTERY 2

CONTACTOR

STARTING

BUS

B AT 1 B U

GEN 1 G

WARNINGPANEL

REVERSE

CURRENT RELAY

REVERSE

CURRENT RELAY

REVERSE

CURRENT RELAY

MAKE-AND-BREAK SWITCH

REVERSE CURRENT RELAY


BUS-TIE

ROTARY

SELECTOR

CONDITIONS:

BUS TIE OPEN, POWER

SELECTOR SWITCH

IN NORMAL

BAT 1 BAT 2

GEN2GEN3 APUGEN1

Figure 2-9. Battery 1 Switch On, All Other Switches



R ev i s i on4

F OR T RA

I NI N G P URP O S E S

ONL Y

OIL

APU STOP

START

IRS 1HRZNIRS 2IRS 3

GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

A B

GEN 3

GEN

DC SYSTEM

OVERHEAD PANEL

BA T 1

GEN 1

WARNING

PANEL


NORM

EXT POWER

OIL

APU STOP

START

IRS 1HRZNIRS 2IRS 3

GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

A B

GEN 3

GEN

G1

GENERATOR 1

TO

AMMETER

G2

GENERATOR 2

TO

AMMETER

G3

GENERATOR 3

START

CONTACTOR

TO AMMETER

G

START CONTACTOR

BATTERY 1

CONTACTOR

BATTERY 1

START

CONTACTOR

GPU CONTACTOR

APU

TO

AMMETER

TO

AMMETER

FIRE

EXTINGUISHERS

COCKPIT

DOME LIGHT

TO

AMMETER

GPU RECEPTACLE

BATTERY

BUS

START

CONTACTOR

BATTERY 2

DC SYSTEM

OVERHEAD PANEL

BATTERY 2

CONTACTOR

STARTING

BUS

B AT 1

GEN 1

WARNING

PANEL

REVERSE

CURRENT RELAY

REVERSE

CURRENT RELAY

REVERSE

CURRENT RELAY


REVERSE CURRENT RELAY


BUS-TIE

ROTARY

SELECTOR

CONDITIONS:

POWER SELECTOR SWITCH

IN NORMAL. ALL

GENERATORS OFF


Figure 2-11. Both Battery Switches On, Bus Tie O




g t a et international



2-14 FOR TRAINING PURPOSES ONLY

N O R M

E X T P O W E R

O I L

A P U

S T O P

S T A R T

I R S

1

H R Z N

I R

S

2

I R S

3

G E N 1

G E N

1 B A T 1 G E

N

3

B A T 1

B A T 2

F L I G H T N O R M

G E N 2

G E N 3

G E N

G 1

G E N E R A T O R 1

T O

A M M E T E R

G 2

G E N E R A T O R 2

T O

A M M E T E R

G 3

G E N E R A T O R 3

S T A R T

C O N T A C T O R

T O A M M E T E R

G

S T A R T C O N T A C T O R

B A T T E R Y 1

C O N T A C T O R

B A T T E R Y 1

S T A R T

C O N T A C T O R

G P U C O N T A C T O R

A P U

T O

A M M E T E R

T O

A M M E T E R

F I R E

E X T I N G U I S H E R S

C O C K P I T

D O M E L I G H T

T O

A M M E T E R

G P U R E C E P T A C L E

B A T T E R Y

B U S

S T A R T

C O N T A C

T O R

B A T T E R Y 2

D C S Y S T E M

O V E R H E A D P A N E L

B A T T E R Y 2

C O N T A C T O R

S T A R T I N G

B U S

B R I G H T

L I G H T S

F I R E

T E S T

D I M

B A T 1

B U S T I E D

G E N 1

B A T 2

H O T

B A T

G E N 3

G

E N 2

W A R N I N G

P A N E L

R E V E R S E

C U R R E N T R E L A Y

R E V E R S E


R E V E R S E


M A K E - A N D - B R E A K S W I T C H

R E V E R S E C U R R E N T R E L A Y

M A K E - A N D - B R E A K S W I T C H

B U S - T I E

R O T A R Y

S E L E C T O R

1 5 0 A

1 3 0 A

8 0 A

1 5 0 A

1 5 0 A

1 3 0 A

1 3 0 A

8 0 A

1 3 0 A

B U S A 1

B U S A 2

B A T T E R Y P O W E R

A P U G E N E R A T O R P O W E R

L E G E N D W

I N D O W S

A 3

G A L L E Y 1 B A R

A 4

G A L L E Y 2 B A R

A 6

H Y D R A U L I C

S T A N D B Y P U M P

A 5

L H M A I N B U S

B U S - T I E

R E L A Y

2 2 5 A

R H M A I N B U S

B 1

B 2

W I N D O W S

B 3

G R O U N D

A P U

A

B

C O N D I T I O N S :

B O T H B A T T E R Y S W I T C H E S

O N , B U S T I E C L O S E D .

B O T H B A T T E R I E S

C H A R G I N G

G E N

2 B A T

2 A P U

F i g u r e 2 1 5

A

P U

G e n e r a t o r O p e r a t i o n




and the APU generator if all four generators are

on and the bus tie is closed, or GEN 2 and theAPU generator are automatically paralleled onthe right main bus when on and the bus tie isopen. An overvoltage and overcurrent protectionsystem in the associated GCUs automatically tripthe affected generator whenever voltage is above32 volts or if output current exceeds design limits.

ControlEach generator is controlled by a two-positionmagnetic switch located on the DC SYSTEMsection of the overhead switch panel. The switchpositions are unlabeled. They must be on (up)for all operation, including engine starting. Gen-erator faults detected by the GCUs, such as over-voltage or overcurrent, cause these switches toautomatically trip to the off (down) position.Physically moving a tripped generator switch tothe on (up) position will reset the generator, pro-vided that the fault no longer exists. The gen-erator switches should be turned off if an engineis inoperative and windmilling in flight.

A guarded line disconnect control switch foreach generator is located on the main electricalbox in the rear compartment (Figure 2-13).These switches provide the control circuits andequalization for the generator line contactors(RCRs) and are for maintenance use only.

Automatic Cabin ElectricalLoad-Shed SystemIn the event of the loss of one generator’s out-

put, certain A6 bus items such as galley, lava-tory, and cabin entertainment system operationwill automatically be load-shed.

After proper electrical load reduction by thecrew an AUTO LOAD SHED switch located onth il t’ id l b l d i th

The auto loa

the ground/flon the ground

IndicatioThe voltmetetion of the ovcation of voltbus tie is open

bus tie is closselective thrswitches locaswitch is labthe right swAPU.” MoviAPU positionthe selected uBAT position

ammeters indciated batteryBAT position

Three amber “GEN 1,” “Gindication foaffected lightcurrent relay

is off the lineto open at the

Operatio

Figure 2-17 been started,terminated aconnecting gits associatedtie is closed,mitted to theB1 and B2 bthrough its mtery 1 conta




G E N

1 B A T 1 G E N

3

G E N

2 B A T 2 A P U

S T O P

I R S

1 H R Z N

I R S

2

I R S

3

G E N 1

B A T 1

B A T 2

F L I G H T N O R M

G E N 2

G E N 3

E N

G 1

G E N E R A T O R 1

T O

A M M

E T E R

G 3

G E N E R A T O R 3

S T A R T

C O N T A C T O

R

T O A M M E T E R

G

S T A R T C O N T A C T O R

B A T T E R Y 1

C O N T A C T O R

S T A R T

C O N T A C T O R

G P U C O N T A C T O R

A P U

T O

A M M

E T E R

G P U R E C E P T A C L E

B A T T E R Y

D C S Y S T E M

O V E

R H E A D P A N E L

B R I G H T

L I G H

T S

F I R E

T E S T

D I M

B A T 1

B U S T I E D

G E N 1

B A T 2

H O T

B A T

G E N 3

G E N 2

W A R N I N G

P A N E L

R E V E R S E


R E V E R S E


R E V E R S E


M A K E - A N D - B R E A K S W

I T C H

B

U S - T I E

R

O T A R Y

S E

L E C T O R

1

5 0 A

1

3 0 A

8 0 A

1

5 0 A

1

5 0 A

1

3 0 A

B U S A 1

B U S A 2

L E G E N D

W I N D O W S

A 3

G A L L E Y 1 B A R

A 4

G A L L E Y 2 B A R

A 6

H Y D R A U L I C

S T A N D B Y P U M P

A 5

L H M A I N B U S

B U S - T I E

R E L A Y

2 2 5 A

A P U

A

B

G E N E R A T O R P O W E R

B A T T E R Y P O W E R

L E C T O R

A T T E R Y

A T I V E




Figure 2-18 depicts that the No. 2 engine hasbeen started. The start was assisted by genera-tor 3, since the APU generator is not operating(see Chapter 7, “Powerplant,” for starting

details). As the No. 2 engine start terminates, thegenerator 2 RCR closes, connecting generator 2to the right main bus; it automatically parallelswith generator 3 since the bus tie is closed. Thewarning panel shows that the GEN 2, GEN 3,BAT 1, and BAT 2 lights are all off while the

GEN 1 and BUS TIED lights remain on.

Figure 2-19 depicts that all three engines are op-erating and all three engine-driven generators areon line. The airplane is ready for takeoff; there-fore, the bus-tie switch is in the FLIGHT NORMposition, separating the left and right distribu-tion buses into separate and independentsystems.

Generators 1

the left mainA1 and A2, and charginging the right mB2 buses as it is also powbattery 2. Therator lights,TIED light a

ing configura

If the No. 2 e2 fails, the bpower to theB1 and B2 bbattery 2.

2 OVERHEADPANEL



R ev i s i on4

2- 1 8

F OR T RAI NI N G P URP O S

E S

ONL Y

NORM

EXT POWER

OIL

APU STOP

START

IRS 1HRZNIRS 2 IRS 3

GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

GEN 3

GEN

G1

GENERATOR 1

G2

GENERATOR 2

TOAMMETER

G3

GENERATOR 3

START

CONTACTOR

TO AMMETER

G

START CONTACTOR

BATTERY 1

CONTACTOR

BATTERY 1

STARTCONTACTOR

GPUCONTACTOR

APU

TOAMMETER

TOAMMETER

FIREEXTINGUISHERS

COCKPITDOME LIGHT

TOAMMETER

GPU RECEPTACLE

BATTERYBUS

STARTCONTACTOR

BATTERY 2

DC SYSTEM

OVERHEAD PANEL

BATTERY 2CONTACTOR

STARTINGBUS

APU

A B

CONDITIONS: BUS TIE CLOSED. BOTH BATTERY SWITCHES ON. POWER SELECTOR

SWITCH IN NORMAL


Figure 2-18. Generato



R ev i s i o

n4

2 - 1 9

F OR T RAI NI N G P

URP O S E S

ONL Y

NORM

EXT POWER

OIL

APU STOP

START


GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

GEN 3

GEN

G1

GENERATOR 1

TOAMMETER

G2

GENERATOR 2

TOAMMETER

G3

GENERATOR 3

STARTCONTACTOR

TO AMMETER

G

START CONTACTOR

BATTERY 1CONTACTOR

BATTERY 1

START

CONTACTOR

GPU CONTACTOR

APU

TOAMMETER

TOAMMETER

FIRE

EXTINGUISHERS

COCKPITDOME LIGHT

TOAMMETER

GPU RECEPTACLE

BATTERYBUS

START

CONTACTOR

BATTERY 2

DC SYSTEM

OVERHEAD PANEL

BATTERY 2CONTACTOR

STARTINGBUS

APU

A B

CONDITIONS: GENERATORS 1, 2, AND 3 OPERATING. BOTH BATTERY SWITCHES ON. BUS TIE OPEN, POWER SELECTOR SWITCH IN NORMAL.


Figure 2-19. Normal F





GROUND POWER UNIT

General

An approved external power unit (GPU) (gen-erator or rectifier) may be connected to the stan-dard three-pin receptacle (Figure 2-20) locatedexternally on the aft right side of the airplane.The unit should be a constant-voltage unit capa-ble of negligible droop when operating at maxi-

mum load. The maximum rating should notexceed 1,200 amps, and the voltage should beregulated at 28 VDC.

Protection

Overvoltage protection is provided by a magnet-ic coil controlled by a PCB (printed circuitboard). When output voltage reaches approxi-mately 31 ±1 VDC volts, the PCB energizes themagnetic coil and disconnects the GPU from thedistribution system. A reset pushbutton isprovided at the main electrical box (Figure2-20); momentarily pushing this switch resetsthe GPU, provided that the fault no longer exists.

Control

The GPU is controlled by a two-position powerselector switch located on the overhead switchpanel (Figure 2-4). The switch positions arelabeled “NORMAL” and “EXT POWER.”When in the EXT POWER position, the bat-tery 1 contactor and battery 2 make-and-break switch open, isolating both batteries (BAT 1 andBAT 2 lights come on).

Figure 2-21 illustrates operation of the GPU.

The bus tie automatically closes (even if the bus-tie switch is off), and the BUS TIED light comeson. The generators, including the APU generator,are inhibited whenever external power is op-

switch, closing it an

to the left main busto buses A1 and A2to the right main busGPU may now be ugines, or it may bechecking the airplancondition prevails, 3 lights remain on.

The bus tie interconngrounded independeposition of the power

The battery bus isthrough the start bus.

DC POWERDISTRIBUT

GENERAL

The distribution busebreaker panels (Figurliner aft of the overhdivided among the le

Four feeder cables (proute power from therear compartment (Fcircuit-breaker panelsare routed along the buses A1 and A2 in and in the left sectbreaker panel. Simil

the right side of the fin the right circuit-brsection of the center c

The circuit breakers the associated circui





GPU RECEPTACLE POW

MAIN ELECTRICAL BOX

NO. 1 BATTERYMAKE-AND-BREAK

2 - 2

OVERHEAD PANEL



R ev i s i on4

22

F OR

T RAI NI N G P URP O S

E S

ONL Y

NORM

EXT POWER

OIL

APU STOP

START


GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

A B

GEN 3GEN

G1

GENERATOR 1

TOAMMETER

G2

GENERATOR 2

TO

AMMETER

G3

GENERATOR 3

STARTCONTACTOR

TO AMMETER

G

START CONTACTOR

BATTERY 1CONTACTOR

BATTERY 1

START

CONTACTOR

GPU CONTACTOR

APU

TOAMMETER

TOAMMETER

FIREEXTINGUISHERS

COCKPITDOME LIGHT

TO

AMMETER

GPU RECEPTACLE

BATTERYBUS

STARTCONTACTOR

BATTERY 2

DC SYSTEM

OVERHEAD PANEL

BATTERY 2CONTACTOR

STARTINGBUS

CONDITIONS: POWER SELECTOR SWITCH IN EXT POWER.


Figure 2-21. G




PRIORITY DISTRIBUTION

Certain circuits are given priority classificationand are normally powered from bus A1. If busA1 fails, these circuits are automatically pow-ered from bus B1. The following systems areincluded in this priority group:

• Warning lights

• Lights and engine indicators test

• Emergency lighting indicator light

• Takeoff warnings

• Fire panel

• Refueling

• Horizontal stabilizer position detector

• Slat monitoring

• Aileron and elevator Arthur monitoring

• Audio warnings

• Audio shutoff

• Baggage compartment door detector

• Center fuel tank monitoring

Figure 2-22 depicts the priority distributionsystem in its normal configuration. The operat-ing power supply is from A1. When power fails

on bus A1, the transfer relay is deenergized,and the circuits are then powered from bus B1.

AVIONIC

Selective powsystem, primground and tTwo latching 2-2 and 2-MASTER” anlocated, one and right circ

pushed in, tnecting powetem. A similoption provirespective ftrolled by twoand R/H AVI

2 - 2

FAILURE B



R ev i s i on4

24

F OR


E S

ONL Y

FAILURE B

WARN LIGHT B(R1WW)

28-VDCBUSB1

AUDIO WARN B(1WL2)

AUDIO WARN A(1WL1)

WARN LIGHT A(L1WW)

28-VDCBUSA1

FAILURE A

AUDIOWARNING

28-VDCPOWERSUPPLY

Figure 2-22. Pr



Table 2-1. ELECTRICAL POWER SOURCE

LEFT CIRCUIT-BREAKER PANEL

PRIMARY BUS A1


FIRE AUDIO WARN A Audio warning RADIO ATC1*WARNING

WARN LIGHTS A Warning panel VOR1* PANEL

DME1*

WARN LIGHTS A Light test ADF1* EX D/N

EXTING 1 Fire NAVIGA- DDRMTION

DETECT 1 Fire ADC1*

LIGHTS WARN SG1* A-B

EADI L

NAVIGA- IRS 1 BAT IRS 1 battery EHSI LTION EFIS C

IRS 1 IRS 1

TEMP PROBE Probe heating

HRZN ST BY Standbyhorizon




Table 2-1. ELECTRICAL POWER SOUR

LEFT CIRCUIT-BREAKER PANEL (C

PRIMARY BUS A2


FIRE EXTING 3 Fire RADIOWARNING

DETECT 3 Fire

BAG COMP Fire

BAT TEMP Battery NAVIGA-temperature TION

BLOWER LH Ventilation

TEST WARN A-B

NAVIGA- IRS 3 IRS 3TION

IRS 3 BAT (Option) IRS 3 battery

ANNUNC LH Radio navlighting

AFCS 1 AP Pilot FGC

AFCS 1 YD Pilot FGC

RADIO ICS LH Intercom




Table 2-1. ELECTRICAL POWER SOURCES (

CENTER CIRCUIT-BREAKER PANEL

PRIMARY BUS A1


LIGHTS ANTICOL FIN Anticollision HYDR L/G COMISC lights

LH EXT External STBY

LIGHT lighting

CKPT LH Lighting HYDR READING

NAV Navigationlights

STROBE Strobe ANTI- WSHLlights ICE LH

COND’G

CENTER Lighting LH PIT

INSTR LH Instrument LH STAlighting HEAT

INV (115V/60 Hz) Option COND

ENGINES IGNTR AUTO Starting CABIN

ENG FAIL 2 Takeoff LH AOwarning

N2 1 Indicators ENGINN1 ITT 2 Turbine temp

AIR FR

CMPTR Engine HP BLComputer

IGNTR 1 StartOIL 1 Engine FLT A/B CO

control CONTROL PITCH

or 115-VAC master






PRIMARY BUS A2


LIGHTS BELTS NO Passenger ANTI-MISC SMK’G signs ICE

COND’G

ENTRY Entrancelighting

SHIELD Glareshieldlighting

DRAIN HEAT Drainanti-icing

LANDING LH Lights

ENGINES N2 3 Indicators

CMPTR 3 Engine FLT

computer CONTROL

IGNTR 3 Starting

OIL 3 Enginecontrol

FUEL 2 Fire

SHUT OFF

FUEL STBY BOOST 2 Fuel

FUEL FLOW 3 Flowmeter





PRIMARY BUS B1


FLT STAB NORMAL Horiz stab ENGINES N2 2CONTROL

AIL FEEL Arthur N1 ITTmonitoring

RH AUTO SLAT Slats CMPT

IGNTR

ANTI- ENGINE 2 Anti-icing OIL 2ICECOND’G AFT SIDE Window CMPT

WINDOW STBY

WIPER RH Wiper LIGHTS LAV M

MISC

COND’G CABIN Conditioning OVERH

BOOTSTRAP Bootstrap FWD C INDI

RH AOA HEAT Probe heat RH CA REA

HYDR L/G INDIC Landing gear CKPT REA

HYDR 2 INDIC HydraulicTAXI

FUEL NORM BOOST 2 Fuel




PRIMARY BUS B1


FLT STAB NORMAL Horiz stab ENGINES N2 2CONTROL

AIL FEEL Arthur N1 ITTmonitoring

RH AUTO SLAT Slats CMPTR

IGNTR

ANTI- ENGINE 2 Anti-icing OIL 2

ICECOND’G AFT SIDE Window CMPTR WINDOW STBY P

WIPER RH Wiper LIGHTS LAV MA

MISC

COND’G CABIN Conditioning OVERH

BOOTSTRAP Bootstrap FWD C INDIR

RH AOA HEAT Probe heat RH CA READ

HYDR L/G INDIC Landing gear CKPT R READ

HYDR 2 INDIC HydraulicTAXI

FUEL NORM BOOST 2 Fuel







PRIMARY BUS B2


LIGHTS GALLEY MASTER Galley FUELMISC

LANDING RH Lights

ANTICOL BELLY Anticollisionlights

RH EXT LIGHT Right externallights

INSTR RH Instrumentlighting

PEDESTAL Instrument HYDRlighting

AFT CABIN CabinINDIRECT lighting

LH CABIN Reading ANTI-READING lights ICE

COND’G

ENGINES APU APU

N1 ITT 3 Turbinetemperature

FUEL 1 Fire

SHUT OFF

REVERSE Engine 2 CONTROL reverser

REVERSE WARN FLT

VALANCE OR





RIGHT CIRCUIT-BREAKER PANEL

PRIMARY BUS B1


FIRE AUDIO WARN B Audio warning RADIO VOR 2*WARNING

WARN LIGHT B Warning panel DME 2 PANEL

ADF 2*

WARN LIGHTS B Light testing EX D/N ATC 2*

VHF 3*

EXTING 2 Fire

DETECT 2 Fire NAVIGA- SG 2*TION

APU Fire EFIS C

EHSI R

EADI R

NAVIGA- IRS 2 BAT IRS 2 battery DDRMTION

IRS 2 IRS 2 ADC 2*





RIGHT CIRCUIT-BREAKER PANEL (C

PRIMARY BUS B2


MISC NOSE FAN Ventilation RADIO

BLOWER RH Ventilation

CREW SEATS Crew seats

EMERG LIGHTS Emergencylightbatteries

NAVIGA-TION

RADIO HF 2 PWR HF 2

PUBLIC Public

ADDRESS address

ICS RH Intercom

NAVIGA- AFCS 2 AP Copilot FGC

TION

AFCS 2 YD Copilot FGC




1. The majority of the DC electrical compo-nents are located in the:

A. Nose compartment

B. Cockpit headliner

C. Rear compartment

D. Underfloor area

2. The battery bus is a hot bus when:

A. The BAT 1 switch is on.

B. Either battery is installed and connected.

C. Any battery or generator switch is on.

D. The BAT 2 switch is on.

3. The DC voltmeters are directly con-nected to:

A. The battery shunts

B. A selector switch below each ammeter

C. Generator shunts

D. Their associated main bus

4. The bus-tie control power source is

supplied:A. From the hot battery bus

B. Directly from either battery

C. From the start bus

D. From the left or right main bus

5. The AVIONICS MASTER switches arelocated on the:

A. Main DC box

B. Center pedestal

C. Circuit-breaker panel

D. Pilot’s and copilot’s panels

6. The line disconthe engine-driv

A. In the nose

B. On the main

C. On the asso

D. In the forw

7. Ventilation foring flight is pro

A. A line supp

B. A bleed-air

C. A blower cgear weight

D. Low-pressur

outlet

8. The battery 2 c2 to the:

A. Right main

B. Battery bus

C. Start bus

D. Associated

9. If the batteries aand properly cnot connect to ciated switchescause is the:

A. Power selNORMAL.

B. Power selecPOWER.

C. External pois open.

D Bus tie swi

QUESTIONS



CHAPTER 3

LIGHTING

CONTENTS

INTRODUCTION ......................................................................

GENERAL..................................................................................

INTERIOR LIGHTING.............................................................

Cockpit Lighting .................................................................

Passenger Cabin Lighting ...................................................

Passenger Ordinance Signs .................................................

Baggage Compartment, Rear Compartment, and Nose Co

EXTERIOR LIGHTING ............................................................

Navigation Lights...............................................................

Logo Lights .........................................................................

Anticollision Lights............................................................

Strobe Lights......................................................................

Landing Lights ....................................................................

Taxi Light...........................................................................

Ice Detection Lights ............................................................

LIGHTING BUS DEPENDENCY............................................

EMERGENCY LIGHTING .......................................................




ILLUSTRATIONS

Figure Title

3-1 Cockpit Lighting Controls..........................................

3-2 Cockpit Dome Lights ..................................................

3-3 Cockpit Reading Lights..............................................

3-4 Glareshield Lighting...................................................

3-5 Circuit-Breaker Panel Lighting ..................................

3-6 BRIGHT–DIM Switch ................................................

3-7 Passenger Cabin Lighting Controls............................

3-8 ENTRANCE Pushbutton............................................

3-9 OCCUPIED Light .......................................................3-10 Passenger Lighting Control Switch............................

3-11 Typical Reading and Table Lamp Switch...................

3-12 Passenger Ordinance Signs .........................................

3-13 Nose Cone, Baggage Compartment, and Rear CompaLighting Components.................................................

3-14 Exterior Light Locations .............................................

3-15 Navigation Lights and NAV Switch...........................

3-16 Anticollision Strobe Lights and ANTICOL Switch ...

3-17 Strobe Lights ...............................................................

3-18 Landing Lights and LANDING Switch .....................

3-19 Taxi Light and TAXI Switch .......................................

3-20 Ice Detection Lights and Control Switch ...................




INTRODUCTION

The Falcon 900 lighting system is divided into interior lighting includes cockpit, passenger compartment, bagpartment, and nose cone lighting. Cockpit lighting incluspecific lighting for instruments and map reading. Pasprovides illumination for warning signs and specific arsafety and convenience. Individual lights are provided fgage compartment, and nose cone. Exterior lighting cotaxi, anticollision, wingtip strobe, and wing ice detecti




The passenger compartment lighting con-sists of fluorescent lighting controlled by

pushbuttons or toggle switches for the en-trance, lavatory, and both sides of the com-partment. Passenger reading lights a nd callsign lights are controlled by switchlights.

The nose cone, baggage, and aft compartmentshave individual lights.

The emergency lighting system ensures illu-

mination of the cockpit, emergency exit, andpassenger compartment entrance door in caseof total electrical power failure.

The airplane exter ior l ighting system isequipped with the following lights:

• Three navigation lights (red on the leftwingtip, green on the right wingtip,

and white on fairing)

• Two landing lwing/fuselag

• One taxi ligh

• Two red strobmounted on toand one on th

• Two strobe li

adjacent to th

• Two wing iceon either sidedetect ice acleading edges

• Two vertical stalled on theizontal stabil



source is the battery. Therefore, it is not af-fected by an electrical system power failure.The six remaining tubes are supplied with 28VDC from the normal electrical system via115-VAC, 50-Hz inverters.

A light located on the lavatory door (Figure3-9) indicates (when illuminated) that the toi-let is occupied.

The front lavatory lighting system is con-trolled by a microswitch acti vated by thedoor latch. The f luorescent tube is sup-plied with power via a 28-VDC-fed 115-VAC, 50-Hz inverter.

Passeng

This indirecered from anThe ceiling fluorescent with individthe cabin en

Passeng

The passenglights and tabpart of the dplied with 2table lamp nearby switc

Figure 3-8. ENTRANCE Pushbutton




PASSENGER ORDINANCE

SIGNSThe Fasten seat belts (Figure 3-12) instructionand the no smoking symbols appear on two lu-minous passenger ordinance signs in the pas-senger cabin and are visible from any sea t.

The fasten seat belts signs are controlled bythe FASTEN BELTS switchlight located on theINTERIOR LIGHTS panel of the cockpit over-head panel. The switchlight has a built-in bulbtest circuit.

The no smoking sign lights are controlled withthe switchlight engraved with the internation-al no smoking symbol. The switchlight is lo-cated on the overhead cockpit interior lightspanel and has a built-in bulb test circuit.

BAGGAGE CO

REAR COMPANOSE CONE L

This lighting consi

• An inspection

• A dome lightment

• Dome lights f

The system compon3-13.

These systems havethe cockpit dome lighby the batteries throuThe nose cone inspwith a built-in switchbaggage compartm

trolled with a micros

EXTERIOR

The exterior lightinanticollision, strobdetection lights. Th

are shown in Figure

NAVIGATION

There are three navia red light on the lefthe right wingtip, vertical stabilizer strolled with the NAV

TERIOR LIGHTS p

LOGO LIGHTS

Two white lights a

Figure 3-11. Typical Reading andTable Lamp Switch


g t in




Figure 3-13. Nose Cone, Baggage Compartment, andRear Compartment Lighting Components

Figure 3-15. Navigation Lights and NAV Switch

ANTICOLLISION LIGHTS

There are two red anticollision strobe lights

Each light has an intensity rating of 100 cadle power. The vertical fin and belly anticosion lights are supplied by power supply box





There are two red anticollision strobe lights(Figure 3-16): one centered on the fin fairingand the other on the underside of the fuselage.

sion lights are supplied by power supply boxthat deliver pulsating high-voltage current. Ttwo power supply boxes are synchronized

that the two lights function simultaneously


STROBE LIGHT

Figure 3-16. Anticollision Strobe Lights and ANTICOL Switch

Figure 3-17. Strobe Lights

EMERGENCY LIGHT

I f l l i l


g t in



ICE DETECTION LIGHTS

Two 75-watt white lights (Figure 3-20) arelocated on either side of the fuselage forwardsection. They enable in-flight detection of iceaccumulation on the wing leading edges. The

lights are controlled with the WING switch lo-cated on the EXTERIOR LIGHTS panel.

LIGHTING BUSDEPENDENCY

Table 3-1 lists the lights and buses that sup-ply the electrical powe r.

In case of total electrical power suure, the emergency lighting systemillumination of the cockpit, emergand main entrance door. This systeplied by three power supply boxes, ta in ing a three-element n ickel-battery normally charged by the onbtrical system. In case of electrical pure , these bat ter ies provide l igapproximately ten minutes. An emlighting power supply assembly isFigure 3-21.

The power supply located behind

furnishes power to the:

• Pilot dome light

• Passenger door EXIT sign

• Main entrance door spotlight


Figure 3-20. Ice Detection Ligand Control Swi

Figure 3-19. Taxi Light and TAXI Switch




QUESTIONS



1. The cockpit dome lights are turned onby:

A. A switch on the overhead panel or aswitch at the passenger entrance

B. A switch at the passenger entrance

C. A switch on the overhead panel

D. A rheostat on the copilot’s side con-sole

2. The emergency lighting switch posi-tions are:

A. OFF, ON, ARMED

B. OFF, STANDBY

C. OFF, CHARGE

D. OFF, ARMED, STANDBY

3. After a to ta l electrical failure, the emer-gency lights have battery power for ap-proximately:

A. 5 to 10 minutes

B. 10 to 20 minutes

C. 25 to 35 minutes

D. 40 to 45 minutes

4. The emluminatbuses ising swi

A. ON

B. OFF

C. ACT

D. ARM

5. The winwith the

A. Ant

B. Lighthe

C. Nav

D. Stro

6. The batthe:

A. Circ

B. Eme

C. Dom

D. Rea

QUESTIONS

CHAPTER 4




CHAPTER 4

MASTER WARNING SY

CONTENTS

INTRODUCTION ......................................................................

GENERAL..................................................................................

WARNING PANEL....................................................................

FIRE PANEL ..............................................................................

HYDRAULIC CONTROL AND INDICATOR PANEL ..........

BATTERY TEMPERATURE INDICATOR .............................MISCELLANEOUS VISUAL WARNINGS ............................

ENG 2 FAIL Warning Light ...............................................

Interstage Turbine Temperature Indicators ........................

AUDIO WARNINGS .................................................................

General ................................................................................

Priority Warnings ................................................................

Warning Voice.....................................................................

INSTRUMENT PANEL INDICATOR LIGHTS ......................

Configuration Panel and Landing Gear Control Handle.....Thrust Reverser Indicator Lights ........................................

OVERHEAD PANEL INDICATOR LIGHTS...........................

ILLUSTRATIONS




ILLUSTRATIONS

Figure Title

4-1 Warning Panel .............................................................

4-2 Fire Panel ....................................................................

4-3 Hydraulic Control and Indicator Panel .......................

4-4 Battery Temperature Indicator <172 ...........................

4-4A Battery Temperature Indicator ≥172 ...........................

4-5 ENG 2 FAIL Light .......................................................

4-6 ITT Indicators...............................................................

4-7 HORN SIL Pushbutton ...............................................

4-8 Configuration Panel and Landing Gear Control Handl

4-9 Thrust Reverser Indicator Lights ..................................

4-10 Overhead Panel Lights.................................................

TABLES

Table Title

4-1 Annunciator Illumination Causes.................................

4-2 Fire Panel Illumination Causes ...................................

4-3 Hydraulic Control and Indicator Panel Illumination Ca

4-4 Battery Temperature Indicator Illumination Causes ...

4-5 ENG 2 FAIL Illumination Causes ...............................

4-6 ITT Light Illumination Causes.....................................

4-7 Audio Warning Causes.................................................

4-8 Audio Warning Testing ...............................................




INTRODUCTION

The master warning system on the Falcon 900 provides a wmalfunctions and unsafe operating conditions which requindication that a particular system is in operation. A systto draw attention to certain system situations.

The warning system makes possible the pre-sentation and testing of warning and indica-

A panel within the center

TEST

MASTER WARN

GENERAL

WARNING PANEL A TEST switch is pr




WARNING PANEL

The warning panel (Figure 4-1 and AppendixB) provides a means of alerting the pilot tocertain system conditions or malfunctions. Theannunciators are either red or amber and illu-minate for the causes lis ted in Table 4-1. Eachannunciator contains two bulbs and an engravedlegend identifying the corresponding system.

nunciator lights and

Electric power is noA1 bus through cirLIGHTS A. In the evtem failure, power isto the B1 bus and fEX WARN LIGHswitch is provided toators during night f

NOTE:ON AIRCRAFT WITHOUT TRANSFER VA

PANEL DOES NOT FEATURE XTK2 OPE

*


Table 4-1 ANNUNCIATOR ILLUMINATIO



Table 4 1. ANNUNCIATOR ILLUMINATIO

• The control switch of theindicated computer is in the OFFor MAN position.

• The indicated computer hasfailed (electrical power supplyfailure, internal failure, orincorrect data).

• The light comes on if:(Specific light to aircraftincorporating M880Amodification).

• The main entrance door is notlocked or the front lavatorycompartment service door isnot locked (on aircraft equippedwith this lavatory compartment).

• The indicated heating systemsare not on.

• Angle-of-attack heating hasfailed.

• The oil pressure of the indicatedengine is lower than 25 psi(1.72 bar).

• Chips are detected in theindicated engine oil system.

Red light with M880A

• The indicated heating systemsare not on.

• Pitot or static pressure probeheating has failed.

• The indicated heating system isnot on.

• Standby pitot pressure probeheating has failed.

Annunciator Cause for Illumination

CMPTR 1

Annunciator

BUS TIED

HOTBAT

FWDDOORS

AUTOSLATS

CMPTR 2

CMPTR 3

OIL 1

OIL 2

OIL 3

L. AOA

R. AOA

L. PITOT

R. PITOT

ST BYPITOT

BAT 1

BAT 2

L. WHLOVHT

R. WHLOVHT


Table 4-1 ANNUNCIATOR ILLUMINATION CAUSE



Table 4 1. ANNUNCIATOR ILLUMINATION CAUSE

The pressure switch located on theindicated engine fuel supply LP lineindicates a pressure drop (pressureless than or equal to 4.6 psi[approximately 320 mb]).

The front-to-rear tank transfer valveis open when it should be closed.

The front-to-rear tank transfer valveis closed when it should be open.

The cabin baggage compartment

access door is not closed.

A fuel level below 200 pounds isdetected in tank group G1 or G3.

• For airplanes SNs 1 to 11—Afuel level below 200 pounds isdetected in tank group G2 (or below 1,100 lb if booster pumps2 are off).

• For airplanes SNs 12 and sub-sequent—A fuel level below 200pounds is detected in tankgroup G2.

The emergency aileron actuator isnot in the neutral position.

A discrepancy is detected betweenthe IAS output of the air datacomputer and the positioninformation supplied by the linearpotentiometer on the aileron Arthur


LOFUEL 1

The ligThe foframethe “ Nsign is

The thare nopositio

The ligduringphase

• Onnot

• Thnot

• Thdoo

• Thacc

• This s

• Bu

• Thto O

• Thrais

• Thaut

On airwhen warnin

The Afailed.

Annunciator Caus

AP

MISTRIM

LOFUEL 3

LO

FUEL 2

AILZERO

AILFEEL

FUELING

REVUNLOCK

AFT CABINISOL

XTK 2OPEN

XTK 2CLOSED

BAG

ACCESS

FUEL 1

FUEL 2

FUEL 3



(option 25-21-01)



MISCELLANEOUS INTERSTTEMPER




VISUAL WARNINGSENG 2 FAIL WARNING LIGHT

THE ENG 2 FAIL warning light (Figure 4-5 andAppendix B) illuminates steady. The light andcauses for illumination are shown in Table 4-5.

TEMPER

The three ITAppendix Bcular dial, aengine overand their cain Table 4-6

The lights can

switch locate

Fig

Annunciato

Table 4-6

Figure 4-5. ENG 2 FAIL Light

• The airplane is on the ground and

the No. 2 engine power lever is atthe 84° setting and the No. 2 en-gine power is less than 85% N1.

• The No. 2 engine S-duct accessdoor is not properly closed.


ENG 2 FAIL

Table 4-5. ENG 2 FAILILLUMINATION CAUSES

AUDIO WARNINGS sound can be silenc




GENERAL

The audio warning system alerts the crew to con-figuration anomalies or certain operational con-ditions. The unit is located in the pedestal andissues warning sounds through the cockpit loud-speaker or the pilot and copilot headsets.

The sounds are generated by a transistorizedaudio warning uni t which includes fourpotentiometers. The potentiometers are des-ignated for:

• Loudspeaker output

• Pilot headset audio output

• Copilot headset audio output

• Rattle noise output

In some cases, the warning sound or voicecan be silenced by pressing the HORN SILpushbut ton (Figure 4-7) located on thepedestal.

The warnings, descriptions of sound, associ-ated warnings, causes of warning, and if the

PRIORITY WAThe stall and VMO /Mity over all others. simultaneously witthe exception of thmovement warning

Next in priority is t

be triggered simultaings, with the excepbilizer-in-movemen

WARNING VO

Single Warnin

A single voice advi

seconds until the copears or the HORN




Table 4-7. AUDIO WARNING CAUSES (Con



SIMULTANEOUSWARNING TYPE OF WARNING OR

SOUND INDICATION

Landing gear Warning voice GEAR Red light on landing The contgear control handle gear dowflashes. locked p

lower thaleast onelevers is power po

STOP onand at legears is

The contdownlocposition,extendedleast oneis not do

Decision height Warning voice Letters DH appear on Preset dMINIMUM both EADIs. reached.

Autopilot Warning voice Illumination of the AP Failure o(Aircraft with AUTOPILOT light on the warning of the auM880C) panel

Takeoff Warning voice Illumination of the T/O The airpConfiguration NO TAKE-OFF CONFIG light on the ground, a

the fire panel power lebeyond 8the follow

• Flapsor mo

• Flap/sCLEA

• Airbraretrac

• Horizout of30' po

• Slats

• Park and d

Multiple WarningsA maximum of three voices can be heard at

B) red light aconfiguratio




A maximum of three voices can be heard atone time. If a fourth signal is received, it isstored until one of the three active causeshas terminated.

If the HORN SIL pushbutton is pressed, onlyone voice at a time disappears.

The audio warnings are tested by the methodslisted in Table 4-8.

INSTRUMENT PANELINDICATOR LIGHTS

CONFIGURATION PANELAND LANDING GEARCONTROL HANDLE

The configuration panel lights and the landinggear control handle (Figure 4-8 and Appendix

The warningbe used to dthe lights an

Table 4-8. AUDIO WARNING TESTING

VMO /MMO Depressing ADC 1 or ADC 2pushbuttons on the pedestal

Cabin altitude Depressing the test push-button on the cabin pressurecontroller accompanied byillumination of the CABIN lighton the warning panel

Fire Setting the warning panelLIGHTS–TEST–FIRE switchto FIRE (all the FIRE lights onthe fire panel illuminate)

Stall On the ground depressing theSTALL 1 or STALL 2pushbutton

Horizontal Action on the pitch trimstabilizer in control


THRUST REVINDICATOR LI


Table 4-9. CONFIGURATION PANELAND LANDING GEAR



The thrust reverser 9 and the Appendix center instrument pcauses for illuminat10.

Figure 4-9.

The cmove

Annunciator Caus

TRANSIT

Table 4-10. TIN

IL

CONTROL HANDLEILLUMINATION CAUSES

• The control handle is in the gear-up position, and the three gearsare not yet uplocked.

• The control handle is in the downposition, and the landing gear isnot fully downlocked.

• The control handle is in the upposition, speed is lower than 160knots, at least one of the powerlevers is at a reduced setting, butnot on STOP for aircraftincorporating M881, and at leastone of the three gears is notdownlocked.

• Steady illuminationAt least one of the six airbrakes isnot in the retracted position.

• Flashing

After automatic retraction of theairbrakes, as long as the handleis in the extended position whilethe airbrakes are retracted

• Steady illumination

All the slats are extended.

• Flashing

Only the outboard slats areextended.

During any movement of the slats orif any one of the slats has failed toeither extend or retract

The indicated landing gear isdownlocked.

• Main gear:


AIRBRAKE

SLATS

LH NOSE RH

MOVING

OVERHEAD PANEL The BRIGHto DIM to re




INDICATOR LIGHTSThe overhead panel lights (Figure 4-10 and Ap-pendix B) are tested by setting the warningpanel TEST switch to LIGHTS.

panel lights

The overheafor illumina


Table 4-11. OVERHEAD PANEL LIGHT ILLUMINATIO



• Steady illumination: It isdepressed to on.

• Flashing: In the event ofautomatic shutdown of the APUby flight/ground contact, ECUoverheat, faulty generator

regulation, or starting faults

The light illumination indicates lowoil pressure or high oil temperature.

The APU generator is off the line.

The APU generator has excitation

voltage. The APU generator switchis on.

The igniter unit of the indicatedengine is energized.

The side tank group interconnection

valve is not closed.

One of the 1–3 or 3–1 crossfeedvalves is not closed.

The fuel level in the associated tank(left or right centerwing tank) islower than 1,000 pounds or the fuellevel in the rear tank is lower than1,100 pounds.

One of the associated crossfeedvalves (1–2 or 3–2) is not closed.

The bleed-air isolation valve isl d


MASTER

OIL

GEN

IGN 1 IGN 2

XFR

XTK

X.BP

LEVEL

X.BP

ISOL

IGN 3

• Gror pocoarpran

No

• AmENonis sy

• AmOvNo

ovis

• Grsecoar

• Amswprthe

• AmOvthe

On athe illgreensame

WING

Eithecircuiand r

Annunciator Caus

APU

APU

APU PANEL

DC SYSTEM PANEL

ENGINES PANEL

FUEL SYSTEM PANEL

BLEED-AIR PANEL

ENG 1 ENG 2

WING

ENG 3

(3 AMBER LIGHTS)

(CENTER AMBERLIGHT)

(2 LH AND RH

AMBER LIGHTS)


Table 4-11. OVERHEAD PANEL LIGHT ILLUM



OPERATION

When airplane power is turned on, the master

warning and annunciator lights circuit is ener-gized with 28 volts. Illumination of one of the annunciators occurs when its correspond-ing system experiences a failure, is turnedoff, or has not been actuated. The annuncia-tor extinguishes only when the illuminatingmalfunction and/or circumstance has beencleared.

Intensity of most annunciators may be de-creased by means of the BRIGHT–DIM switch.When placed in the DIM position, each an-nunciator’s intensity is reduced by approxi-mately half. To return the annunciators tonormal intensity, the switch must be placed tothe BRIGHT position.

NOTEPower for the master warning panelis normally supplied from bus A1. If bus A1 fails, bus B1 automaticallypowers the warning panel

The FASTEN BELTS passengerordinance sign is illuminated.

The EMERG LIGHTS selector is inthe ON or OFF position, and theaircraft is electrically powered.

The no smoking ordinance sign isilluminated.


FASTENBELTS

EMERGLIGHTS

Annunciato

INTERIOR LIGHTS PANEL

OFF

LHAVIONICS

MASTER

OFF

RHAVIONI

MASTE


QUESTIONS



1. The warning panel can be checked by po-sitioning the control switch to:

A. BRIGHT

B. LIGHTS

C. FIRE

D. DIM

2. When the warning panel CABIN light il-

luminates, the accompanying sound is:

A. The word “cabin”

B. A continuous variable 660–3,330-Hztone

C. An intermittent 1,660-Hz tone

D. A high-pitched siren

3. The audible word “minimum” indicatesthat:

A. Fuel level is low.

B. Cabin pressure is low.

C. Preset decision height is reached.

D. Angle of attack is negative.

4. The HORN SILlence the audib

A. Fire

B. VMO /MMOC. Stall

D. Altitude de

5. The TEST pus

temperature in

A. Warning pa

B. Overhead p

C. Copilot’s s

D. Indicator

6. When more thative, they may

A. Be silenced

B. Not be sile

C. Be silenced

D. Be silencecause

CHAPTER 5




FUEL SYSTEM

CONTENTS

INTRODUCTION ......................................................................

GENERAL..................................................................................

FUEL STORAGE......................................................................

General................................................................................

Wing and Center Wing Tanks............................................

Front and Rear Fuselage Tanks..........................................

FUEL SYSTEM CONTROLS ...................................................

FUEL DISTRIBUTION .............................................................

General................................................................................

Fuel Tank Pressurization System.......................................

Fuel Transfer System ..........................................................

Engine Feed System ...........................................................

Crossfeed System ...............................................................

FUEL INDICATING SYSTEMS...............................................

Fuel Quantity Indicators .....................................................

Fuel Flow/Fuel Used Indicators .........................................

Fuel Temperature Indicator—Aircraft Lower Than SN 70 or without Option 28-40-01

Gravity Refueling ..........................................................................

D f li




Defueling .......................................................................................Drain Valves ..................................................................................

Anti-icing Blending Apparatus......................................................

LIMITATIONS......................................................................................

Fuel Quantity.................................................................................

Fuel Specifications ........................................................................

Pressure Refueling.........................................................................

Fuel Additives ...............................................................................

QUESTIONS.........................................................................................

ILLUSTRATIONS

Fig re Title




Figure Title5-1 Fuel Tanks ...................................................................

5-2 Fuel System Controls and Indications........................

5-3 Tank Pressurization and Quantity Indication .............

5-4 Fuel Distribution.........................................................

5-5 Transfer Valve Control Logic—Aircraft without SB F

5-5A Manual Control of Transfer Valve—Aircraft without

5-6 Jet Pump (Typical)......................................................

5-7 Crossfeed X-BP 1↔ 3—Normal Configuration...

5-8 Crossfeed X-BP 1↔

3—Pump 1 Inoperative.......

5-9 Crossfeed X-BP 1↔ 3—Pump 3 Inoperative.......

5-10 Crossfeed X-BP 1↔ 2 and 3↔ 2—Normal C

5-11 Crossfeed X-BP 1↔ 2 and 3↔ 2—Pumps 1

5-12 Crossfeed X-BP 1↔ 2 and 3↔ 2—Pump 2 In

5-13 Fuel Quantity Indicator (Typical)................................

5-14 Fuel Flow/Fuel Used Indicator (Typical)....................

5-15 Refueling System Controls and Indicators..................

5-16 Pressure Refueling ......................................................

5-17 Fuel Filler Port ............................................................

5-18 Gravity Refueling........................................................

5-19 Fuel Sump Drain Valve ...............................................

TABLES

Table Title




Table Title5-1 Fuel Tank Capacities ..............................................................

5-2 Fuel Specifications.................................................................


F



INTRODUCTIONDuring normal operations, the Falcon 900 fuel system csystems: the left, center, and right. Each subsystem normallengine. Interconnect and crossfeed valves allow fuel tranfuel feed from any tank in the event of fuel imbalance oplane may be pressure- or gravity-refueled or defueled.

Fuel storage consists of three tank groups: theleft wing and center wing, the forward andrear fuselage and the right wing and center

tank-to-tankfor fuel transa DC opera

F

0

2

4 6

8

10

MAINFUEL

LBS X 100

FUEL FLO W

GENERAL

Throughout this chapter a color code is used todesignate tank group fuel or tank group func-tion. Amber designates group 1, green desig-nates group 2 and yellow is used for group 3

compartments with fflow inboard and intpump compartmenttions of each wing i




TANK GROUP LITERS

KILOGRAMS(SPECIFICGRAVITY =

0.803)

U.SGALLO

Group

Left wing and center wing

Group

Right wing and center wing

Group

Front and rear tanks

2,756.5

2,756.5

1,655.0

1,525.0

907

907

544

501

3,433.0

3,433.0

2,061.0

1,899.0

Front

Rear

1

3

2

nates group 2, and yellow is used for group 3.

FUEL STORAGE

GENERAL

Three fuel tank groups comprise the Falcon

900 fuel storage. Group 1 consists of the leftwing and the left center wing tanks. Group 2consists of the forward and rear fuselage tanks.Group 3 consists of the right wing and rightcenter wing sections. All tanks a re part of theairplane structure . Table 5-1 lists the fuel tank capacities.

WING AND CENTERWING TANKS

Group 1 and group 3 integral tanks normally sup-ply the No. 1 and No. 3 engines, respectively.The tanks (Figure 5-1) are d ivided in to

tions of each wing icenter wing sectionsflow fuel from the N

The tanks are confthe front and rear witween the center wiend rib. Each wingtions, outboard and

rated by a rib with iflapper valves. Theand bottom surfaces3 tanks. As shownpump compartmentof each center wing one electrical, submpump for its respecboard end o f eacpressure-relief val

tank pressures do npressure. The interare covered with a sbacterial growth anright wing has sumfiller port, covered

Table 5-1. FUEL TANK CAPACITIES

R ev i s i on 3 . 0

1

SIDECENTER WING

TANK NO. 1 PUMP COMPARBAFFLE



5 - 3


P O S E S

ONL Y

SEALEDWALL

NO. 3 BOOST PUMPCOMPARTMENT

PRESSURIZATION

INTERCONNECTION

FRONTSPAR

FRONTTANK

CENTER WINGFEEDER

JUNCTIONPLATE

FLAPPERVALVES

FILLERPORT

PUREARSPAR

BAFFLE

NO. 3 PUMPCOMPARTMENT

BAFFLE

Figure 5-1. F

FRONT AND REAR FUSELAGETANKS

The front and rear fuselage tanks comprise

tion, only tenergized. 2↔3 crby selectio




The front and rear fuselage tanks comprisethe group 2 tanks, which normally supply theNo. 2 engine and auxiliary power unit (APU).These tanks form a structural part of the fuse-lage. As shown in Figure 5-1, the front tank is located immediately forward of the wingcenter section. The rear tank is aft of the mainlanding gear wheel wells. Two electric, sub-

merged, centrifugal boost pumps are locatedin the forward compartment of the rear tank.

The front and rear tanks are interconnected bytwo pipes which allow air and, if necessary,fuel circulation between the tanks . Two fueltransfer pipes also connect the two tanks.Sump drains are provided. No gravity fillerport is provided for the group 2 tanks; when

the airplane is gravity-refueled, the transfersystem must be used to fill the tanks.

FUEL SYSTEMCONTROLS

Fuel system controls are located on the

overhead panel and the center instrument panel.The overhead panel is arranged schematically,representing the basic fuel system configuration.Controls located on the overhead panel include:

• Two BOOSTER switches (one for eachwing tank)—These are two-positiontoggle switches which are used to turnthe group 1 or group 3 boost pump on oroff.

• A single BOOSTER switch (for group2)—This is a three-position switch withpositions marked “OFF–ST-BY–NORM.”It controls the two group 2 boost pumps

yfeed switcswitch is inpump will tin conjunct

• Three rotary clabeled “X-Bcrossfeed from

gine when cir

• A three-positnect) toggle swright positionsolenoid valvpressure to thtion valve be3 tanks. This

the lef t or rthrough a logthe group 1 anvalves, allowthe group 1 switch and Xo n l y w h e n1↔3 inte

• A switch iden

equipped witXTK 2 (aircrThis three-p“CLOSED–Oi t ive manuCLOSED) or front/rear tanfer solenoid)

Controls located o

panel include:

• Three guardedswitches—Whcloses the res




XTK

X.BP

XTK 2*

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

AUTO

OPEN

CLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C

DISCH

FAULT

FIR

2

1

0

FUEL 1 FUEL 2 FUEL 3XTK 2OPEN

XTK 2CLOSED

BAGACCESS

BAG 1 COMP 1

BLEED AIR

STABILIZ

RELAY TE

BOOSTERPUMP 1SWITCH

BOOSTERPUMP 2SWITCH

FUELTEMPERATURE

INDICATOR(OPTION)

GROUP1 3

INTERCONNECTIONSWITCH

X-BP2 3

CROSSFEEDROTARYSWITCH

X-BP1 2

CROSSFEEDROTARYSWITCH

BOOSTERPUMP 3SWITCH

X-BP 1 3ROTARYINTERCONNECT

SWITCH

FORWARD/REAR TANKTRANSFER SWITCH

*

FUE

*



* *

A main tenance pane l i s loca ted on thecopilot’s right side wall. Figure 5-2 shows thelocations of the fuel system controls and theannunciator indications, which are discussedl i hi h

• A water t rap

• A pressure ga

During refueling, ve




later in this chapter.

FUEL DISTRIBUTION

GENERAL

Distribution of fuel from the wing and fuselagetanks is aided by the tank air pressurization sys-tem. Fuel transfer by jet pump or via gravityis possible from the front fuselage tank to therear tank. Transfer is also possible betweenany two lateral tank groups for fuel leveling,if necessary. Jet pumps transfer fuel internallyin all tank groups. Boost pumps are the primarymeans of supplying fuel to the engines.

For operators with an approved MinimumEquipment List, dispatch with a boost pumpinoperative is possible, provided that the wingtanks are pressurized. An air pressure gage islocated in the rear equipment compartment toconfirm tank pressurization.

FUEL TANKPRESSURIZATION SYSTEM

The fuel tanks are automatically pressurized bylow-pressure bleed air from No. 1 and No. 2 en-gines as soon as either engine is started (Figure5-3). The system is completely automatic. Thepressure is regulated by the following:

• A pressure-reducing valve downstreamfrom the engine bleed

• Two valve boxes which include pres-sure-relief and negative pressure-relief valves

trically controlled the wing roots.

FUEL TRANSF

Front-to-Rear Aircraft FittedTransfer Valvewithout SB F9

Description

In addition to the twinterconnect pipesrear fuselage tanks, installed (Figure 5

sitioned for gravityto rear only until ththird of its capacityfor gravity transfernormally using a jrear tank. This pipefrom the bottom ofvalve in this pipe is ocontrol circuit to

sumption of fuel fvalve closes to constank capacity whiltank are consumed

The logic control c

• A fuel level dtank at a level

higher than thfuel flow cuto

• Two level detank at the 1,1


g t in



Revision 3.01 FOR TRAINING PURPOSES ONLY

XTK

X.BP

XT

X.BP

ENG 1 ENG 2

FUEL SYSTEM

BOOSTER

BOOSTER

NORM

ST-BY

OFF

°C

FUEL 1

LOFUEL 1

BLEEDOVHT

ECUOVHT

COND GOVHT

LOFUEL 2

REVUNLOCK

LOFUEL 3

FUEL 2

-2P BK CABIN

FUEL 3

FUELING

LEVEL

LEVEL

G1

G2

G3

G2

LEVEL LESS THAN 200 LB

LEVEL LESTHAN 1,000

CENTER WINGTANK

PROBE

SUMP DRAIN VENT VALVE

FRONT TANK

GRAVITY FUE

FLVA

WING NEGATIVEPRESSURE RELIEF

VALVE

VALVE BOXES

AUTOMATICDRAIN

LEGEND

TANK PRESSURIZATION

ELECTRICAL CONNECTION

REAR TANK

PRESSUREGAGE

DRAINBOWL

PRESSURE REDUCER

AUTOMATIC DRAIN

NO. 2 ENGINE LP AIR

NO. 1 ENGINE L

* NOT FEATURED ON AIRCRAFT WITHOUT


≥ 159

Figure 5-3. Tank Pressurization and Quantity Indication



FILLERPORT

CENTERWING TANKS

PROBE




;

;

G1

BOOSPUMP

NO. 1 ENGINE

NEGATIVEPRESSURERELIEFVALVE

INTERCV

(NORMA

R1 CROSSFEEDVALVE


MANIFOLD


GROUP 1CROSSFEED

MANIFOLD

GROUP 2FUEL

MANIFOLD

; ; ; ;

; ; ; ;

FUEL

SHUTOFFVALVE

; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;y y y y

;

;

y

y

;

; ;

;

;

;

;

;

; ;

;

;

;

The transfer valve is open when:

• The front tank contains

more than 1,200 pounds.

Figure 5-5 illvalve control

Operation


Frontk



or

• The rear tank containsless than 1,100 pounds.

The transfer valve is closed when:

• The front tank contains

less than 1,200 pounds.

or

• The rear tank containsmore than 1,400 pounds.

The transferfront tank legardless of ttransfer valvless than 1,11,400 poundfer from the f

ward shift of

In the event tion malfunccauses the X

1

OPEN

FRONT TANK REAR TANK

ANY TANKLEVEL1,200

LB

TRANSIT B

FRONT TANK

1,200LB

2

CLOSED


1,200LB

FRONT TANK

VALVPOSIT

1

2

OPE

CLOS

1,200LB

1,400OR MORELB

LESS THAN1,100 LB

3

OPEN


1,200LB

tank consumption

Reartank

consumption

warning panel light to illuminate. When illu-minated, the lights have the following mean-ings:

• XTK 2 CLOSED—This indicates thath f l i l d h i

pet and padding andtape covering the haexplains operation.

The following two c




the transfer valve is closed when itshould be open.

• XTK 2 OPEN—This indicates that thetransfer valve is open when it should beclosed.

The CLOSED–OPEN–AUTO switch on the

overhead panel can directly control openingand closing of the valve, overriding the con-trol logic. The XTK lights will indicate thecontrol switch positions until the valve posi-tion agrees with the switch position.

The transfer valve has a clutch-release actu-ator installed which can be used to open thevalve manually from inside the pressurized

cabin. The control is built into the cabin floor,even with the left No. 11 window. It consistsof a retractable handle (see Figure 5-5A) towhich access is gained by lifting up the car-

The following two cuse of the manual transfer valve:

• The valve hast ion, and thtrapped in the

• The valve has

and the rear cbe affected.

In the second climit is reachebalance is greand if the rema

each tank grouand 1,500 pounhowever, use ocan be avoide

CONTROL

LEVER

consumed first.

The following applies to aircraft not fittedwith the XTK 2 fuel transfer valve (aircraft




with SB F900-48):

The fuel flow, provided by gravity trans-fer from the front to rear fuselage tanksassisted by the jet pump in the rear tank,is in most cases higher than the amountof fuel consumed from the rear tank (supplied to engine 2).

This results in the following consump-tion sequence:

• Consumption of fuel in forward tank until empty, then,

• Consumption of fuel in rear tank

This consumption sequence leads to a

variation in the CG of the aircraft, whichhas to be compensated by an aft CG lim-itation for the zero fuel weight.

Internal Tank Transfer

Internal tank transfer is accomplished by jetpumps using motive flow fuel from the respect-ive boost pumps. Each wing tank group usesfour jet pumps; the fuselage tanks use two jetpumps for internal transfer in addition to a third jet pump used for fuel transfer from front tank to rear tank. Refer to Figure 5-4 and Engine FeedSystem, this chapter, for additional information.

ENGINE FEED SYSTEM

GeneralThe group 1 and group 3 fuel distribution(feed) systems each include the followingcomponents:

The group 2 distribution system includes thefollowing components:

• Two boost pumps• One fuel manifold

• Bus A1—Gro

• Bus A2—Gro

• Bus B1—Gro

• Bus B2 Gro




• One boost crossfeed manifold

• Three jet pumps

• One indicating pressure switch

Boost Pumps

Four identical, three-phase, AC-powered,submerged, centrifugal fuel boost pumps areinstalled in the Falcon 900 fuel system. Eachpump has a built-in inverter which converts28 VDC to 115 VAC 400 Hz. Group 1 andgroup 3 tanks each have one pump installed;group 2 tanks have two pumps. Group 1 andgroup 3 pumps deliver fuel under low pres-sure to their respective interconnection and

crossfeed manifolds and then to their re-spective engines. Group 2 pumps deliver fuelto the group 2 manifold and crossfeed man-ifold, then to the No. 2 engine. Pump dis-charge is 7.25 psi with a flow rate of 10.56 U.S.gallons per minute at 51,000 feet altitude.

Control switches for the boost pumps are lo-cated on the overhead panel. Electrical power

is supplied to the fuel boost pumps from thefollowing primary buses:

• Bus B2—Gro

Refer to Fuel Systefor additional infor

LEGEND

HIGH VOLUME,LOWER PRESSURE

THAN MOTIVE-FLOW FUEL

Jet PumpsFuel system jet pumps use motive-flow fuel

(see Figure 5-4) delivered from the fuel boostpumps. The jet pumps are identical in the threet k d t th t i

• Fuel sis shutIt is c

OFF swter inst




tank groups and operate on the ventur iprinciple (Figure 5-6).

In each tank group 1 and 3, four jet pumps si-phon fuel from the outboard, inboard, andcenter wing tanks and deliver it to the re-spective boost pump compartment. In tank

group 2, three jet pumps are installed in therear tank. One pump is used to siphon fuelfrom the front tank when the transfer valve isopen; the other two pumps siphon fuel fromthe aft part of the rear tank and deliver it to theboost pump compartment. The jet pumps en-sure that the boost pumps stay submerged inthe boost pump compartments. The jet pumpcircuit is equipped with a filter and a check

valve with a clogging indicator for systemprotection (clogging indicators are only onaircraft prior to SN 96).

Groups 1 and 3Interconnection ManifoldsThe groups 1 and 3 interconnection manifoldsare mounted in a dry bay outside the centerwing tank on the rear spar. They receive fuelfrom the respective boost pumps. An electricallydriven spherical core valve (R1 and R3 cross-feed interconnect valves on Figure 5-4) on eachmanifold supplies motive-flow fuel to the jetpumps in normal operation. Both valves arecontrolled by an X-BP interconnect rotaryswitch and are normally positioned closed;when selected to open, fuel is directed to theX-BP crossfeed interconnect system. Includedin each manifold is a filter for the motive-flowfuel and a clogging bypass valve with an indi-cator. An outlet enables a direct connection

motor

• A n e lvalve—requir1↔switch

• A man(normthe nofrom trespecvalves

Group 2

The manifoland connect

• The grgine su

• Group

In addition, 1 and 3 cros

• GroupNo. 3 e

• Groupengine

Group 2 The group 2

the forward bthree valves

• Fuel shshut ofIt is c


XTK

X.BPLEVEL LEVEL

BOOSTER BOOSTER



(normal/drain) drain valve—When inthe normal position, the fuel feed linefrom the boost pumps is connected to theNo. 2 engine via the fuel shutoff valve.

Indicating Pressure SwitchesLocated in the nacelles, each engine feed sys-tem has a pressure switch installed which il-luminates annunciator lights FUEL 1, FUEL

2, and FUEL 3 when the corresponding fuelpressure is below acceptable limits (4.6 psi).

CROSSFEED SYSTEM

1↔ 2 or

X-BP 1↔

Description

The crossfeed systegroup to be consupump from the oppoin Figure 5-4, the isystem are:

• Two booster cr(R1 and R3 in Fin the side-engfolds and cont

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

BP1

XT

R1

G1

G1 G3

ENGINE 1

AUTO

OPEN

CLOSED

BOOSTER

NORM

ST-BY

OFF

°C

LEGENDGROUP 1BOOST

GROUP 3

BOOST


TANKINTERCONNECTION

NOINTERCONNECTION

LIGHTS

XTKX-BPBP1BP3

NEUTRALCLOSEDONON

XTKX-BP

OUTOUT



*

*

Figure 5-7. Crossfeed X-BP 1↔ 3—Normal Con


XTK

X.BPLEVEL LEVEL

BOOSTER BOOSTER



XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

BP1

R1

G1

G1

ENG

AUTO

OPEN

CLOSED

BOOSTER

NORM

ST-BY

OFF

°C

LEGEND

GROUP 3BOOST


TANKINTERCONNECTIO

NOINTERCONNECT

LIGHTS

XTKX-BPBP1

BP3

NEUTRALOPENOFF

ON

XTKX-BP

OUTON

XTK

X.BP

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

BP1

R1

G1

G1

G1 G

ENG

AUTO

OPEN

CLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C


TANKINTERCONNECTIO

LIGHTS

XTKX-BP

RIGHTOPEN

XTKX-BP

ONON


*

*


XTK

X.BPLEVEL LEVEL

BOOSTER BOOSTER



XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

BP1

XTK

R1

G1

G1 G3

ENGINE 1

AUTO

OPEN

CLOSED

BOOSTER

NORM

ST-BY

OFF

°C

LEGENDGROUP 1BOOST


TANKINTERCONNECTION

NOINTERCONNECTION

LIGHTS

XTKX-BPBP1BP3

NEUTRALOPENONOFF

XTKX-BP

OUTON

XTK

X.BP

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

BP1

XTK

R1

G1

G1

G3

G3

G1

ENGINE 1

AUTO

OPEN

CLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C


TANKINTERCONNECTION

LIGHTS

XTKX-BPBP1

LEFTOPENON

XTKX-BP

ONON


*

*


XTK

X.BPLEVEL LEVEL

BOOSTER BOOSTER LEGEND

GROUP 1 BOOST

GROUP 2 BOOST



located in the group 1 interconnectionmanifold

The two crossfeed interconnect valves havetwo functions:

• To interconnect the fuel feed systems of No. 1 and No. 3 engines

• To cut off the fuel supply to the jet pumpson the side with the inoperative boostpump

The amber X-BP light on the overhead panel il-luminates when the associated crossfeed valve hasfailed to fully close.

Interconne

The interconnnects tank grofuel pressurenection valvoverhead panthe valve is no

XTK Solen

The XTK sol1 interconneWhen openeinterconnectithen illumina

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

AUTO

OPEN

CLOSED

BOOSTER

NORM

ST-BY

OFF

°C GROUP 3 BOOST


TANKINTERCONNECTIO

NOINTERCONNE

LIGHTS

BP1

BP2BP3X-BP 1–2X-BP 2–3

ON

NORMON

X-BPX-BP

OFFOFF

EN

CLOSEDCLOSED



*

Figure 5-10. Crossfeed X-BP 1↔ 2 and 3↔ 2


XTK

X.BPLEVEL LEVEL

BOOSTER BOOSTERLEGEND

GROUP 2 BOOSTGROUP 3 BOOST

BP1



XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

AUTO

OPEN

CLOSED

BOOSTER

NORM

ST-BY

OFF

°C

XTK

X.BP

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

AUTO

OPEN

CLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C

BOOST PUMP 1 FAILURE OR G1/G3 L

B

ST-B

ENGINE 1

LEGEND

GROUP 1 BOOST

GROUP 2 BOOST

BP1

B

ST-B

ENGINE 1


LIGHTS

BP1BP2BP3

X-BP 1–2X-BP 2–3

OFFNORMON

X-BPX-BP ONOFFOPENCLOSED


LIGHTS

NOT FEATURED ON AIRCRAFT

WITHOUT ELECTRIC TRANSFER

VALVE XTK 2



VALVE XTK 2

*

*

*

*


XTK

X.BPLEVEL LEVEL

BOOSTER BOOSTERLEGEND

GROUP 1 BOOST

GROUP 3 BOOST



XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

AUTO

OPEN

CLOSED

BOOSTER

NORM

ST-BY

OFF

°C

XTK

X.BP

XTK 2

X.BP X.BPLEVEL

ENG 1 ENG 2

FUEL SYSTEM

ENG 3

AUTO

OPEN

CLOSED

LEVEL LEVEL

BOOSTER

BOOSTER

NORM

ST-BY

OFF

BOOSTER

°C

BOOST PUMPS 2 FAILURE—EN

GROUP 3 BOOST


LIGHTS

BP1BP2BP3X-BP 1–2X-BP 2–3

ONOFFON

X-BPX-BP

OO

OPENCLOSED


LIGHTS

BP1 ON



VALVE XTK 2



VALVE XTK 2

*

*

*

*

spectively. Note that both the XTK and the in-terconnection valves are closed with respectto the interconnection lines and the No. 1 and

No. 3 boost pumps are operating.

Groups 1 and 3 crossfeed controls include:




p

• The X-BP 1↔ 3 rotary switch

• The three-position (left–center off–right)XTK switch, which functions only whenthe X-BP switch is set to open

If the No. 1 boost pump fails or is shut down,the corresponding crossfeed interconnect valve(labeled R1 on Figure 5-7) opens when selec-tion of X-BP 1↔ 3 is made, as controlledby a logic circuit of the correct crossfeedinterconnect valve, provided that the oppositeinterconnect valve is closed. This results in cut-ting off motive-flow fuel to the jet pumps on

the same side (which are off anyway becausethe boost pump is not operating).

Figure 5-8 shows configuration of the fuelpanel switches and the valve positions re-quired for crossfeed in the event of failure of the group 1 boost pump. Also shown is groups1 and 3 tank level equalization with the group1 boost pump inoperative or turned off. Figure

5-9 shows the same operations but with thegroup 3 boost pump inoperative or turned off.Included in both figures is a tabular explana-tion of the indications and operation.

X-BP 1↔ 2 or 3↔ 2Crossfeed

When the X-BP 1↔

2 or 3↔

2 rotaryswitch is positioned to open the respective X-BP 1 or 3 crossfeed valve, interconnection is es-tablished between the fuel supply system of thecorresponding engine (No. 1 or No. 3) and that

Figure 5-13 Fu

(T

set to normal, opening either X-BP 1↔2 or3↔ 2 crossfeed valve causes the standbygroup 2 boost pump to be energized. Thus, con-

sumption of the fuel in the front and rear tankstakes precedence over the corresponding wingtanks group without need to cut out the corre-

di id I thi th titi

tank.

• Green

This inulation(on air




sponding side pump. In this way, the quantitiesof remaining fuel can be equalized (the amountof fuel in group 2 being greater than that ingroups 1 or 3) without having to shut down thepumps.

Figure 5-10 shows the normal configuration of the group 2 fuel system. It normally supplies onlythe No. 2 engine. Figures 5-11 and 5-12 showfailed boost pump configurations and levelequalizing. Tabular data also shows switch andlight configurations as well as engine feed.

FUEL INDICATING

SYSTEMSFUEL QUANTITY INDICATORS

The three fuel quantity gages (Figure 5-13) areidentical. They have a round dial with a sin-gle needle and are gradated from 0 to 8,000pounds. Refer to Instrument Markings inChapter 1, “Aircraft General,” of this manual,for color pictures of the instrument markings.Following are the fuel quantity indicator in-strument markings.

Tank groups 1 and 3:

• Yellow arc ................. 0 to 1,000 pounds

Tank group 2:

• Yellow arc ................. 0 to 1,000 pounds

• Yellow arc ..... .. 2,200 to 2,400 poundsThis indicates the total q antit at hich

fer val

FUEL FLUSED IN

The fuel flo14) serve ducation is insthour by a neefrom 0 to 30 dow on the faquantity of f

A button is instrument fthe fuel-uselows displayseconds’ du

A FUEL FLthe cockpit depressed, tslew to 1,20

FUEL TEINDICATLOWER TOPTION

A digital disome aircraft

5-2) indicatfuel tank. Th

FUEL IND

that fuel pressure is low in the respectiveengine supply line.

Low Level WarningThe LO FUEL 1, LO FUEL 2, and LO FUEL 3li h h h f l l l i h i

• The access doconnector or tis open.

• Either the DEREFUELING

Th l




lights warn that the fuel level in the respectiveboost pump compartment is below 200 pounds.

NOTE

• For airplanes SNs 1 to 11—A fuel

level below 200 pounds is detectedin tank group G2 (or below 1,100pounds if booster pumps 2 are off).

• For airplanes SNs 12 and subse-quent—A fuel level below 200pounds is detected in tank groupG2.

1,000-Pound Level DetectionThree LEVEL lights are located next to theBOOSTER pump switches on the overheadpanel. During normal operations, each light willilluminate when the respective tank group fuellevel is below 1,000 pounds in groups 1 and 3 and1,100 pounds in group 2.

XTK OPEN and XTK CLOSEDWarnings—Aircraft Fitted withFuel Transfer Valve XTK 2(Aircraft without SB F900-48)

These lights signal malfunction of the trans-fer valve in the group 2 fuel transfer line. TheXTK OPEN light comes on to signal that thevalve is open when it should be closed. TheXTK CLOSED light comes on to indicate thatthe valve is closed when it should be open.

FUELING Warning

• The lever neawhich controraised.

• B2 bus is not

X-BP WarningThese lights are locand illuminate to wcrossfeed valves ha1↔3 light indicaX-BP 1↔3 cros


g t in

BAG1COMP1

;




20

15

10

5

0

FUEL QTY

LBS X 1000

20

1510

5

0

FUEL QTY

LBSX 100

BAG1COMP1

BLEEDAIR

FUELFLOW

TEST

ISOL

;

;

;

;

HEAT

GROUNDINGRECEPTACLE

LEFT REFUELINGSWITCH

GROUP 2INDICATION

FUELINGOK LIGHT

REFUELINGAUTOMATIC STOP

TEST PUSHBUTTON

REFUELING MODESELECTOR

(OPTIONAL)

STOPFUELING LIGHT

GROUP 1 INDICATION

PARTIAL REFUELINGSELECTOR

PRESSURE FUELING PANE

FULL FFULL

LEFT CENTER RIGH

ON

FULL

TEST

STOPFUELING

CLOSED

VENTVALVE TEST

OPEN

FUELINGOK

PARTIAL

OFF

ON

OFF

ON

OFF


FULL FULLFULL

LE FT C ENT ER R IGH T

ON

FULL STOPFUELING

CLOSED

ON

OFF

VENTVALVE TEST

OPEN

FUELINGON

PARTIAL

OFF

ON

OFF

ON

OFF

Figu



G1 END-OF-REFUELING

THERMISTOR

PRESSURIZATIONINTERCONNECTION

CENTER WINGTANK





FULL FULLFULL

LEFT CENTER RIGHT

ON

FULL STOPFUELING

CLOSEDTEST

DEFUELING

ON

OFF

VENTVALVE TEST

OPEN

FUELINGOK

PARTIAL

OFF

ON

OFF

ON

OFF

G1

PRESSURE REFUELING

ELECTRICAL

LEGEND

F

• A green FUELING OK light which illu-minates to indicate that the vent valvesare open and fueling may take place

• A FULL–PARTIAL toggle switch whichis used in conjunction with the partial re-fueling selector (when installed)

•The amb

If the STOP ing the refuefueling immrefueling sw




• A partial refueling selector which is usedwhen only partial refueling of the airplaneis desired

Normal OperationThe red STOP FUELING light should illuminatewhen the refueling control panel door is openedin conjunction with a previously opened con-nector panel door. This indicates that the ventvalves are closed. The FUELING light on thecockpit warning panel also illuminates if poweris turned on. On airplanes SNs 16 and subse-quent, a light is located in the upper part of the

refueling control compartment.

After opening the connector panel door, the ventvalve control lever must be moved before the re-fueling coupling (nozzle) can be connected.Moving the lever causes the vent valves to open.After the vent valves open, the red STOP FU-ELING light should go out and the green FUEL-ING OK light illuminate. Ensure that the nozzleis grounded to the airplane and the fuel truck and

that both the truck and the airplane are properlygrounded. The coupling cap may then be removedand the nozzle connected.

After the nozzle is connected, place the three re-fueling toggle switches to ON. This opens the re-fueling solenoid valves (Figure 5-16), which willbe powered only if the vent valves are fully open.Check that the fuel truck pressure available is

30–50 psi, and then start refueling.

After refueling is started, push the TEST but-ton to test the automatic refueling stop system,and check for the following:

After compling switches

Partial R

To prepare fprocedure aunder Norm

To refuel torefueling sesired. PlacePARTIAL. Tthe quantity

up equally bcan be checkdicators. Eamatically clogroup probeceived a quaselected. Thand the refuefrom the batt

in the cockp

GRAVITY

The airplaneity filler porwing’s uppeairplane is Refueling p

quantity ind

Refueling thquires press

When opened, the interconnects the grfueling system (Se

all the tanks througsupply circuit is enboost pumps and opvalves for group 1 a




NOTE

If gravi ty-refuel ing the airplane

through only one wing filler port,open the crossfeed valve and turn onthe boost pump for that side only.

• Place the GRAVITY FUELING switch onthe cockpit maintenance panel to ON.This opens the defueling/refueling valvein the Group 2 manifold (Figure 5-18) andthe vents.

• The CENTER refueling switch in thepressure-refueling panel can now beplaced to the ON position to open thecenter tank’s pressure-refueling valve.

NOTE

Gravity refueling should not be done

using battery power only because of the high current draw during refueling.

DEFUELING

valves for group 1 a

DRAIN VALVE

The sump drain vala means of eliminati

contaminants whicthe tanks. Fuel sampfore and after refuesample can be takenis allowed to settletime permits. A totlocated under the wcraft with SN 158 aand subsequent havdrain valve.

ANTI-ICING BAPPARATUS

Figure 5-20 shows tcal for adding fuelthe wing through th

necessary. The fueadding the additive has been added.

Figure 5-17. Fuel Filler Port

FALCON 900 PILOT TRAINING MANUALg t

in

FRONT

TANKGRAVITY

FUELING

VENT

VALVE

F U E L

J E T A

, JE T A1, J E T B , J P 4 ,

J P 5

C A U T

I O N : PRESSU R I Z E D

T A N K

F

O R F U E L I N G

I N S T R U C T IONS SE E O P E R G

M A N U

A L

T O T A

L A I R

C R A F T U S AB LE QUA N T

I T Y 2 8

4 5 U S G A L




;

; ;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

G1

D

R

BOOSTER

CROSSFEED

MANIFOLD

CROSSFEED

VALVE

GROUP 1 BOOST

LEGEND

GROUP 3 BOOST

ELECTRICAL

BAG 1 COMP 1BLEEDAIR

FUEL FLOW

TEST

ENG INDIC

TEST

GRAVITY

FUELING

STABILIZERRELAY TEST

COOLINGFANANTI-SKID

ISOL

;

;

;

;;

;

;

;

;

;;

;

;

;

;

;

;

;;

;

;

;

;

;

;

;

;

;

;

;

;

HEAT

ON

OFF

RESET

MAINTENANCE PANEL

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

@ @ @

@ @ @

; ; ;

; ; ;

y y y

y y y

; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;@ @ @ @ ; ; ; ;y y y y ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;@ @ @ ; ; ;y y y

;

;

;

;

y

y

y

y

;

;

;

y

y

y

; y

;

;

;

y

y

y

;

;

y

y



Fuel used must conform to the following spec-ifications. Table 5-2 is representative of thefuel definition as of December 1987.




DESIGNATION

GARRETT

SPECIFICATION

EQUIVALENCE

WIDE-CUT-TYPE FUEL

HIGH-FLASHPOINT-TYPE

FUEL

CIS

KEROSENE

EMS 53111

EMS 53112

EMS 53113

EMS 53116

GOST

10227-86

ASTM D 1655 Jet A –4CAN 2-3.23 Jet A

ASTM D 1655 Jet A1 –4CAN 2-3.23 Jet A1DERD 2494 AVTURDERD 2453 AVTUR/FSii

MIL-T-83133 JP8 –5AIR 3405C -AIR 3405C -

ASTM D 1655 Jet B –5CAN 2-3.22 Jet B

MIL-T5624 JP4 –5AIR 3407B -DERD 2486 AVTAGDERD 2454 AVTAG/FSiiCAN 2-3.22 -

AIR 3404C - –4AIR 3404C -DERD 2498 AVCATMIL-T-5624 JP5

DERD 2452 AVCAT/FSiiCAN 3GP24 -CAN 3GP24 -

T1 –6TS1 REGULAR –6

Table 5-2. FUEL SPECIFICATIO

FREEPOI(° C

1.The groups 1 and 3 tanks are described as:A. Integral tanks

B. Bladder tanks

C. Strap-in metal tanks

6.Internal tank faccomplished

A. Bleed-air p

B. Jet pumps u


QUESTIONS



C. Strap in metal tanks

D. Isolated tanks

2. Group 2 tanks are described as:

A. Integral tanks

B. Bladder tanksC. Strap-in metal tanks

D. A structural part of the airplane

3. The Minimum Equipment List allows dis-patch with:

A. Two boost pumps inoperative, pro-vided that they are not in the same

tank groupB. One boost pump inoperative, pro-

vided that it is in group 2 tanks

C. One boost pump inoperative, pro-vided that the wing tanks arepressurized

D. Two boost pumps inoperative, pro-vided that the fuselage tanks are

pressurized

4. Group 2 fuel is normally used as follows:

A. 1,000 pounds from wing tanks, thenthe rear tank fuel

B. Front and rear tank fuel, sequentially

C. All rear tank fuel, then the fronttank fuel

D. Not until the wing tank fuel isdepleted

5. The front-to-rear transfer valve manualt l h dl i ft

p pfrom the enfuel pumps

C. Jet pumps ufrom the bo

D. DC-powere

7. The number ofstalled in the F

A. Two

B. Three

C. Four

D. Five

8. Motive flow foby the:

A. Engine-dri

B. Transfer pu

C. Boost pum

D. Scavenge p

9. The jet pumps

A. All fuel is pump comp

B. Low-pressuengine fuel

C. The boost p

D. A and B

10. During normalBOOSTER swi

A Neither bo



22. Maximum fuel truck pressure forpressure refueling is:

A. 30 psi

B. 40 psiC. 50 psi

D. 55 psi




CHAPTER 6

AUXILIARY POWER U

CONTENTS




INTRODUCTION ......................................................................

GENERAL..................................................................................

MAJOR SECTIONS..................................................................

Air Intake ............................................................................

Compressor .........................................................................

Combustor...........................................................................

Turbine ................................................................................

Exhaust................................................................................

Accessory............................................................................

OPERATING PRINCIPLE........................................................

APU SYSTEMS .........................................................................

Oil System..........................................................................

Fuel System........................................................................

Ignition System ...................................................................

Instrumentation ...................................................................

Control and Starting...........................................................

Electrical Power Supply......................................................

ILLUSTRATIONS

Figure Title

6-1 APU Installation........................................................

6-2 Major Sections...........................................................




6-3 APU Layout...............................................................

6-4 Operating Principle ...................................................

6-5 Oil System Operation................................................

6-6 Start and Speed Governing Logic .............................

6-7 Fuel System Operation..............................................

6-8 Overhead Switch Panel .............................................

6-9 Battery Start—APU (RPM Less than 50% N1) ........

6-10 GPU Start—APU (RPM Less than 50% N1) ...........


AUXILIARY



INTRODUCTION

This chapter describes the auxiliary power unit (APUseries airplanes. The primary objective for installing APU

airplane self-supporting while on the ground, especially whbases where compatible ground servicing equipment may

APUs are generally independent of all installed airplane s


GENERATORCOOLING AIR

APU COMPARTMENT



COOLING AIRINLET

GENERATORCOOLING AIR

OUTLET

VENTILATIONAIR INLET

LEFT RIGHT

GENERAL

The Falcon 900 incorporates an APU manufac-

tured by Garrett Turbine Engine Company; itis designated GTCP36-150 (F). The unit istransversely mounted in the rear fuselage(Figure 6-1), under the No. 2 engine, and is en-closed in a carbon and titanium container The

AIR INTA

The APU ai

unit; the opeof foreign mthe intake fr6-3) locatedof the aft fu




closed in a carbon and titanium container. TheAPU is certificated only for ground opera-tion; the control circuits are wired throughthe landing gear ground contact switches, thuspreventing in-flight operation or starting.

MAJOR SECTIONS

For descriptive purposes the APU engine is di-vided into six major sections (Figure 6-2):

• Air intake

• Compressor

• Combustor

• Turbine

• Exhaust

• Accessory

COMPRE

The compre

impeller witimpeller is riThe functionrequired airand direct it which conveenergy. Thiscombustion ply for grou

cupied areas

COMBUS

The combuslar combustioplenum. Air the compres

LEGEND

ACCESSORY GEAR

AIR INTAKE

COMPRESSOR AIR

TURBINE

COMBUSTOR

EXHAUST

ACCESSORY GEAR

AIR INTAKE

COMPRESSOR

TURBIN




APUBLEEDVALVE

NO. 2 ENGINEBLEED-AIR

LINE

NO. 2 ENGINEAIR INTAKE

THERMALEXPANSION

CONNECTOR

APUCOMPARTMENT

GENERATORVENTILATION

AIR INLET

APU COMPARTMENTVENTILATION

AIR INLET

GENERATOR

VENTILATIONAIR OUTLET

APU

DRAIN MANIFOLDSB-900-109

Figure 6-3. APU Layout



APU SYSTEMS

The APU systems consist of the following:

• Oil

• Fuel• Ignition

• Instrumentation

• Control and starting

• Electrical power supply

• Bleed-air supply

OIL SYSTEM

The oil system (Figure 6-5) provides for cool-ing and lubrication for the rotor bearings and

A gear pump, driveprovides the pressulubrication. The putrolled by a pressursure is sensed by a pon an amber light lcontrol panel (“Annwhen oil pressure dAPU N1 is greater automatically shut dtemperature greatilluminate the OILshut down the APUarmed only when th

is latched in (on). Asure line removes oil. Ferrous metal magnetic drain pluginspection

LEGEND

COMBUSTIONCOMPRESSOR AIRAIR INTAKE

Figure 6-4. Operating Principle


MAINSHAFTBALL BEARING

DRIVE PAD(NOT USED FOR FALCON 900)

VENT

COOLING FINS



FUEL SYSTEM

The fuel system consists of an electronicallycontrolled fuel control unit (FCU), a fuel pump,a bypass fuel filter, a torque motor, an ultimaterelief valve, a differential pressure regulator, afuel solenoid shutoff valve, a flow divider, andtwo fuel manifolds terminating in six s implexatomizers (three primary and three secondary),located in the combustion chamber. The FCUi t ll d b l t i t l it (ECU)

at a constantthe FCU.

Ultimate The ultimate

pressure of tcially at shu

Fuel Sup

PRESSUREVALVE

FILTEROILSUMP

GEAR-TYPE PUMP

SUCTIONMAGNETIC

DRAIN PLUG

OILTHERMOSTAT

FUEL CONTROL UNIT

STARTER- GENERATOR

BALL BEARING

LOWOIL PRESSURE SWITCH

Figure 6-5. Oil System Opera

Differential Pressure Regulator

The differential pressure regulator maintainsa constant differential across the meteringvalve since flow is proportional to meteringvalve position. Fuel in excess of the me-tered volume is returned to the pump inletby a bypass line. The metered fuel for start-ing, acceleration, steady-state operation,

Electronic Co

The solid-state ECUfor starting, speed gfor the APU. The nals representing Amounted on the accis backed up by atopping circuit dur




ing, acceleration, steady state operation,and deceleration is a sole function of theECU.

topping circuit durfrom light-off to ssignal is received couple probe in the

APU RPM APU EGT

MAX CURRENT SCHED

TURBINE SPEED (%)25 TO 95% 16 SEC

TIMED ACCEL

LEGEND

FUEL CONTROL TORQUEMOTOR METERING VALVE

00

25 50 75 100

C U R R E N T ( M A

)

1,000

1,100

1,200

1,300

1,400

1,500

1,600

T E M P E R A T U R E

( F )

ECU

Speed GoverningThe speed governing system is a solid-statetimed-acceleration control, which operates

on the principle of governor reset acceleration.The governor is initially preset at 25% of max-imum design rpm. Following light-off, thegovernor set point increases as a function of time, and the resulting signal is transmitted toh h FCU i l

• T5 open

• N1 mon

• Overcur• Low oil

• High oi

• Takeoff




the torque motor on the FCU metering valve,which responds by adding or subtracting fuelto maintain the rpm at the governor reference.This reference increases as a function of time

until the 100% reference is attained. This valuewill now be held constant throughout the en-tire operating range.

The T5 loop has no function during APU start-ing unless T5 temperature exceeds the designvalue.

If the APU meets any resistance to accelera-tion between light-off and self-sustaining rpm,the closed loop temperature topping circuit willassume precedence, signaling the torque motorto reduce fuel flow.

In summary, the electronic speed governingloop is constantly monitoring rpm and mod-ulating fuel flow by comparing the monopoleinputs to the ECU’s 100% reference rpm, gen-erating the required error signal to vary the cur-rent to the torque motor, and repositioningthe fuel metering valve accordingly.

Figure 6-6 depicts the start and speed gov-erning logic of the APU’s ECU.

ProtectionThe APU incorporates a comprehensive pro-tection system which will automatically shutthe unit down when any of the following con-di i i

• Faulty A

• All battswitche

• APU fir

• Environ

When an APlisted conditsystem is reAPU MASTexplained in

OperatioAs shown infuel at boostfilter from tThe HP pumrequired valscreen to the

of the fuel mtorque motosignals frommally the conals supplieon the accestemperatureimum, the Tin an output

back fuel. Ting system, stantaneouslloads) and m

DC POWER

EGRPM

APU MASTER

PUSHBUTTON

FROM

NO. 2 ENGINE

FUEL SUPPLY

ECU

SAFETYCIRCUITS




The ignition system is controlled entirely bythe ECU. The ignition is turned on at 10%rpm and continues until rpm reaches 97% +4seconds, at which time it is turned off by asignal from the ECU.

The ignition system installed on theAPU is a high-energy type capable of inducing an output voltage of be-tween 5 000 and 6 000 volts Even

INSTRUMENT

The APU incorpodicator located onoverhead switch phourmeter located

The N1 indicator is 0 to 120. The dial facber, and red. The grto 105%, the amber 110%, and the red raThe gage is a sweep

WARNING

LEGEND

TANK BOOST PRESSURE

HP PUMP PRESSURE

BYPASS FUEL

METERED FUEL

PRIMARY NOZZLE FUEL

SECONDARY NOZZLE FUEL

ELECTRICAL

FILTER BYPASS

TORQUE MOTOR

S

S

FILTER

HP FUEL PUMP

ULTIMATE RELIEF

REGULATOR

SCREEN

FUEL

SUPPLY

SOV

/P

Figure 6-7. Fuel System Operation

the power sePOWER posto close theTIED light i

The folscriptionot be c




The T5 indicator is graduated in degrees Cel-sius from 0 to 1,000. The dial face is color-coded green, amber, and red. The green rangeextends from 150 to 679°C, the amber rangeextends from 679 to 732°C, and the red radialmarker is located at 732°C. Operating poweris provided from the ECU through the latched-in (on) position of the APU MASTER push-button. When power is turned off, the needledrops to an ambient temperature indication.

An hourmeter is mounted in the APU enclosureand is accessible by opening the No. 2 enginecowl. An option will allow installation of thehourmeter in the rear compartment. The unitreceives power from the ECU when enginerpm is 97% +4 seconds. When the APU is shutdown, power is removed from the hourmeterby the ECU when rpm decays below 97%.

CONTROL AND STARTING

Prior to starting the APU, a safety check mustbe performed. This should include an oil levelcheck as well as a security check of the APU

not be cAPU mproved

Battery S(RPM LesSelect the NSTBY (Figu2 light on thgoes off. Roizontal posit

light on the wgreen APU (to the latchegreen APUlatched-in poon in the MApower is supsupply shutoAPU GEN am

rpm gage negage indicatcombustion

If the Aing a shdelayed

least 20normal

Momentarily

Figure 6-8. Overhead Switch Panel

terminated; the APU OIL light must go off be-fore 50% rpm. On some installations, at ap-proximately 60% rpm compressor dischargepressure opens the surge valve. RPM must con-tinue to increase rapidly and smoothly; at 97%+4 seconds ignition is terminated, and thehourmeter begins to record operating time. TheAPU GEN light goes out, indicating that thegenerator is on line and battery charging is tak-i l Th APU ti t l t t

inc orporated, is enenoid valve opens, fubustion chamber thronozzles, and combsmooth accelerationat the flow divider rvalve opens, and fusecondary spray nozshuts down starter opti d b f 5




ing place. The APU continues to accelerate toapproximately 101.6%. APU bleed air is avail-able for ground heating or cooling if and whenselected.

APU generator voltage is indicated on both DCvoltmeters. APU line load may be checked bymoving the selector switch under the rightammeter to the APU position. When eitherammeter selector swi tch i s a t the BATposition, the ammeters indicate recharging of the associated battery with the bus tie closed.

If the APU GEN light remains onabove 97%, the APU start relay mayhave failed to open. Push to unlatchthe APU excitation pushbutton; thegreen light should go out, and theAPU automatically shuts down. Have

maintenance investigate and correctthe problem.

Figure 6-9 depicts a battery start of the APU.Prior to initiating the start, both battery switchesand the bus-tie switch are on, BAT 1 and BAT2 lights are out, and the BUS TIED light is on.The APU excitation pushbutton is latched in, andits green light is on. The APU MASTER push-

button is latched in, and its green light is alsoon. The amber APU OIL and GEN lights are bothon. The No. 2 engine standby boost pump is on,the APU fuel supply shutoff valve is open, and

tinues, and before 5When a surge valvesor discharge pressuAt 97% +4 secondsnition, the generatorlight goes out. If theswitch is on, the Aareas for ground cooThe APU rpm stabilconstant at approximlizes at approximate

GPU Start—A(RPM Less thaFigure 6-11 depictsas the power sourcethe APU start cycle,operating. When thmoved to EXT POWsimultaneously: the

and No. 2 battery maBAT 1 and BAT 2 ligtery make-and-breabecause DC powerPOWER position ofall generators includinhibited, and the GGPU output is routea diode to the batter

the No. 1 battery maleft main bus and bu

Note that in this con

CAUTION

R ev i s i on 3

NORM

OIL

APU STOP

START

GEN 1

BAT 1

GEN 1 GEN 3

BAT 2

GEN 2 APU

BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

GEN 3

I RS 1 H RZ N I RS 2 I RS 3

TO

AMMETER

DC SYSTEM

OVERHEAD PANEL

GEN

EXT POWER



6 - 1 3

F

OR T RAI NI N G P URP O S E S

ONL Y

G1

GENERATOR 1

G2

GENERATOR 2

TO

AMMETER

G3

GENERATOR 3

START

CONTACTOR

TO AMMETER

G

START CONTACTOR

BATTERY 1

CONTACTOR

BATTERY 1

START

CONTACTOR

GPU CONTACTOR

APU

TO

AMMETER

TO

AMMETER

FIRE

EXTINGUISHERS

COCKPIT

DOME LIGHT

TO

AMMETER

GPU RECEPTACLE

BATTERY

BUS

START

CONTACTOR

BATTERY 2 BATTERY 2

CONTACTOR

STARTING

BUS

Figure 6-9. Battery Start—AP

6 - 1 4

F O

NORM

EXT POWER

OIL

APU STOP

START

GEN 1

BAT 1

GEN 1 GEN 3

BAT 2

GEN 2 APU

BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

GEN 3

GEN

I RS 1 HR ZN I RS 2 I RS 3

G1

GENERATOR 1

TO

AMMETER

START

CONTACTOR

DC SYSTEM

OVERHEAD PANEL



R ev i s i on 3

OR T RA


O

NL Y

Figure 6-10. GPU Start—APU

G2

GENERATOR 2

TO

AMMETER

G3

GENERATOR 3TO AMMETER

G

START CONTACTOR

BATTERY 1

CONTACTOR

BATTERY 1

START

CONTACTOR

GPU CONTACTOR

APU

TO

AMMETER

TO

AMMETER

FIREEXTINGUISHERS

COCKPIT

DOME LIGHT

TO

AMMETER

GPU RECEPTACLE

BATTERY

BUS

START

CONTACTOR

BATTERY 2 BATTERY 2

CONTACTOR

STARTING

BUS

When using a GPU to start the APU, the startcycle differs from the previously described bat-tery start of the APU because the batteries arecompletely isolated, even from the battery bus,and the No. 2 battery contactor remains open.Therefore, when the APU START switch is mo-mentarily pushed, the only contactor which re-sponds is the APU start contactor; it closes andremains closed until the APU start is termi-nated by the ECU at 50% When the APU ac-

unlatched plight goes obelow 0, and

When thSTOP pmatic shnot be re




nated by the ECU at 50%. When the APU accelerates above 97% rpm, the APU GEN lightremains on since all generator operation is in-hibited as long as the GPU generator continuesto power the distribution system.

ELECTRICAL POWER SUPPLY

The electrical power supply from the APUgenerator is described in detail in Chapter 2,“Electrical Power Systems.”

BLEED-AIR SUPPLYThe bleed-air supply from the APU is describedin detail in Chapter 11, “Air Conditioning.”

APU SHUTDOWNSYSTEMS

The APU incorporates automatic and selectiveshutdown systems. The automatic shutdownsystem is a function of the ECU and is de-scribed in detail under Electronic Control Unit(ECU) in this chapter.

The normal means of shutting down the APUis by momentarily pushing the STOP pushbut-ton (Figure 6-8). The STOP pushbutton trans-mits an input to the ECU which it interprets asan overspeed (114%) signal. The overspeedlogic shuts down the APU by closing the APU

not be repushbutand rela

If the APU fpushbutton iavailable. MMASTER puthe APU fuelfrom the ECshutoff valvbe written u

the problem automatic ov

LIMITA

The APU mu

Operation ocabin and noauthorized.

Maximum N

Exhaust gas

If no IT10 secoSTOP putes bef

NOTE

The duration of operation on amberrange 679 to 732°C (1,255 to 1,350°F)

must be as short as possible (intendedfor emergency operation only).

Maximum generator output:

• Transient .......................... 350 amps




(One minute maximum)

• Stabilized.......................... 300 amps

Refer to fuels and oils for the engine.

QUESTIONS


1. Fume extraction and cooling in the APUcompartment is achieved by:

A. A fan driven by the starter-generator

B. Ram air

C. An exhaust gas venturi effect

D. A bleed-air-driven ground blower

6. The T5

A. Provduri

B. Modvalv

C. Maifuel



2. The APU pushbutton must be latched(green light on) initially:

A. After the APU rpm reaches 97% +4seconds

B. At 10%, to turn on the ignition

C. To open the fuel supply valve

D. To excite the starter-generator, be-fore pushing the START switch

3. Automatic APU shutdown occurs if:

A. The battery and engine generatorswitches are turned off.

B. Light-off does not occur within tenseconds.

C. The OIL light is on at 20% rpm.

D. EGT hangs at 25% rpm.

4. Pushing the APU MASTER pushbutton

(green light on):A. Turns on the OIL and GEN lights

B. Supplies power to the ECU

C. Opens the fuel supply valve

D. All the above

5. At 97% +4 seconds rpm the ECU:

A. Terminates the start cycle

B. Turns off the OIL lightC. Terminates ignition

D. Opens the bus tie

D. Conerat

7. The GT

A. Vari

B. Con

C. Twi

D. Sing

8. When thmomenof the A

A. Bothswit

B. Theclos

C. TheAPU

D. The

9. During

A. Botmai

B. Theratio

C. Botstar

D. Boting

CHAPTER 7

POWERPLANT

CONTENTS

INTRODUCTION




INTRODUCTION ......................................................................

GENERAL..................................................................................

ENGINES ...................................................................................

General ................................................................................

Ratings ................................................................................

Characteristics.....................................................................

Major Sections ....................................................................

OPERATING PRINCIPLES......................................................

ENGINE SYSTEMS .................................................................

General ................................................................................

Oil System...........................................................................

Fuel System.........................................................................

Ignition System...................................................................

Instrumentation ...................................................................

Engine Power Control.........................................................

Engine Starting ...................................................................

Engine Failure (No. 2 Engine)............................................

Maximum Engine Rotor Speed (N1 and N2)..................................

Maximum Interstage Turbine Temperature (ITT) .........................

Starting Time .................................................................................

Fuel Computers..............................................................................

Thrust Reverser..............................................................................

A d Oil




Approved Oils................................................................................

Oil Pressure....................................................................................

Oil Temperature .............................................................................

Placards and Instruments ...............................................................

QUESTIONS.........................................................................................

ILLUSTRATIONS

Figure Title

7-1 Engine Installation......................................................

7-2 TFE731 Engine (Typical)...........................................

7-3 Major Sections............................................................




7-4 LP Compressor Surge Bleed Valve ............................

7-5 Surge Bleed Control Solenoid Valves ........................

7-6 LP Compressor Surge Bleed Valve—Open.................

7-7 LP Compressor Surge Bleed Valve—Closed .............

7-8 LP Compressor Surge Bleed Valve—1/3 Open...........

7-9 Operating Principles (Gas Flow) ................................

7-10 Oil System Components and Servicing.......................

7-11 Oil System Miscellaneous Components......................

7-12 Indication and Warning Systems.................................

7-13 Oil System Operation..................................................

7-14 Fuel System Components............................................

7-15 Fuel System Layout.....................................................

7-16 Digital Electronic Fuel Computer (Typical)................

7-17 PT2TT2 Probe ..............................................................

7-18 Surge Bleed Valve Operation ......................................


7 20A P I S

7-23 Ignition System Components .................................................

7-24 Ignition System Operation .....................................................

7-25 Engine Instrument Panel ........................................................

7-26 ITT Thermocouple System Operation ...................................

7-27 Maintenance Test Panel..........................................................

7 28 Center Pedestal




7-28 Center Pedestal.......................................................................

7-29 Battery Start—No. 3 Engine (RPM < 50% N2)......................

7-30 Battery/APU Generator-Assisted Start—No. 2 Engine(RPM < 50% N2) ...................................................................

7-31 Battery/APU Generator-Assisted Start—No. 3 Engine(RPM < 50% N2) (No. 2 Generator On)................................

7-32 Battery/APU Generator-Assisted Start—No. 1 Engine

(RPM < 50% N2) (No. 2 and No. 3 Generators On)..............7-33 GPU Start—No. 2 Engine (RPM < 50% N2).........................

7-34 Airstart Envelope....................................................................

7-35 Thrust Reverser—Stowed ......................................................

7-36 Thrust Reverser—Deployed...................................................

7-37 Thrust Reverser—Hydraulic Operation .................................

TABLES

Table Title

7-1 Computer Malfunction Analysis ............................................


P



INTRODUCTIONThis chapter describes the powerplants installed on the Faddition to the engine, the chapter also includes descriptioas oil, fuel, ignition, instrumentation, engine power contverser, and associated powerplant limitations.

The contents of this chapter must not be construed as opefor pressure, temperature, and thrust are used only for theivalues must be obtained from the pertinent material issuplicable manufacturer, the Federal Aviation Administratioof the country of origin.


ENGINES

GENERAL

The engines (Figure 7-1) are installed in a con-ventional trijet configuration. The No. 1 andNo. 3 engines are pylon-mounted on the leftand right sides of the rear fuselage. The No. 2engine is mounted internally in the tail cone.

The maximum conand Mach 0.8 is 98TFE731-5AR-1C orthe TFE731-5BR-1

CHARACTER

The major charactemodular design coengine maintenandowntime.



The engines (Figure 7-2) are manufacturedby Garret t Turbine Engine Company in

Phoenix, Arizona, a division of the HoneywellEngine and System Division. They are desig-nated TFE731-5AR-1C or TFE731-5BR-1C,if SB F900-100 is incorporated.

RATINGS

The rated static thrust at sea level and 23°C(73.4°F) is 4,500 pounds of thrust for theTFE731-5AR-1C or 4,750 pounds of thrustfor the TFE731-5BR-1C at sea level and25°C (77°F).

downtime.

MAJOR SECT

General

For descriptive purpinto 10 major sectio

• Air inlet

• Fan

• Planetary gea

• Low-pressure

• High-pressure

• Combustor

• High-pressure

• Low-pressure

• Exhaust and c• Accessory ge

Air Inlet

The air inlet is formis directed into the fair inlet duct. Aft ofvided into two flow

ters the core enginefull-length bypass d

F


FAN INLET HOUSING

FAN ROTOR ASSEMBLY

FAN SPINNER

OIL FILLER TUBE

IGNITION U



assembly consists of the fan disc and a coni-cal inlet spinner. The fan blades have integralpart-span shrouds which butt each other toform a continuous part-span ring.

The fan performs a twofold function: its outerdiameter accelerates a moderately large airmass at a relatively low velocity into the full-l e n g t h b y p a s s d u c t ; i t s i n n e r d i a m e t e r

by the 731 from the rea10,416 for TFTFE731-5B

PlanetaryThe planetarately aft of t

FLOW D

TRANSFER SHAFTCOVER TUBE

ACCESSORY DRIVE GEARBOX

FUEL PUMP ANDINTEGRAL FUELFILTER BYPASS

INDICATOR VALVE

TRANSFER GEARBOX ANDN

2 TRANSDUCER

Figure 7-2. TFE731 Engine (

FUENOZZ

LOW-PRESSURECOMPRESSOR

FAN REDUCTIONGEAR

FAN

BYPASS AIR

INLET

AIR INLET

7 - 4

F OR



CORE AIRINLET

ACCESSORY GEAR TRANSFER GEAR HIGH-P

COMP

R ev i s i on1

T RA


ONL Y

Figure 7-3. Ma

LP Compressor

General

The LP compressor has four axial stages. Theairflow to the LP compressor is supplied by theinner diameter of the fan. The compressor dif-fusion process increases air pressure and ve-locity and directs the air mass rearward througha diffuser to the HP compressor.

Th LP i d i b h LP bi




The LP compressor is driven by the LP turbine.The direction of rotation, as viewed from the

rear, is clockwise.

Surge Bleed System

The LP compressor has a surge bleed systemsince axial compressors tend to stall and surgeunder certain conditions. For each compressorrpm, there is a relationship between its pressureincrease and its air mass. Under certain

conditions such as low power, acceleration, anddeceleration, this balance may be disturbed. Thecompressor, in effect, is an air pump, but the airmoving through it meets many obstacles, such ascombustion pressure. When overfueling occurs,such as during acceleration, combustion pressuremay exceed compressor outlet pressure,compressor airflow slows down, and compressorblade stall occurs, resulting in a pileup of air

followed by a surge.

Stall and surge conditions may also occur duringperiods of HP/LP spool rpm mismatch, since airstacks up at the HP compressor, producing stallsand surges which, if not controlled, could causeflameout and/or mechanical failures.

A safe surge margin has been determined for theTFE731 engine. It is controlled by an electro-pneumatic surge valve (Figure 7-4) mounted atthe 3-o’clock position on the engine. The surge

Figure 7

The solenoids in turn are alternately controlledby outputs from the digital electronic fuel com-puter (to be described in the Fuel Computersection).

The surge bleed valve has three positions: open,1/3 open, and closed. The solenoid valves arecalled A and B and they control P3 (HP com-pressor discharge) air for operation of the surgebleed valve.

I Fi 7 6 h bl d l i i

and permitting the spilling LP air into

In Figure 7-7 the su

since P3 air is shut noid A, while the dmits P3 air, overcomthe valve.

In Figure 7-8 the s1/3-open position be

i d i i




In Figure 7-6 the surge bleed valve is open sincesolenoid B is energized, venting chamber B

energized, permittin

LP COMPRESPILL TO BY

ORIFICE

CHAMBER

CHAMBER A

SPRING

PORTPORT A

POPPET

FLOW

FLOW

SOLENOID A


ORIFICE

PORT A

POPPET

FLOW



ORIFICE

C

CHAMBER A

SPR

PORT A

POPPET

FLOW

SOLENOID A

CHAMBER A

SP

VENT

SOLENOID A

ENERGIZEDP

3 SUPPLY PRESSURE

Figure 7-7. LP Compressor Surge Bleed V

produce a balance which will allow the valveto move to, and remain in, a 1/3-open position.This position is called the fail-safe position be-cause the valve will assume it when electrical

power fails or if the fuel computer is turnedoff. (See Fuel Computer later in this chapter.)

The computer controls the surge bleed valveand maintains the safe surge margin based onN1 and N2 mismatch and the ratios of fueloperating schedule.

chamber and form the fan bypass air toing for the hot sectreenters the bypass

the combustor.

HP TurbineThe single-stage, axidly connected to thigh-pressure rotortracts sufficient en




HP Compressor

The HP compressor is a single-stage centrifu-gal compressor. It is located aft of the LP com-pressor. A face shroud on the compressor discdirects all airflow through the disc blades.The velocity energy leaving the compressor isconverted to high-pressure energy by a di-vergent duct. It is then directed rearward to thecombustor.

The HP compressor is driven by the HP tur-bine. The direction of rotation, as viewed fromthe rear, is clockwise.

The LP and HP compressors provide a i rfor cooling, combustion, and the airplane’spneumatic services.

Combustor

The combustor consists of a reverse-flow com-bus t ion chamber , loca ted in a p lenum,surrounding the turbine. The combustionchamber includes 12 equally spaced duplexfuel-spray nozzles and two high-energy ig-niter plugs. The function of the combustion

chamber is to control the mixing of fuel andair and, following ignition of the mixture bythe two igniter plugs, to contain the combus-tion gases and direct them for expansion

tracts sufficient encombustion gases to

and the accessory g

The HP turbine nocooled by comprescored passages.

The HP turbine anding assemblies consengine. The directio

The rpm of the HP sthe TFE731-5AR-1CTFE731-5BR-1C a

LP Turbine

The three-stage axiaconnected to the fouthe LP rotor shaft.

cated within the HPof rotation is clockw

The LP turbine discdiameter from frontfore, extracts an equthe expanding gasedrive the four-stagsingle-stage fan thro

The LP turbine, the Lform the LP spool the LP spool is 21 0

The combustion gases from the core engine aredirected into a 14-petal, daisy-shaped coremixer, which directs the combustion gases tomix with fan bypass air in the bypass duct. This

principle increases direct thrust and also re-duces gas-velocity-generated noise. The mixedgases are directed to the atmosphere throughthe exhaust duct and provide the propulsiveforce for the airplane.

Accessory Gear

pressure of tLP compresscompressor, creased by a

to the combu

A precise amflow combu jec ted by tThe mixture energy igni




The accessory gear (housed in two separate

gearboxes) is mounted on the lower forwardside of the engine. It consists of a transfergear and an accessory drive gear. The trans-fer gear is driven by a towershaft throughbevel gears on the HP rotor shaft, which in turndrives the accessory gear through an inter-connecting horizontal drive shaft.

The following accessories are driven by the ac-

cessory gear:

• Fuel pump assembly and fuel controlunit (FCU)

• Lubricating pump

• Hydraulic pump

• DC combination starter-generator

A dual-wound monopole is installed in thetransfer gearbox to provide rpm input signals tothe analog N2 gage in the cockpit and to thedigital electronic fuel computer.

OPERATINGPRINCIPLES

energy ignithrough the

enough enerthrough the mgear throughshaft, whichthe transfer g

The LP turbdrive the LPshaft, which

etary gear an

The remaininate through tpass airflow it to the atmofor the airpla

When the en

the ignition s

ENGIN

GENERA

The engine s

• Oil

• Fuel

CORE ENGINE



7 - 9

F O

R T RAI NI N G P URP O

S E S

ONL Y

LEGEND

INLET AIR (CORE AND BYPASS)

LP COMPRESSION

HP COMPRESSION

CORE ENGINE EXHAUST

MIXED EXHAUST

143 LB/SECTOTAL

Figure 7-9. Operating P




OIL PUMP FUE

OIL SYSTEM

General

The oil system is a self-contained, pressure-scavenge system, which provides for coolingand lubrication of the main engine bearings,the planetary gear system, and the accessorygears and bearings.

The oil is contain(Figure 7-10) mounof the fan inlet casfiller, dipstick, an

engine is serviced ththe left engine macrossover tube. Thserviced through thetube. In all casescowling must be ope


SEGMENTEDOIL-TO-AIR



SWITCH CLOSES WHEN REARCOMPARTMENT DOOR IS OPENED.

REAR DOORMICROSWITCHBATTERY BUS

SEGMENTED OIL-TO-AIRHEAT EXCHANGER

OIL TANK

SIGHT GAGE

FILLER AND DIPSTICK

OIL-TO-FUELHEAT EXCHANGER

LRIGHT SIDE

CRASH LOGIC

26/28 VDC

A panel (Figure 7-10) in the rear compartmentfacilitates checking of the oil level. The panelcontains a quantity gage powered from a trans-mitter in each engine’s oil tank. Operating and

control power is supplied to the panel from thebattery bus through a circuit breaker, crash logicPC, and a microswitch, which closes when therear compartment door is opened. The gage iscalibrated in quarts and indicates oil to beadded to the selected tank. Power to the gage iscontrolled by an ON–OFF switch. Tank selection is made with a three-position

i h l b l d “ G

red. These inpostflight anthey must bchange) by u

located belomomentary mal (down)

Fuel HeaAn oil-to-fu(Figure 7-11




momentary contact switch labeled “ENG

3–ENG 2–ENG 1.” When the gage powerswitch is turned off, the sweep needle will goto a position labeled “OFF.” The oil quantityshould be checked within 10-60 minutes afterengine shutdown and serviced with the typeand brand specified in the AFM.

Pump

The oil pump (Figure 7-11) contains one pres-sure element and four scavenge elements. Itis installed on the aft face of the accessorygearbox and is driven by the accessory gear.

Regulator and Relief ValveA combination regulator and relief valve (Fig-

ure 7-11) controls and limits the pressure of the oil system. The regulator controls pressureat 42 +4 psi when N2 rpm is 68% or greater.If the pressure threshold is exceeded, the re-lief valve opens and limits pressure to the de-sign limit.

Filter

A filter (Figure 7-11) with a ∆P bypass is in-stalled in the pump pressure line. When the dif-ferential across the filter increases to between

gheat of the o

mation on thtion, see Fue

Oil CooliOil cooling exchanger a(oil cooler).

The oil-to-segmented hca ted in thincorporatessure bypass 65 and 77°Coil which passegments is engine rotorbearings, and

The coolingexchanger is

The lubricatiings and geabearings requit is directedchanger (Fiside of the ebly. This coo

valve (Figure 7-11). This valve begins to closebetween approximately 27,000 and 30,000feet to maintain an absolute engine case pres-sure of approximately 4 psi. This prevents oil

pump cavitation at higher altitudes.

Oil ScavengingA positive scavenge system is provided by thefour scavenge elements of the oil pump. Thescavenged oil is returned to the tank by a com-mon scavenge line. An indicating chip detector,

in the oil outlet lineexchanger and is trture scale on the aperature gage on th

(Figure 7-12). The brated in degrees Cefor operation.

OperationThe engine oil supptank (Figure 7-13),oil tank by the press




capable of attracting ferrous metal and fuzz,

is installed on the accessory gearcase; all scav-enge oil passes over the chip detector. Thecircuitry of the chip detector is in parallelwith the low oil pressure indicating system of the associated engine. It is described laterunder Indication.

Indication

Engine oil pressure is sensed at the outlet of the oil-to-fuel heat exchanger and transmittedto dual-scale, dual-needle pressure/tempera-ture gages on the engine instrument panel(Figure 7-12). The pressure scales are cali-brated in psi, and they require 28 VDC foroperation.

Oil pressure is also sensed by pressure switchesin each engine’s oil pressure line to providewarning of abnormally low pressure during op-eration. When oil system pressure is 25 psi orless, the appropriate OIL 1, OIL 2, or OIL 3light on the warning panel (See Appendix B)illuminates.

The OIL 1, OIL 2, and OIL 3 lights are wiredin parallel with the associated chip detector(Figure 7-12); therefore, if an accumulation of ferrous metal or fuzz occurs at a chip detec-tor, the appropriate low oil pressure warning

oil tank by the press

scavenge oil pump.and limited by a retransmitted throughThe oil from the heasegmented oil-to-airof this oil is directrotor bearings and ffer gear and the accing oil is further coo

fuel-to-oil heat exclubrication of the pscavenged by the fouand returned to thscavenge line. An atained in the lubricapressurizing valve of the pumps at hig

FUEL SYSTEMGeneral

Basically, the Thydromechanical flfuel to the combustrequirements.

The hydromechanicaacting alone or, nwith a digital electr


E E S W H E N R E A R

D O O R I S O P E N E D .

R D O O R

O S W I T C H

I N D I C A T I N G

P A N E L

( R E A R C O M P A R T M E N T )

L O G I C :

E A S T O N E O F T H E

E R A T O R O R B A T T E R Y

C



L O W - P R E S S

U R

S W I T C H

S W I T C H C L

O S

C O M P A R T M E N T

R E

A R

M I C

R O

B A T T E R Y B U S

W A R N I N

G P A N E L

O N A I R C R A F T W I T H O U

T T R A N S F E R

V A L V E X T K 2 , T H E W A R

N I N G P A N E L

D O E S N O T F E A T U R E X

T K 2 O P E N

O R C L O S E D L I G H T S .

T E M P E R A T U R E

R E S E T

S W I T C

H

C R A S H

A T

L E

G E

N E

P R E S S U R E

C R A S H L O

G I C

C I R C U I

T

2 6 / 2 8 V D C

O I L P R

E S S U R E

T R A N S

M I T T E R

C H I P

D E T E C T O

R

O I L 1

O I L 2

O I L 3

F O

VENT

BREATHERPRESSURIZING

VALVE

ANTISYPHONORIFICE

OIL TANK

LUBE ANDSCAVENGE OIL PUMPOIL PUMP

INLET

TRANSFER GASSEMBLY SC

NO. 4 AND NO. 5 B

TRANSFER GEASSEMBACCESSORY DRIVE

GEARBOX ASSEMBLY



7 - 1 3

O

R T RAI NI N G P URP O S E S

ONL Y

COMMONSCAVENGE

CHIPDETECTOR

∆ P INDICATOR

FUEL IN F

OIL TEMPERATU

FUEL HEAT

BYPASS VALVE

FILTER

REGULATORVALVE

FILTER INDICATOR

BYPASS VALVE

SUPPLY

PRESSURE

SCAVENGE

VENTING

FUEL

AMBIENT AIR

LEGEND

Figure 7-13. Oil Sy

The automatic mode is the desired mode underall operational conditions. When in the auto-matic mode, the hydromechanical fuel controlresponds to commands from the electronic fuel

computer and, in response to these commands,automatically provides engine start control,rpm governing, temperature limiting, acceler-ation and deceleration control, steady-state op-eration, and surge bleed valve operation.

In addition, the digital fuel computer incor-porates a full-time parameter monitoring sys-tem which will automatically produce reversion

the high pressure dshut off by the power

Anti-ice Sensor

Fuel temperature is filter. The sensor mevalve in the FCUtemperature decreasefuel is then directed exchanger to the inlwill close the antemperature increase




tem which will automatically produce reversion

to manual control if spurious signals exist.

Components

General

The engine fuel system components (Figure 7-14) include a two-stage pump, a relief valve, abypass filter, anti-ice valve and sensor, fuelheater, fuel control unit (FCU), flow divider, fuelmanifolds, duplex spray nozzles or atomizers,and a digital fuel computer.

Fuel Pump

The two-stage fuel pump consists of a certrifugallow-pressure (LP) stage and a positivedisplacement high-pressure (HP) stage. The LPstage receives inlet fuel at fuel tank boost pump

pressure. The inlet pressure is increased by theLP pump to between 30 and 40 psi.

Filter

A micronic filter (Figure 7-15) with a bypassremoves solid particles from the fuel. A ∆P of between 9 and 12 psi will open the bypass andpermit unfiltered fuel to enter the system. A

bypassing fuel filter is identified by its asso-ciated magnetic indicator. Filter outlet fuel is di-rected to the HP pump, which develops a

p

Fuel Heater

A fuel heater provchanging to prevensystem from cloggicomponents. Fuel fis thermostatically cof the fuel pump. Anprevents excessive p

FCU

The FCU is clampedand driven by the pshaft. The FCU inmechanical N2 goverultimate overspeed sovalve determines t

(relative to P3) to m(in the manual modeautomatic mode, a bmetered fuel to meetresponse to signacommanded by the fu

Mechanical Gov

The mechanical goflyweight-speeder spit controls N2 rpm dvaried by power lev

R ev i s i on 3

F O

FUEL MANIFOLDS

F R O

N T



7 - 1 5

O

R T RAI NI N G P URP O S E S

ONL Y

PUMPASSEMBLY

FILTER

FLOW DIVIDER

FUELCONTROL

UNIT

OIL/FUEL

HEAT EXCHANGER

PRIMARYSYSTEM

SECONDARYSYSTEM

FUEL MANIFOLD

RIGHT SIDE

FUELHEATER

11 1

2

3

4

567

8

9

10 7 3 1+

=

DIGITAL FUEL COMPU

Figure 7-14. Fuel Sy

7 - 1 6

F OR T RA

FUEL HEATER

FUE

PUMP

F I L T E R C O U P

FUEL FROMTANK



R ev i s i on 3

I

NI N G P URP O S E S

ON

L Y

11 1

2

3

45

67

8

9

10 7 3 1

0 0- 0K 0 3-B B 9 6- T5 0 9-N 10 1- T2 0 4- TM 0 7- MS 1 0- N20 2- BA 0 5-P L 0 8- OS 1 1-C C

+

=

FUNCTIONSELECTORSWITCH

AIR FILTER

TEST

DIGITAL FUEL COMPUTER

PT2TT2

(INLET AIR SENSOR)

INPUT/OUTPUT

INPUT28 VDC

IGNITION

SCREEN

TORQUE MOT

UP L I N G

TANK

Figure 7-15. Fuel

Fuel Shutoff Valve

A fuel shutoff valve consists of a rotary valveand a hydraulically (fuel pressure) operatedvalve. The rotary valve is operated by the powerlever and opens when the power lever is moved5° or more from the off position.

Ultimate Overspeed Solenoid Valve

This electrovalve is computer-controlled and isnormally closed. When energized open, it by-passes the high-pressure fuel from the piston of the fuel shutoff valve, allowing the spring to

The fuel systbe identified

Fuel ComGeneral

In normal opstart to shutby an associputer (Figurgines No. 1baggage com




the fuel shutoff valve, allowing the spring to

close the valve, and the engine flames out. Itsoperation is described later under Fuel Computer.

Flow Divider

The flow divider functions to direct metered fuel,initially to the primary fuel manifold and to the12 primary outlets in the duplex fuel nozzles inthe combustion chamber. When fuel flow demand

reaches approximately 150 pounds per hour, fuelis also directed into the secondary manifold, andsecondary ports in the duplex fuel nozzles.

baggage com

puter is locaoperating pofrom the 28-has a dual pbreakers eapower for op

Character

Computer optain engine pprogrammed

11 1

2

3

4

567

8

9

10

00-001-T02-B

PT2 PORT AIR FILTE

Computed ou tpu t s a r e de te rmined andtransmitted to a torque motor in the FCU toproduce the precise metered fuel to meet thepower requirement under all operating con-

ditions. In addition the computer maintains thedesign safe surge margin by controlling the LPcompressor surge bleed valve.

The following parameters are sensed by thecomputer:

• Power lever angle (PLA)

• PT2TT2

N1 is sensed by alocated at the aft eshaft. One winding inputs to the compu

rpm. The other winN1 indicator.

N2 is sensed by a simthe transfer gearcamonopole transmits representing N2 sposupplies inputs to the

ITT is sensed by ten




PT2TT2

• N1 (LP spool rpm)

• N2 (HP spool rpm)

• ITT (interturbine temperature)

• AB (LP compressor surge bleed area)

• Flight/ground signal

• Power increase signal

PLA is transmitted to the computer by a po-tentiometer on the FCU.

Total engine inlet pressure and total engine in-let temperature are sensed by the PT2TT2 probe(Figure 7-17) mounted in the nacelle inletduct and transmitted to the computer.

ples located betweeThe parallel circuicomputer representture (ITT). The comthe engine’s safe-sucompressor stall an

Figure 7-18 showbleed valve. Since

deenergized, the surbalanced in a 1/3-o

The computer is protrol as a function oThe ground idle prsecond N2 accelerflight idle program iacceleration rate to

computer, controllingidle to flight idle, isright main landing ge

The computer is repcreased power at higtakeoff conditions. M(N2 and ITT increasea PWR INC pushbutt

Computer Outpu

The computer functi


SOLENOID A

SOLENOID B

SURGE BLEED VALVECONTROL SOLENOIDS

F R O N

T



• Determine an acceleration schedule tomaintain the present surge margin or toprevent excessive ITT.

• Determine a deceleration rate to pre-

vent blowout, or control the decelerationrate to prevent excessive HP/LP rpmmismatch.

• Providing fun

• Provid

• Providtions.

• Termi

PT2

T2T

28 VDC

FROMAIRPLANESYSTEM

CONDITION: SOLENOIDS DEENERGIZED, VALVE 1/3 OPEN

SURGE BLEEDVALVE


;

;

11 1

2

3

45

678

9

10 7 3 1

0 0- 0K 0 3- BB 9 6- T5 0 9- N10 1- T2 0 4- TM 0 7- MS 1 0- N20 2- BA 0 5- PL 0 8- OS 1 1- CC

+

=

Figure 7-18. Surge Bleed Valve Op

• Monitor all computer functions, and ini-tiate fail-down to manual if voltages areout of range.

• Alternate overspeed (N1 and N2) protec-

tion in manual mode.

• Automatic transfer to N2 loop if T5 loopfails at maximum or cruise thrust.

• Open the surge bleed valve when thepower lever is at idle.

ControlEach engine fuel coassociated three-pothe ENGINES sectipanel (Figure 7-19)“CMPTR 1,” “CMPhave posit ions labe“OFF.” When the sengines are controllputer throughout the




INTENTIONALLY LEFT BLA




When the switches are at MAN, all computerfunctions, except N1 and N2 alternate ultimateoverspeed, are disabled. When the switchesare at OFF all computer functions are disabled.

NOTE

The computer switch must be posi-tioned to MAN when it is establishedtha t a computer mal funct ion i sadversely affecting engine operationor performance , even when theassociated computer fail light is not on.

The power incontrol push(See Appendengines, and

indicate modDepressing tfects a modeing a discretmodification23°C operati(0.7 percentabove 23°C 11.34 and 1



When computer failure or electricalpower failure occurs, the fuel systemautomatically reverts to manual mode.Acceleration time will increase to ap-proximately 10 to 25 seconds. Fuelconsumption will increase by ap-

proximately 5% for any given powersetting and idle rpm will be higher.

Indication

Three lights on the warning panel (Appendix B)are lableled “CMPTR 1,” “CMPTR 2,” and“CMPTR 3.” The appropriate light will be onwhen a computer switch is at the MAN or OFF

positions or if the computer switch is at theAUTO position and the computer has failed. If the failure results from transient conditions, thecomputer may be reset by moving the affectedswitch from AUTO to OFF and then back toAUTO. If the light persists, the switch must bemoved to the MAN position. This position pro-vides alternate ultimate overspeed control for theN1 spool at 107%, and at 109% for the N2 spool.

Figure 7-20 shows the interface between thefuel computer and the engine.

power leverproportionaposition to taheat indicatfor the TFE7the TFE731

Since the usethe detriment

on a digital copower is usesystem is equi

Monitoring

The digital etegral monito(built-in test e

The monitorfunctions, suresponds by

The BITE syself-testing identificatioin nonerasabmay be reca

sonnel when

Computer fa

CAUTION

F OR

PT2TT2



7 - 2 1


E S

ONL Y

FUEL CONTROL

MAX 120°

POWER LEVER

IDLE

20°

0°PLA

FUEL

IN

P3

LEGEND

AB = SURGE BLEED AREA PLA = POWER LE

N1 = LOW-PRESSURE SPOOL SPEED TM = DC TORQUE

N2 = HIGH-PRESSURE SPOOL SPEED OS = OVERSPEE

PT2 = ENGINE INLET TOTAL PRESSURE MM = MANUAL M

TT2 = ENGINE INLET TOTAL TEMPERATURE WF = FUEL FLOW

ITT = INTERSTAGE TURBINE TEMPERATURE P3 = COMPRESSO

Figure 7-20. Engine-C


POSITION/FUNCTION SWITCH REA

1—Normal run position Stores adjustment—no

2—Self-test Checks display lamps,detected faults

3—Fault diaplay Displays detected fault

4—Idle adjust Display indicates idle rp

5—Specific gravity Display indicates speciffrom 1-11.

6 Engine flat rate Display indicates flat ra

Table 7-1. COMPUTER MALFUNCTION ANALY



6—Engine flat rate Display indicates flat ra

7—FR/MN Display indicates % N2

8—T5 limiter Display indicates limite

9—T5 bias CL/CR Display indicates T5 in

10—Fault erase Displays 88, cal switcheflashes, displays 00, an

completed).

CODE COMPONENT

00 —

01 T2 sensor

02 Solenoid A

03 Solenoid B

04 Torque motor

05 Power lever pot

06 ITT

07 Manual mode solenoid

Table 7-2. CODE, AFFECTED PARAMETER, AND TYP

moving the affected computer switch to theMAN position will provide an alternate over-speed protection for N1 at 107% and for N2 at109%. The mechanical governor is still effec-

tive, and if operational, it will limit N2 to 105%in both automatic and manual operating modes.

Computer Interrogation

The computer may be interrogated for identifi-cation of stored faults or for testing and adjust-ment. When the rotary switch is placed in anyposition from 1 to 10, an LCD (liquid crystal dis-

play) on the face of the computer (Figure 7 16)ill di l di i l d l i h l i

The engine is assucruise power settingis at the AUTO povalve (1) is energ

mechanical N2 goveThe power lever is ad(2) is open, directinshutoff valve (3). Thepressure regulator ful(4). The computer omotor (5) determine tpass valve (6) relativtablishing the mete

meteringvalve(4) to th




play) on the face of the computer (Figure 7-16)will display a digital code relative to the selection.

Table 7-1 shows computer malfunction analy-sis, relative to the function selector switchposition. This switch must be returned toPosition 1; if not, the affected computer faillight will continue to flash on and off regard-less of computer switch position.

Table 7-2 shows the code, the affected parame-ter, and the type of fault rela tive to each posi-tion of the function selector switch.

NOTE

Position 11 is an unused position onFalcon 900 airplanes.

Fuel Enrichment

The computer is programmed to provide auto-matic fuel enrichment during engine starting. Thisassists engine acceleration and also enhances thenozzle spray pattern for more efficient combustion.

The automatic fuel enrichment increases me-tered fuel volume by approximately five ratios.

It is modulated by T5 and automatically ter-minated when ITT reaches 200°C. As the en-gine continues to accelerate toward idle rpm,

metering valve (4) to thfrom there through thprimary fuel manifoldvider to the secondarymanifolds supply the the combustion chamautomatically maintaprevent fuel filter icinice sensor (11) and th

When an N2 overspeand the mechanical to the governor recomputer will energsolenoid (13); at eiththe fuel shutoff valvwill flame out.

Figure 7-22 shows

in the manual modeIt is assumed that the power. The associatether at MAN or OFFvalve (1) is deenergizN2 governor will contlever position by vatmosphere (2). Thetioned relative to preferential to meterrequirement. The excthe inlet of the HP p

R ev i s i on 3

F OR

T

6

FROM TANKBOOST PUMP

BYPASS6.5 PSID

ANTI-ICESENSOR

HPPUMP

FILTER40-MICRON

METAL

INLET

FILTER BELLO

11



7 - 2 3

TRAI NI N G P URP O S

E S

ONL Y

Figure 7-21. Fuel System Op

11 1

2

3

45

678

9

10 7 3 1+

=

TEST

28 VDC

LPPUMP35–40PSID

RELIEF1,450 PSIG

CMPTR

AUTOMANOFF OIL FLOW


PT2TT2

FLT

GRND

FUEL/OILHEATER

ANTI-ICEVALVE

BYPASSVALVE

SPEED

GOVERNOR(105% AUTO)

OVERBOARD

DRAIN

DRIVE SHAFT

TOPOWER LEVERPOTENTIOMETER

MANUAL MODPISTON

MANAD

LEGEND

BOOST PUMP PRESSURE

LP PUMP PRESSUREHP PUMP PRESSURE

REGULATED PRESSURE

SECONDARY NOZZLE FUELP

3 AIR PRESSURE

12

7 - 2 4

F OR T RAI N

I N

4

FROM TANK

BOOST PUMPBYPASS6.5 PSID

ANTI-ICESENSOR

HPPUMP

FILTER40-MICRON

METAL

LPPUMP

RELIEF1,450 PSIG

INLET

FILTER BEL

5



R ev i s i on 3

N G P URP O S E S

ONL Y

2

11 1

2

3

45

678

9

10 7 3 1 +

=

TEST

28 VDC 35–40PSID

CMPTR

AUTOMANOFF OIL FLOW

DIGITAL FUEL COMPUTERPT2TT2

FLT

GRND

FUEL/OILHEATER

ANTI-ICEVALVE

BYPAVALV

SPEEDGOVERNOR(105% AUTO)

OVERBOARD

DRAIN

DRIVE SHAFT

TOPOWER LEVERPOTENTIOMETER

MANUAL MOPISTON

MA

LEGEND

BOOST PUMP PRESSURE

LP PUMP PRESSURE

HP PUMP PRESSURE

REGULATED PRESSURE

SECONDARY NOZZLE FUEL

P3 AIR PRESSURE

Figure 7-22. Fuel System O

are valid. Normally the mechanical governorshould limit N2 rpm relative to power lever po-sition or to 105% N2 when an overspeed occurs.

The P3 limiter (6) functions in all operating

modes to limit the ultimate value of P3, therebyaffording structural protection to thecompressor casing.

IGNITION SYSTEM

GeneralThe Falcon 900 uses a high-energy ignitionsystem consisting of a dual-ignition exciter,two high-tension cables, and two igniter plugs.

The ignitiomodes: grou

Control

The ignition tor switchesoverhead swswitches arSTART,” “Mfor the switposition powMAL” and NORMAL pbatteries, AP




two high tension cables, and two igniter plugs.

Figure 7-23 shows the locations of the com-ponents of the ignition system.

batteries, APcombinationposition, theThe normal p

DUAL IGNITION UNIT

IGNITION BOX

START. When in this position, the ignition op-erates when the associated power lever ismoved from the stop to the idle position.Ignition will continue until terminated by anelectronic speed switch in the fuel computerat approximately 50% N2. Ignition may be ter-minated up to 50% at any time by moving theswitch to the MOTOR–START STOP position.

The AIR START position bypasses the powerlever microswitch and provides continuous ig-nition on the ground or in flight. Normally thisposition is used only for airstarts or ground

checks, (or at other times as outlined in the AFM

).

when the switch is inthe power lever micputer are bypassedignition. The stall ithe autoignition relinformation on stalControls” chapter i

INSTRUMENT

GeneralThe engine-related

in three vertical roment panel (Figure 7




, ( )The MOTOR–START STOP pos it ion isprimari ly used to dry motor the enginewithout fuel and ignition, for maintenancepurposes, or for clearing the engine followingan aborted ground start. This position may alsobe used to terminate a ground start prior toreaching 50% N2. It must be used to terminatestarting and ignition whenever manual modestarts are made.

The stall ignition system is an automatic an-tiflameout system, which is armed for opera-tion whenever the switch is in the GRD STARTposition. For more information on the stallignition system, see the “Flight Controls”chapter in this training manual.

IndicationAmber rectangular lights labeled “IGN 1,”“IGN 2,” and “IGN 3,” are below the start se-lector switches (Appendix B and Figure 7-23).The associated light will be on whenever poweris available to the ignition exciter. These lightsare not an indication that the ignition exciteror the igniter plugs are operational.

Operation

ment panel (Figure 7a typical layout con

• N1 (LP spool

• ITT (interturb

• N2 (HP spool

• Fuel flow• Fuel quantity

• Oil pressure/t

N1 (LP Spool R

The N1 gage is analo

is calibrated in percefrom 0 to 110%. A lected rpm. A three-decimal place is locThe digital indicatpercent. The N1 gaindicator and is usedN1 gage receives imonopole located spool shaft . Onesupplies the N1 inppreset compensatorN1 inputs to the fue

ITT (Interturbine Temperature)The ITT gage is analog/digital. The analogscale is calibrated in degrees Celsius from 0 to1,000, and a sweep needle indicates the set

value. A four-digit display provides instanta-neous ITT indication. The digital indicator islocated on the dial face. A red light on the dialface will flash on and off when ITT indica-tion exceeds 952°C (or 974°C when the powerincrease system is used for the TFE731-5AR-1C) and 980°C (or 996° C when the power in-crease system is used for the TFE731-5BR-1C).

ITT sensingcuits, connecumferentiaLP turbines.informationond circuit sufuel compupower for opor fails, the sbelow 0°C an




IDLE GENSW

AUTO-IGNITIONCUTOFF

LOGICGATE

TOSTARTER

START RELAY

MOTORSTARTSTOP

CMPTR

50% N2

IGNITER PL

IGNITIONRELAY

IGNITIOEXCITE

GRDSTART

AIRSTART

G C U

for operation; when psweep needle will go

Fuel FlowSee Fuel System, th

Fuel Quantity

See Fuel System, th

Oil Pressure/T

See “Oil System,” t




Figure 7-26 depicts the operation of the ITTthermocouple system.

N (HP Spool RPM)

Engine Instrum

A test system is proN2 gages. All three aa push-button switcnance test panel (Fig

sidewall.

The test switch is laThe test system indharness, signal procdisplays, and the N1test requires approxthe test switch is mosweep needle will m

at 106% and hold stwill initially indica106%, and finally cvalue of the N1 covalue must not be gre0% of actual N1. Thto the blue dot on 1,000°C for the TFEfor the TFE731-5BR

will show 8888 and tred light on the dial

Figure 7-25. Engine Instrument Panel


TO ITTGAGE

TO FUELCOMPUTER



LEGEND

CIRCUIT 1 CIRCUIT 2

10 9 8 7 6 5

Figure 7-26. ITT Thermocouple System

(Figure 7-28). The power lever is mechanicallyconnected to the input power lever on the FuelControl Uni t (FCU). The power levermechanically sets the FCU governor (primar-ily for manual mode operation) to maintain N

2relative to power lever position, and simulta-neous ly pos i t ions a wiper on the FCUpotentiometer, which transmits PLA informa-tion to the fuel computer for automatic mode

operation. The powpositions: stop, idlelatch on the power lthe lever can be movepositions. An adjustat the full power po




Reverser LeverA thrust reverser control lever is piggyback-mounted on the No. 2 power lever.

Friction ControlPower lever friction is preset by maintenance.There is no crew adjustment.

Quadrant Proximity Switchesand Solenoids

Proximity switches are mounted in the powerlever quadrant to perform a variety of functions

Ground SEngine crankthe starter-ge

There is no starting, bu2–3–1 withavailable. Wthe recomme

Control

All switches located on the7 19) Electri




during power lever travel. These proximityswitches are identified as follows:

• Thrust reverser arming (No. 2 engine)

• Engine fail warning (No. 2 engine)

• Thrust reverser control (No. 2 engine)

• Air-conditioning cutoff

• Stop/s tart

• Landing gear warning

• Engine warnings (T/O CONFIG)

In addition, the power lever quadrant containstwo electromagnetic lock solenoids. One sole-noid locks the reverser lever at the deployposition until the reverser doors reach thefully deployed position, and the other solenoidlocks the power lever in the idle position whenthe reverser is out of the stow position.

ENGINE STARTING

General

Engine starting is divided into two generalcategories: ground start and airstart.

7-19). Electridescribed inchapter. Enguntil all safecrepancies arare outlined iUnder all nomust be in the

The followinplified descrivarious powecycle. It is assare positione

Battery S

Push and momswitch; checgage to a minrotation. Theidle positioncomes on. Chmonitor ITTand check tha

tor all engineidle is achiev



R ev i s i on4

F OR

T RA

NORM

EXT POWER

OIL

APU STOP

START


GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

GEN 3GEN

G1

GENERATOR 1

TOAMMETER

G3

GENERATOR 3

STARTCONTACTOR

TO AMMETER

DC SYSTEM

OVERHEAD PANEL

CONDITIONS:BUS-TIE CLOSED,

POWER SELECTOR NORMAL,

BAT 2

GEN 2 APU

BAT 1

GEN 3GEN 1



7 - 3 3

AI NI N G P URP O S

E S

ONL Y

G2

GENERATOR 2

TO

AMMETER

G3

G

START CONTACTOR

BATTERY 1CONTACTOR

BATTERY 1

STARTCONTACTOR

GPU CONTACTOR

APU

TOAMMETER

TOAMMETER

FIREEXTINGUISHERS

COCKPITDOME LIGHT

TO

AMMETER

GPU RECEPTACLE

BATTERYBUS

STARTCONTACTOR

BATTERY 2 BATTERY 2CONTACTOR

STARTINGBUS

BOTH BATTERY SWITCHES ON,START BUTTON 3 PUSHED

Figure 7-29. Battery Start—N

7 - 3 4

F OR T RAI N

I N G

NORM

EXT POWER

OIL

APU STOP

START


GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

GEN 3GEN

G1

GENERATOR 1

TOAMMETER

G3

GENERATOR 3

STARTCONTACTOR

TO AMMETER

START CONTACTOR

BATTERY 1BATTERY 1

DC SYSTEM

OVERHEAD PANEL

CONDITIONS:

BOTH BATTERY SWITCHES ON,BUS-TIE CLOSED,POWER SELECTOR SWITCH NORMAL,APU GENERATOR ON,

NO. 2 ENGINE START

BAT 2

GEN 2 APU

BAT 1

GEN 3GEN 1



R ev i s i on 3

P URP O S E S

ONL

Y

G2

GENERATOR 2

TOAMMETER

G

BATTERY 1CONTACTOR

STARTCONTACTOR

GPU CONTACTOR

APU

TOAMMETER

TOAMMETER

FIREEXTINGUISHERS

COCKPIT

DOME LIGHT

TOAMMETER

GPU RECEPTACLE

BATTERYBUS

STARTCONTACTOR

BATTERY 2 BATTERY 2

CONTACTOR

STARTING

BUS

Figure 7-30. Battery/APU Generator-Assis

R ev i s i on4 . 0 2

F OR

T RA

NORM

EXT POWER

OIL

APU STOP

START


GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

GEN 3GEN

G1

GENERATOR 1

TOAMMETER

G3

GENERATOR 3

STARTCONTACTOR

TO AMMETER

DC SYSTEM

OVERHEAD PANEL

CONDITIONS:

BOTH BATTERY SWITCHES ON,BUS-TIE CLOSED,POWER SELECTOR SWITCH NORMAL,APU OPERATING,

BAT 2

GEN 2 APU

BAT 1

GEN 3GEN 1



7 - 3 5


ONL Y

G2

GENERATOR 2

TOAMMETER

G

START CONTACTOR

BATTERY 1CONTACTOR

BATTERY 1

STARTCONTACTOR

GPU CONTACTOR

APU

TOAMMETER

TOAMMETER

FIREEXTINGUISHERS

COCKPIT

DOME LIGHT

TOAMMETER

GPU RECEPTACLE

BATTERYBUS

STARTCONTACTOR

BATTERY 2 BATTERY 2

CONTACTOR

STARTINGBUS

NO. 2 GENERATOR ON,

NO. 3 START

Figure 7-31. Battery/APU Generat(RPM < 50% N2) (No.

its GEN light will go out. All three generatorswill operate in parallel since the bus tie isclosed. Battery 1 and battery 2 will reconnectto their associated main buses, the batterylights will go off, and charging will recur unti l

a start is initiated on the No. 1 engine.

Figure 7-32 depicts what occurs when the No.1 engine START switch is pushed. The battery1 and battery 2 make-and-break switches willopen (the BAT 1 and BAT 2 lights will come on).Simultaneously the APU generator reversecurrent relay will open, (the amber APU GENlight will come on) and disconnect the APUgenerator from the distribution buses. The bat-tery 2 contactor will close and parallel battery

ground for thematically closi

GPU Start No.An approved GPU mgine when required.the unit must be checoperating. For more control and protectioSystems” chapter in

Figure 7-33 shows and operating. Moswitch from NORMclose the GPU con




2 with battery 1 on the start bus. The APU startcontactor will close and connect the APUgenerator to the start bus. Finally, the No. 1 en-gine start contactor will close, connecting thestart bus to the windings of generator 1, andinitiate the start cycle. Generators 2 and 3 in par-allel will supply the distribution buses. Fromhere on, the start cycle is identical with that de-scribed for a battery start of the No. 2 engine.

NOTE

When a generator-off light remains onafter the start terminates, it may indi-cate that the associated start contactorfailed to open, and the appropriate ac-tion must be taken by the crew.

When the No. 1 engine start terminates, gener-ator 1 will connect to the left main bus, and theGEN 1 light will go off. The APU generator willreconnect to the right main bus, and the APU GENlight will go off. Battery 1 and battery 2 will re-connect to their respective buses, and the BAT 1and BAT 2 lights will go off. The BUS TIEDlight will remain on until the bus-tie switch isrotated to the FLIGHT NORMAL position.

close the GPU conopen the battery 1 coand-break switch, dfrom the main and battery-off lights wsupplied to the bus-the bus-tie switch

position. The GPU,operation of all engigenerator to preparalleled with the connects to the stacontactor, to the lefbattery 1 make-andright main bus thrPushing the No. 2

the No. 2 start contaof the GPU to thstarter-generator. Fsequence is identicbattery start of the N

All generator, battewill remain on as nected and operatin

Engine Start L

R ev i s i on 3

F OR

T RAI

NORM

EXT POWER

OIL

APU STOP

START


GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

GEN 3

GEN

G1

GENERATOR 1

TO

AMMETER

G3

GENERATOR 3

START

CONTACTOR

TO AMMETER

DC SYSTEM

OVERHEAD PANEL

CONDITIONS:

BOTH BATTERY SWITCHES ON,BUS-TIE CLOSED,POWER SELECTOR NORMAL,APU OPERATING,

O G O O

BAT 2

GEN 2 APU

BAT 1

GEN 3GEN 1



7 - 3 7

NI N G P URP O S E S ONL Y

G2

GENERATOR 2

TOAMMETER

G

START CONTACTOR

BATTERY 1

CONTACTOR

BATTERY 1

STARTCONTACTOR

GPU CONTACTOR

APU

TOAMMETER

TOAMMETER

FIREEXTINGUISHERS

COCKPITDOME LIGHT

TOAMMETER

GPU RECEPTACLE

BATTERYBUS

STARTCONTACTOR


STARTINGBUS

NO. 2 GENERATOR ON,NO. 3 START

Figure 7-32. Battery/APU Generat(RPM < 50% N2) (No.

7 - 3 8

F OR T RAI N

I N G P

NORM

EXT POWER

OIL

APU STOP

START


GEN 1 BAT 1 BAT 2

FLIGHT NORM

GEN 2

APU

GEN 3

GEN

G1

GENERATOR 1

TO

AMMETER

G3

GENERATOR 3

STARTCONTACTOR

TO AMMETER

START CONTACTOR

BATTERY 1CONTACTOR

BATTERY 1

DC SYSTEM

OVERHEAD PANEL

CONDITIONS:BATTERY AND GENERATOR SWITCHES ON,APU INOPERATIVE,BUS-TIE SWITCH—FLIGHT NORMAL,

POWER SELECTOR SWITCH —EXT POWER

BAT 2

GEN 2 APU

BAT 1

GEN 3GEN 1



R ev i s i on 3

P URP O S E S

ONL

Y

G2

GENERATOR 2

TOAMMETER

G

CONTACTOR

STARTCONTACTOR

GPU CONTACTOR

APU

TO

AMMETER

TOAMMETER

FIREEXTINGUISHERS

COCKPITDOME LIGHT

TOAMMETER

GPU RECEPTACLE

BATTERYBUS

STARTCONTACTOR


STARTINGBUS

Figure 7-33. GPU Start—No.

It does not apply to the GPU generator, sincethe selection of EXT POWER by the powerselector switch inhibits the APU generatorand all engine generators until the sw itch isreturned to the NORMAL position.

AirstartsAirstarts consist of starter-assisted starts andwindmilling starts. All airstarts must be per-formed within the airstart envelope (Figure 7-34) con ta ined in Sec t ion 3 , “AbnormalProcedures,” of the AFM .

Starter-As

Starter-assiswhenever thand N1 rota

checklist muThe procedidentical wbattery starts

Windmillin

An immediaand high-spe

any altitudedepicted in tthat the cau

WARNING




An engine must not be airstarted un-less the cause of flameout or thereason for shutdown is determined asnot be ing caused by f i re or me-chanical failure.

caused by mcomputer is

To perform power lever tstart selector

that the IGNwithin 10 se

;

;

;

;

;

;

;

;

;

ALTITUDE (X 1,000 FT)

30

20

10

MAXIMUM AIRSTART ALTITUDE IN MANUAL MODE

instruments until the engine stabilizes at idle.Turn the start selector switch back to GRDSTART, and set power on the engine as required.

Windmilling Airstart (N2 15% orGreater and N1 Greater Than 10%)

To perform an airstart when windmilling rpmis 15% or greater is identical to the procedurefor a high speed airstart, except that when N2reaches 50%, the start selector switch is movedback to the GRD START position.

NOTEWait 10 seconds between two con-secutive airstarts and do not perform

door in the engine not closed.

This engine-fail motional only when the

A failure sensing unnose gear proximityerated when No. 2 eor higher and No. 2 e

The No. 2 engine fain flight if the enginot closed.

Indication

A light labeled “EN




secutive airstarts and do not performm or e t ha n t h r e e uns uc c e s s f u lairstarts.

Aborting Airstarts

An airstart must be aborted when any of the

following conditions are present:

• No oil pressure within 10 seconds oflight-off

• No ITT within 10 seconds after movingthe power lever to idle

• Rapid ITT increase (approaching 952°C forthe TFE731-5AR-1C, or 978°C for the

TFE731-5BR-1C)

• N 1 r e m a i n i n g c l o se t o z e r o w h e nN2 = 20%

• N2 not increasing rapidly and smoothlyafter light-off

If the IGN light remains on after moving the startselector switch to GRD START, move the switch

to MOTOR–START STOP, and check that thelight goes off.

A light labeled ENlocated on the pilocome on when anscribed conditions

THRUST REV

General

A hydraulically opeverser is mounted duct of the No. 2 enfair with the aft enreverser is in the st35). When deploye

a vertical axis (Figu2 engine exhaust athrust outward and fceleration of the air

The reverser controthe main landing getrolled by the touchconsequently, the

tional only with we

ControlThe reverser(Figure 7-28power lever.

positions: stThe lever is position untat idle. The istow positiosolenoid, peverse thrust ofully deploy

Figure 7-37 dhydraulic de




When the revposition, a mpletes a circreverser selsolenoid, whlocking actu

verser actuathe latchessolenoid is valve in the rdraulic pressploy the clamthe fully depclosed, and locking sole

reverser leveWhile the resition the poand mechani

Moving the rreverse thrusreached. Thelocked in th

verser lever tspool valve tof the revers

Figure 7-35. Thrust Reverser—Stowed

Pressure SourceThe thrust reverser is powered from the No. 2hydraulic system. The thrust reverser subsys-tem includes a pressure storage accumulator,

charged through a check valve. The accumu-lator stores fluid under nitrogen pressure andis capable of one complete reverser cycle whenNo. 2 system pressure is not available.

Indication

Two lights (see Annu7-25) provide indicoperation. One light

SIT,” will illuminatlock or clamshell ddeployment and reming bellcrank controlthe fully deployed p


NO. 2 SYSTEM

CHECK VALVE



LEGEND

NO. 2 SYSTEM PRESSURE

NITROGEN

RETURN

HOSES

CUTOFFVALVES

DEPLOYMENT SOLENOID

FLIGHTDEPLOYMENT SAFE

ISOLATION VATHERMAL

EXPANSIONVALVE

E2 E3

ACCUMULATOR

LOCKINGSOLENOID

The second l ight (green) , labeled “DE-PLOYED,” will come on when a synchroniz-ing bellcrank controll ing the clamshellsreaches the fully deployed position and re-mains on. The operating sequence of these

two lights indicates normal operation of the re-verser. During a stow cycle the DEPLOYEDlight will go out when the clamshells leave thedeployed position, and the TRANSIT lightwill come on. The TRANSIT light will go off when the clamshells are stowed and locked. Alight labeled “REV UNLOCK” on the warn-ing panel will come on only during a retrac-tion cycle and go off when the retraction cycle

is complete. This light will also come on duringflight if either one of the clamshell lock latchesis not in the locked position.

There iduring may be

i s no dingestio

Any cycthe grodown, clamsh

cated bUNLOC




p

Automatic Stowing

The thrust reverser incorporates an automaticstow feature which will automatically apply

and maintain hydraulic pressure on the stowside of the reverser actuator when any one of the clamshell door latches moves from thelocked position.

Emergency StowAn emergency stow system is controlled by aguarded two-position switch on the engine in-

strument panel (Figure 7-25). The switch is la-be led “STOW” and “NORMAL” and i sguarded in the NORMAL position. When theswitch is moved to the STOW position, the cir-cuit will bypass all normal sequencing andapply and maintain hydraulic pressure to thestow side of the reverser actuator.

Inoperative ReverserFlight is authorized with an inoperative re-verser provided that it is locked in the stow

LIMITA

This airplanwith the lim

the AFM , reg

The followinthe AFM .

THRUST (UNINSTALEVEL, I

Takeoff

TFE731-5AR

TFE731-5BR

Maximum

TFE731-5AR

TFE731-5BR

THRUST SETTING

The engine low-pressure rotor speed N1 isused as the thrust setting parameter.

The takeoff and maximum continuous thrustmust be based on the N1 values in Section 5 of the AFM:

• Takeoff thrust (5-minutetime limit) ............ Refer to Section 5

of the AFM • Maximum

continuous ............ Refer to Section 5

of the AFM

MAXIMUM ENGINE ROTORSPEED (N AND N )

Takeoff—With incre(5 minutes maximum

TFE731-5AR-1C ........TFE731-5BR-1C ........

STARTING TIM

Windmilling airstaN2 to 60% N2) .... . .

Ground start (from to idle) .................

FUEL COMPUThe fuel computers off Flight with one




SPEED (N1 AND N2)

CONDITION OF USE N1 N2

Takeoff and maximumcontinuous ........................ 100% 100%

N1 N2

Takeoff (With increasedthrust)...................................... 100% 100.8%

Transient (5-second maxi-mum allowable)...................... 103% 103%

100% N1 = 21,000 rpm—TFE731-5AR-1CTFE731-5BR-1C

100% N2 = 29,692 rpm—TFE731-5AR-1C

100% N2 = 30,300 rpm—TFE731-5BR-1C

With increased thrust100.8% N2 = 29,989 rpm—TFE731-5AR-1C

100.8% N2 = 30,540 rpm—TFE731-5BR-1C

MAXIMUM INTERSTAGETURBINE TEMPERATURE (ITT)

S i d/ i l

off. Flight with one may be permitted iAnnex section of th

THRUST REV

The thrust reverseruse only.

APPROVED O

Type II oils: AeroExxon 2380, Mobilacccordance with Company: EMS 53

These brands may

OIL PRESSUR

THRUST MIN

S ET TI NG P RE

Takeoff ormaximumcontinuous ....... . ..

Idl

NOTE

The OIL 1, OIL 2, and OIL 3 lightsin the warning panel illuminate foran oil pressure below 25 psi or chip

detector contamination.

OIL TEMPERATURE

Operational limits:

From sea levelto 30,000 feet ........................ 127°C maximum

Above 30,000 feet ................. 140°C maximum

Transient at anyaltitude.................................. 149°C maximum

(2 minutes)




.............................................. (2 minutes)

Minimum(for exceeding idle power) ................... 30°C

PLACARDS

AND INSTRUMENTSAll placards and instrument markings required byFAR 25 are contained in the Operating Manual.


1. The TFE731 engine is defined as a:

A. Twin-spool aft fan

B. Twin-spool turbofan

C. Free-spool turbofan

D. Single-spool turbofan

2. N1 or N2 transients of 103% are accept-able for:

A. Continuous operation

B. Five minutesC . One minute

D. Five seconds

A. Fuel compposition

B. N os e ge a rfailed

C. Engine powdoor is not

D. Fuel comp

7. The takeoff ththe TFE731 en

A. Fuel f low

B. N1

C ITT

QUESTIONS



3 . If the No. 2 engine fuel compu ter failswhen lined up for takeoff:

A. Do not take off and have theproblem corrected.

B. Match N1 to that of the highestengine.

C. Match ITT to that of the lowestengine.

D. Set takeoff power to the N1 or ITTlimit.

4. The minimum oil pressure to sustainfl ight is :

A. 55 psi

B. 25 psi

C. 38 psi

D. 46 psi

5. The maximum oil temperature to sus-

tain flight above 30,000 feet is:A. 140°C

C. ITT

D. N2

8. The maximum10% N2 to ligh

A. 60 secondsB . 20 seconds

C . 45 seconds

D. 10 seconds

9. The fuel comreset if it trips

A. Input volta

B. N2 input fa

C. PT2TT2 inp

D. Momentarthrust

10 . When the fuel MAN posi t io

tored i s :A. T5

d


CHAPTER 8

FIRE PROTECTIO

CONTENTS

INTRODUCTION......................................................................

GENERAL ................................................................................

FIRE DETECTION PANEL .....................................................

FIRE DETECTION SYSTEMS ...............................................



Engine Fire and Overheat Detection .................................

APU Fire Detection ...........................................................

Baggage Compartment Fire Detection .............................

Landing Gear Wheel Well Overheat Detection .................

Fire Test ..............................................................................

FIRE EXTINGUISHING .........................................................

General................................................................................

Engine Fire Extinguishing..................................................

APU and Baggage Compartment Fire Extinguishing ........

PORTABLE FIRE EXTINGUISHERS ....................................

Automatic Fire Extinguishers (Option 26-20-10)..............

QUESTIONS..............................................................................


ILLUSTRATIONS

Figure Title

8-1 Fire Panel and Warning Lights...................................

8-2 Continuous-Loop Fire Detectors ...............................

8-3 Fire Detection System Block Diagram .....................

8-4 Fire Test Circuit Block Diagram ...............................

8-5 Fire-Extinguisher Bottles ...........................................

8-6 Fire-Extinguishing System.........................................

8-7 Extinguishing System Controls and Operation



8-7 Extinguishing System Controls and Operation .........

8-8 Identification of Discharge Head Connections ..........

8-9 Halon Extinguisher (Typical).....................................

TABLE

Table Title

8-1 Fire Detector Characteristics .....................................


FIRE

; ; ; ; ;

; ; ; ; ;

; ; ; ; ;

; ; ; ; ;

; ; ; ; ;

; ; ; ; ;

; ; ; ; ;

; ; ; ;

; ; ; ;

; ; ; ;

; ; ; ;

; ; ; ;

;;

;;

;;

;;

; ;

; ;

; ; ; ;

; ; ; ;

; ; ; ;

; ; ;

; ; ;

; ; ;

; ; ;

; ; ;

FIR

FW



; ;

; ;

; ;

; ;

; ;

; ;

; ;

; ;

; ;

; ;

; ; ; ; ; ;

; ; ; ; ; ;

; ; ; ; ; ;

; ; ; ; ; ;

; ; ; ; ; ;

INTRODUCTION

The Falcon 900 has a fire protection system to detect and tions of the airplane. The detection system activates warn

the cockpit. The extinguishers are controlled by crew co

GENERAL

The crew is warned of the presence of fire,overheating, or smoke (in the baggage com-partment) by detection units. For each zone the

fire panel or warning lights illuminate. Anaudio warning sounds when overheat, fire, or

k i t i it d th th th

supply line iated shutoff

The distribuplied by fixeth


Each of the main landing gear wheel wells isprovided with an overheat detection system.The brakes are a potential source of heat whichmay lead to excessively high temperatureswhen the gear is retracted and the doors locked

closed. The gear wells are not equipped withan extinguishing system.

A test circuit provides for detection and ex-tinguishing system monitoring.

FIRE DETECTION

PANELThe fire panel (Figure 8-1) is equipped withextinguisher discharge control switches, lo-

There are five red ffire warning panel.minates if a fire is dthe area of the APUbaggage compartme

illuminated until tpast. Light illuminan aural warning, wthe HORN SIL pus

The applicable FAUfault occurs in thecircuit. The FUELfor each engine) con

each engine feed linor closing is verifilight going out whtion is completed. I



extinguisher discharge control switches, located above the corresponding fire warninglight (FIRE 1, FIRE 2, and FIRE 3, FIRE APU,and FIRE BAG COMP) (all colored lights areshown in Appendix B).

T h e t h r e e e n g i n e e x t i n g u i sh e r D I S CHswitches have three positions, labeled “0,”“1,” and “2.” The APU and baggage com-partment discharge switches have two posi-tions, labeled “0” and “1.”

Normally, all of the DISCH switches are safe-tied with break-away wire to the 0 position.The 1 position of the engine DISCH switches

is powered by the main bus as follows: No. 1engine, bus A1; No. 2 engine, bus B1; No. 3engine, bus A2. The 2 position is powered bythe battery bus. The switches must be pulledout prior to upward movement to position 2.The baggage compartment discharge switch ispowered by bus A2 and the battery bus; theAPU discharge switch is powered by bus B1and the battery bus.

The APU discharge switch, when moved to

pamber FAULT lightitoring the loop. If fault is indicated in

FIRE DETESYSTEMS

ENGINE FIREDETECTION

The continuous-loo

fine-bore capillary depending on wherdetectors are permasponder, and the comically sealed. The metal core saturatemaintains pressure is the sensitive elemfunctions in two mo

• General oveture detection


FAULT

210

TRANS

FIRE 1

DISCH

FAULT

210

TRANS

FIRE 2

FUELSHUT

OFF

DISCH

FAULT

210

FIR

DIS HDISCH

WARNING PANEL

TEST

FIRELIGHTS

FUELSHUT

OFF

FUEL SHUTOFFLIGHTS (3)



HORN SILPUSHBUTTON

FIRELIGHTS

L PITOTL PITOT R PITOTR PITOT

R AOAR AOA

HOTBATHOTBAT

FLAPASYMFLAPASYM

R. WHLOVHTR. WHLOVHTL. WHLOVHTL. WHLOVHT

L AOAL AOA

ST BYPITOTST BYPITOT

UTO

SL TS

AUTOSLATS

• Local overheating (pinpoint detection)—If a sudden local temperature rise oc-curs (actual flames), the metallic corereleases a high quantity of gas to imme-diately activate the pressure switch. This

process is reversible, allowing the de-tectors to function repeatedly. As thetemperature returns to normal, the gas isreabsorbed by the metallic core, and thedetector returns to its initial state of readiness.

Figure 8-2 shows the location of engine andAPU fire detectors.

The detector unit is a sealed responder con-taining two pressure switches:

• Pressure warning switch—The contactsclose at the occurrence of its preset gen

• Integrity presare normallypressure in th

When the detector detected, the applicpanel illuminatessounds.

If the detector unit icontacts are open. Ilight of the corresporeveals such a defethe same type and dthe valves for whicare preset (Table 8-




close at the occurrence of its preset gen-eral or localized overheating conditions.

NO. 3 ENGINE

NO.APU


Table 8-1. FIRE DETECTOR CHARACT

NUMBER OF AMBIENTLOCATION DETECTORS DETECTION THRESHOLD DETEC

ENGINE One per engine 400°F (204°C) 9

FIRE ZONE

APU One 400°F (204°C) 9

LANDING GEAR One per 270°F (132°C) 8WHEEL WELLS wheel well(SNs 001–095)

LANDING GEAR One per 3WHEEL WELLS wheel well(SNs 096–999)

APU FIRE DETECTION warning circli h



The fire detector in the APU compartment isthe same type used in the engine nacelles. Itoperates in two modes: general overheat andlocal overheat detection. Detection of an over-

heat condition illuminates the red FIRE APUlight and sounds the aural warning.

BAGGAGE COMPARTMENTFIRE DETECTION

Fire protection in the baggage compartment isprovided by an optical smoke detector utiliz-ing light reflection from smoke to excite a

photosensitive element.

When a smoke concentration of 15 to 30% pen-etrates the optical detector, the illumination of a photoelectric cell is modified, triggering awarning signal on an electronic circuit. TheFIRE BAG COMP warning light will illuminatewith a response time of less than four seconds.

To ensure that the warning signal is providedwithout delay, the smoke detector is positioned

light source

LANDINGOVERHE

Continuous-in the engineor overheat signal for owheel wellsspective redwarning panextinguishin

lights are teon the warni

A block diasystem is pre

FIRE TES

The various

with the TESing panel (Fi


ENGINE 1DETECTION

MONITORING UNIT

ENGINE 1FIRE

DETECTOR

• Continuity of bottle cartridge

• Operation of the fire detection units’warning function

• Integrity of the detector heads

• Operation of the red warning lights

• Operation of the aural warning system

Placing the switch in the LIGHTS position teststhe wheel well detector units, the associatedwarning lights and, on aircraft SNs 096 through999, the associated PCBs.

FIRE EXTIN

GENERAL

All of the fixed fire ethe same extinguis

romethane (CF3Br)

13B1 and in the Uni

Each of the extingu

colored, electrical



ENGINE 2FIRE

DETECTOR

ENGINE 3FIRE

DETECTOR

APUFIRE

DETECTOR

SMOKEDETECTOR

LEFT GEAR

WELLDETECTOR

ENGINE 2DETECTION

MONITORING UNIT

ENGINE 3DETECTION

MONITORING UNIT

APUDETECTION

MONITORING UNIT

PCB


Figure 8-

WARNING PANEL

TESTFIRELIGHTS

ENGINE 1DETECTOR

ENGINE 1EXTINGUISHERS

discharge heads. Each head is connected to one

of two extinguishing circuits. For the No. 1 and

No. 3 engines, a single bottle is discharged; for

the No. 2 engine, two bottles are discharged si-

multaneously because of the greater area of the

engine nacelle. Interconnection of the engine ex-

tinguishers enables successive use of two bot-

tles to extinguish a fire in one engine. The

bottles are shown in Figure 8-5.



R. ADAL. AOA

R. WHLOVHT

L. WHLOVHT

L. PITOT R.PITOTST BYPITOT

EXTINGUISHERS

AUDIOWARNING

ENGINE 2DETECTOR


ENGINE 3DETECTOR


APUDETECTOR

APUEXTINGUISHER

SMOKE

DETECTOR

8 - 8

F OR T RAI NI

N G P URP O

SMOKEDETECTOR

SELF-TEST

UNIT (2)

LANDING GEAR WELLLOCALIZED DETECTOR

(AIRCRAFT SNs 96 AND SUBSEQUENT)

PORTABLE HALON



R ev i s i on4

O S E S

ONL Y

LEGEND

FIRE DETECTORS

ENGINE BOTTLES

ENGINE EXTINGUISHING

PORTABLEHALON FIRE

EXTINGUISHER

WHEELWELL

PORTABLE HALONFIRE EXTINGUISHER

Figure 8-6. Fire-Exti

The bottles have pressure gages and a correc-tion table for checking nitrogen charge at vary-ing ambien t t empera tu res . A p ressure -relief valve dumps extinguishing agent in theevent of overpressurization. The exhausted

agent is vented overboard through the generaldrain mast in the rear compartment.

Figure 8-6 sdetector uni8-7 illustrattrols and sytifies discha


NORMAL

ENGINE 1 FIREFUELSHUTOFFSWITCH

FIRE

PANEL

210

210

FIRE 1

FAULT

DISCH

FIRST SHOT SECON

ENGINE1

1 2 3



ENGINE3

NORMAL

1 2 3

210

210

FIRE 2FAULT

DISCH

ENGINE 2 FIRE

210

210

FIRE 3

FAULT

DISCH

ENGINE 3 FIRE

FIREPANEL

FIREPANEL

NORMAL

FIRST SHOT


BAGGAGEFAULT

ENGINE2

1 2 3

FIRST SHOT (LEFT)

FIRST SHOT (RIGHT)


A1 B12 A12 B1

(3WB1)

A

(22WB-A) (22WB-B)

B

(22WB)

NO. 1 ENGINE

NO. 2 ENGINE

(L3WB2)

BAGGAGECOMPARTMENT APU

AB3

(R3WB2)



Figure 8-8. Identification of Discharge Head Conn

A1 FIRST SHOT, NO. 1 ENGINE

B12 SECOND SHOT, NO. 2 ENGINE

A12 FIRST SHOT, NO. 2 ENGINE

B1 SECOND SHOT, NO. 1 ENGINE

B3 SECOND SHOT, NO. 3 EN

A22 FIRST SHOT, NO. 2 ENGIN

ENGINE FIRE EXTINGUISHING

When an engine fire is indicated by illumina-tion of a FIRE light, the fuel valve to that en-

gine should be closed by raising the guard andactuating the FUEL SHUT OFF switch (Figure8-1). With the valve in transit, the amber TRANSlight illuminates, going out as the valve closes.Referring to Figure 8-7, note that moving theDISCH switch above the FIRE 1 light to the 1position discharges the No. 2 bottle into theNo. 1 engine nacelle. If the fire persists, ad-vancing the switch to the 2 position fires the No.

3 bottle into the same nacelle.

nacelle when the DIMoving the switch tthe No. 2 and No. 5

APU AND BAGCOMPARTMEEXTINGUISHI

The No. 1 bottle (Fthe APU and the bthe FIRE APU or thilluminates, actuatiabove the illuminainto the designated


of the cabin on the left side (See Figure 8-6).Each bottle contains 2.5 pounds (1.13 kg) of Halon 1211. The extinguishers can be used tofight Class B and C fires (liquids and electri-cal). A typical Halon 1211 extinguisher is

shown in Figure 8-9.



AUTOMATIC FIREEXTINGUISHERS(OPTION 26-20-10)

Two automatic fire extinguishers are installedin the front and rear lavatory trash cans. TheseHalon 1301 fire extinguishers are automati-cally discharged when the ambient tempera-ture reaches 76°C (170°F)

Figure 8-9. Halon Extinguisher

(Typical)


1. Placing the TEST switch in the FIRE po-sition tests the:

A. Continuity of the red warning lightsB. Continuity of cartridge circuitry

C. Integrity of the detector heads

D. All of the above

2. If the FIRE 1 light illuminates:

A. The No. 2 bottle automatically dis-charges into the No. 1 engine nacelle.

B. Crew action must be taken; bottle dis-charge does not automatically occur.

C. The No. 3 bottle automatically dis-charges into the No. 1 engine nacelle.

D Th N 1 d N 3 b ttl t

5. When an engin

A. Use the ap

release thethat nacelle

B. Wait until nacelle befo

C. P os it i on switch to Pbackup bott

D. A ct ua te

switch to cto tha t enDISCH swi

6. Actuating the d

QUESTIONS



D. The No. 1 and No. 3 bot tles auto-matically discharge into the No. 1 en-gine nacelle.

3. Illumination of an L. or R. WHL OVHTlight:

A. Automatically discharges a bottle intothe related wheel well

B. Triggers the aural “gear” warning fromthe audio control box

C. Warns of an overheat condition in thewheel well, but no provision is made forfire extinguishing in that area

D. A and B

4. If the discharge of one bottle into the No.1 or No. 3 engine nacelle does not extin-guish a fire:

A. A second bottle can be discharged intothe nacelle.

B. Two more bottles can be dischargedinto the nacelle.

C There is no provision for discharging

FIRE APU war

A. Dischargesof the No. 1

B. Di scha rgeof the No. 1

C. A r m s N o.switch musits contents

D. Dischargesof two bottl

7. The portable H

effective againA. Liquid and

B. Liquid, elefires

C. Liquid fires

D. Electrical f

CHAPTER 9

PNEUMATICS

CONTENTS

INTRODUCTION ......................................................................

GENERAL..................................................................................

ENGINE BLEED AIR...............................................................

General ................................................................................

LP Bleed Air .......................................................................




HP Bleed Air......................................................................

Bleed-Air Mixing...............................................................

APU BLEED AIR......................................................................

DISTRIBUTION ........................................................................

General ................................................................................

Anti-icing Systems Not Activated .....................................

Anti-icing Systems Activated .............................................

OPERATION ..............................................................................

Electric Motor-Operated Isolation Valve ...........................

APU Bleed Valve ................................................................

INDICATION .............................................................................

QUESTIONS..............................................................................

ILLUSTRATIONS

Figure Title

9-1 Bleed-Air System—Off..............................................

9-2 BLEED AIR Panel ......................................................

9-3 Bleed Air System—Power Level at Idle ....................

9-4 Bleed-Air System—Power Lever at Takeoff .............

9-5 Bleed-Air System—Anti-ice Systems On..................

9-6 Bleed-Air System—APU On ......................................

TABLES

T bl Titl




Table Title

9-1 Ground and Takeoff Conditions.................................

9-2 Flight Conditions........................................................

9-3 APU Bleed Air ............................................................

9-4 Bleed-Air Distribution................................................


L R

BLEED AI

5 1 5

2 0



INTRODUCTION

The pneumatic system on the Falcon 900 uses engine and Aconditioning, pressurization, and ice and rain protection sythe bleed-air sources, indication of malfunction, and warni

GENERAL

Air from the three engines or the APU is used bythe pneumatic system. The bleed air is di-id d i t t t i i bl d i d

ENGIN

GENERA

discharge. Figure 9-1 illustrates the bleed-airsystem.

Main HP circuit bleed valve operation iscontrolled by HP1, PRV 2, and PRV 3 switches

located on the BLEED AIR panel (Figure 9-2).

LP BLEED AIR

The main bleed-air source for LP bleed air isone inboard bleed port on the No. 1 and No.3 engines and the left side of the No. 2 engine.There is an auxiliary outboard bleed port on theNo. 1 engine and the right side of the No. 2 en-gine. The outboard bleed port of the No. 3 en-gine is not used.

HP BLEED AIR




The main bleed-air source for HP bleed air isthrough interconnected bleed ports on each en-gine. The auxiliary bleed-air source is one bleed

port on the upper centerline of each engine.

BLEED-AIR MIXING

At the outlet of each engine, airflows from themain LP and HP ports are mixed by venturi ac-tion so that the resulting pressure is greater thanLP bleed-air pressure. Figure 9-3 illustrates the

bleed-air system with the power levers at idle.LP and HP characteristics relative to enginestatus are listed in Tables 9-1 and 9-2.

Table 9-1. GROUND AND TAKEOFF CONDITI

Figure 9-2.

LP AIR

Pressure Temperature PresISA CONDITIONS


g t in

NO. 3ENGINE

MAIN LPBLEED PORT AUXILIARY HP

BLEED PORT

MAIN HPBLEED P

PRESSURE-REGULATING VALVE 3

LP 3 BLEED-AIRTEMPERATURE PROBE


CREW AIR-CONDITIONING SYSTEM



AIRINTAKE

ANTI-ICING



Revision 3

NO. 1ENGINE

PASSENGER AIR-

CONDITIONING SYSTEM

LP1 BLEED-AIRTEMPERATURE PROBE

BLE

MAIN HPBLEED PORT

MAIN LPBLEED PORT

AIRINTAKE

ANTI-ICING

AUXILIARY LPBLEED PORT

AUXBL




BLE

WI

CREW AIR-C

FUEL TANK PRE

PRES

PASSENGER AIR-C

MAIN LBLEED POR

AIRINTAKE

ANTI-ICING

;

ON IFISOLATION

VALVECLOSED

OVERHEAD PANEL

ISOL

HEAT

APU

NORM

PRV3PRV2

LPLP

ISOLATION

PASSENGER CREW

BLEED AIR

BAG

AUTOON

OFF

HP1

ISOL



9-4

MAIN LPBLEED PORT

AIRINTAKE

ANTI-ICING

; ; ; ; ; ; ; ; ;

; ; ;

; ; ; ;

;

;

;

;

;

; ;

;

; ;

; ;

;

;

;

;

; ; ; ; ;

FLASHING IF

CORRESPONDING

SWITCH

OFF

ON IF:• APU BLEED VALVE OPEN IN FLIGHT• APU BLEED VALVE OPEN ON THE GROUND AND POWER LEVER AT A POSITION ABOVE 54˚ FCU

WARNING PANEL

BLEEDOVHT BLEEDAPU

MAIN BLEED PO


Figure 9-3.

The three airflows are then mixed and supplya single air system, which can be divided intotwo separate subsystems by means of an elec-

LP bleed air (is available started. The


LPFLIGHT

CONDITIONS Pressure Temperature % per Pre

(absolute) engine (ab

Takeoff with 85.55 psi 455° F 188anti-icing: (5,900 mbar) (235° C) 2.4 (13,00

• TS = 0° C • N1 = 94%

Cruise without 25.81 psi 401° F 7anti-icing: (1,780 mbar) (205° C) 1.0 (4,83

• M = 0.75• Z = 41,000 ft• TS = –56.5° C• N1 = 98.8%

Table 9-2. FLIGHT CONDITION



p y ytric motor-operated isolation valve. The valveis normally selected open as long as no fault(leakage, smoke, etc.) is detected in either of

the subsystems.

One side of the isolation valve, connected tothe No. 1 and No. 3 engines, supplies com-pressed air to the cockpit air-conditioning andthe wing anti-icing systems. The other sidesupplies the passenger cabin air-conditioningand the No. 2 engine air intake and S-ductanti-icing systems.

The auxiliary HP bleed-air systems are com-pletely independent. Each supplies the air in-take anti-icing of the corresponding engine.

The auxiliary LP bleed-air systems of the No.1 and No. 2 engines are interconnected andsupply the fuel tank pressurization system.

Backflow from a running engine into an inop-erative engine, or an engine at a lower power

valves.

HP bleed air

by switches isurization je

APU BL

When the airfrom the APbleed-air mcabin air-cois also supplcontrolling ttank pressurof APU blee

DISTRI

GENERA

of HP and LP air to the pilot and passengerair-conditioning systems as well as to the wingand No. 2 engine S-duct anti-icing systems. AnLP air circuit feeds the fuel tank pressurizationsystem. Three separate HP circuits feed the anti-

icing systems of the No. 1, No. 2, and No. 3engine air intakes. A pressurization jet pumpand waste water circuit is supplied HP air from

the No. 1 and No. 2 e

In the main circuiadditional pressure to

Table 9-4 lists bleed

source and engine po


Table 9-4. BLEED-AIR DISTRIBUTION

LP

AIRPLANE SYSTEMSAUXILIARY

MAIN (NO. 1 AND(3 ENGINES) NO. 2 ENGINES) (3 E



Fuel tank pressurization 1-2

Pressurization jet pump

Waste water system

No. 1 engine air intake and heatexchanger inlet anti-icing

No. 2 engine air intake anti-icing


Low power setting 1-2-3Wing

anti-icingHigh power setting 1-2-3

No. 2 Low power setting 1-2-3engineS-ductanti-icing High power setting 1-2-3

Low power setting





B

W

CREW AIR

FUEL TANK P

PR

PASSENGER AIR

MAIN HP

MAIN BLEED PO

AIRINTAKE

ANTI-ICING

; ;

ON IF

ISOLATIONVALVE

CLOSED

WARNING PANEL

OVERHEAD PANEL

ISOL

HEAT

APU

NORM

PRV3PRV2

LPLP

ISOLATION

PASSENGER CREW

BLEED AIR

BAG

AUTOON

OFF

HP1

ISOL



9-8

MAIN HPBLEED PORT

MAIN LPBLEED PORT

AIRINTAKE

ANTI-ICING

; ; ; ; ; ; ; ; ;

; ; ;

; ; ; ;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

; ; ; ;

;

;

FLASHING IF

CORRESPONDING

SWITCH

OFF

ON IF:• APU BLEED VALVE OPEN IN FLIGHT

• APU BLEED VALVE OPEN ON THE GROUND AND POWER LEVER AT A POSITION ABOVE 54˚ FCU

BLEED

OVHT

BLEED

APU

Figure 9-4.



g t in

C

FUEL

PASSEN

MBLEE

AIRINTAKE

ANTI-ICING

ON IFISOLATION

VALVE

CLOSED

WARNING PANEL

OVERHEAD PANEL

ISOL

HEAT

APU

NORM

PRV3PRV2

LPLP

ISOLATION

PASSENGER CREW

BLEED AIR

BAG

AUTOONOFF

HP1

ISOL




NOTE: HP1, PRV2, AND PRV3 WILL AUTOMATICALLY CLOSE FOR 18 SECONDS WHEN ANTI-ICING OPERATION AND HIGH POWER SETTINGS ARE SELECTED.

BLEE

MBLEE

AIRINTAKE

ANTI-ICING

FLASHING IF

CORRESPONDING

SWITCH

OFF

ON IF:• APU BLEED VALVE OPEN IN FLIGHT• APU BLEED VALVE OPEN ON THE GROUND AND POWER LEVER AT A POSITION ABOVE 54° FCU

BLEED

OVHT

BLEED

APU

Figu



MABLEED

WING AN

CREW AIR-CONDI

FUEL TANK PRESSUR

PRESSURIZ

PASSENGER AIR-CONDI

MAIN HPBLEED PORT

MAIN LPBLEED PORT

AIRINTAKE

ANTI-ICING

; ;

ON IFISOLATION

VALVECLOSED

WARNING PANEL

OVERHEAD PANEL

ISOL

HEAT

APU

NORM

PRV3PRV2

LPLP

ISOLATION

PASSENGER CREW

BLEED AIR

BAG

AUTOON

OFF

HP1

ISOL




BLEED PORT

MAIN LPBLEED PORT

AIR

INTAKEANTI-ICING

; ; ; ; ; ; ; ; ;

; ; ;

; ; ; ;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

; ; ; ;

;

;

FLASHING IF

CORRESPONDING

SWITCH

OFF

ON IF:• APU BLEED VALVE OPEN IN FLIGHT• APU BLEED VALVE OPEN ON THE GROUND AND POWER LEVER AT A POSITION ABOVE 54̊ FCU

BLEEDOVHT

BLEEDAPU

Figu

54° PLA), and the degree of valve opening ismonitored by the APU electronic control unit sothat the APU T5 limit is not exceeded (the APUbleed valve is fully open as long as the T5 limitis not reached).

INDICATION

Indication of bleed-air system operation is ac-complished with BLEED OVHT and BLEEDAPU lights on the annunciator panel (AppendixB) and an ISOL light on the BLEED AIR panel(Figure 9-2).

The monitoring system consists of three tem-perature probes and an overheat detection box.

The three probes are located on the bleed-airducts downstream of the LP/HP air mixingventuri of each engine.

causes illumiThe BLEEDswitch (HP ing to the ovposition.

If the BLEEDcan easily overheating action to takoff one at a twhen the ligh

The ISOL ligvalve open,

sources. The the isolation v

The BLEEDas the APU billuminates:




venturi of each engine.

The overheat detection box receives the tem-

perature signal from the temperature probesand compares it with a preset value of 635° F(335° C) with anti-icing off or 545° F (285° C)with anti-icing on. When the detected temper-ature exceeds this threshold, the detection box

• When and at to a pgine p

• When and th


1. The pneumatic system uses bleed airfrom the:

A. No. 1 and No. 3 engines

B . No. 2 engine

C. APU

D. All the above

2. The engines supply the pneumatic sys-tem with:

A. HP bleed air

B. Mixed bleed a ir

C. LP bleed ai r

D. All the above

3. The switches for bleed-air control arelocated on the:

A O h d l

7. The No. 1 engionly when the

A. Airplane is

B. APU is inoC. Wing anti-

D. Air-conditing.

8. As soon as thea i r i n t a k e aswitched on a

than 378.5° FA. PRV 2 valv

B. PRV 2 valvregulating

C. PRV 2 valv

D PRV 2 valv

QUESTIONS



A. Overhead panel

B. Center console

C. Pi lot’s s ide panelD. Lower instrument panel

4. The main bleed-air source for LP bleedair is:

A. An inboard port on the No. 2 engine

B. An inboard port on the No. 1 andNo. 3 engines

C. T h e l ef t s i de p o r t on t h e N o. 2engine

D. Both B and C

5. Mixed bleed-air pressure is:

A. Less than LP bleed-air pressure

B. More than HP bleed-air pressure

C. More than LP bleed-air pressure

D. Only present on the ground

D. PRV 2 valv

9. The APU blee

A. The BLEEB. Three pow

switches

C. The APU e

D. All the abo

10 . The crew can dcircuit is over

1, PRV 2, and A. Off

B. AUTO

C. Off one at

D. NORM one

CHAPTER 10

ICE AND RAIN PROTEC

CONTENTS

INTRODUCTION......................................................................

GENERAL .................................................................................

NACELLE INTAKE ANTI-ICING............................................

General................................................................................

No. 1 Nacelle and Ram-Air Inlet Anti-icing.......................

No. 2 Nacelle and S-Duct Anti-icing..................................




No. 3 Nacelle Anti-icing.....................................................

WING LEADING-EDGE SLATS ANTI-ICING.......................

General................................................................................

Control ................................................................................


Operation ............................................................................

MAIN LANDING GEAR BRAKE HEATING SYSTEM (SB 3

General................................................................................

Control ................................................................................


Operation ............................................................................

AIR DATA SYSTEMS ANTI-ICING...................................................

General............................................................................................

Control............................................................................................

Monitoring......................................................................................

Indication and Warning..................................................................

Operation ........................................................................................

COCKPIT WINDOWS DEFOGGING AND DEMISTING ................

General............................................................................................

Windshield Defogging...................................................................

Control............................................................................................

Demisting.......................................................................................




Operation ........................................................................................

CABIN WINDOW DEMISTING.........................................................

WINDSHIELD WIPERS.......................................................................

General............................................................................................

Control............................................................................................

Operation ........................................................................................WATER WASTE DRAIN ANTI-ICING...............................................

General............................................................................................

Control............................................................................................

Operation ........................................................................................

LIMITATIONS......................................................................................


ILLUSTRATIONS

Figure Title

10-1 Nacelle Air Intakes......................................................

10-2 No. 1 Nacelle Air Inlet and Ram-Air Inlet..................


10-4 No. 1 Nacelle and Ram-Air Inlet Anti-icing Operatio

10-5 No. 2 Nacelle and S-Duct Anti-icing Operation .........

10-6 No. 3 Nacelle Anti-icing Operation ............................10-7 Wing Leading-Edge Slats Anti-icing Operation .........

10-7A Wing Leading-Edge Slats Anti-icing and Brake Heati

10-8 Windshield Anti-icing Operation ...............................



10-9 DV and Side Window Heat Operation .......................

10-10 Air Data Systems Components (Typical)....................

10-11 Air Data Systems Anti-icing Operation......................

10-12 Footwarmer/Defogging Control Lever .......................

10-13 Cockpit Windows Defogging/Demisting Operation...

10-14 Windshield Wiper Operation ......................................

10-15 Water Waste Drain ......................................................

10-16 Water Waste Drain Anti-icing Operation ....................

TABLES

Table Title

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;


ICE AND RAIN



INTRODUCTION

This chapter describes the ice and rain protection systemAll values used throughout this chapter, such as for pressused for their illustrative meanings only. Pertinent values muals and publications issued with the airplane.

GENERAL

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

NACELLE INTAKEANTI-ICING

GENERAL

The nacelle intakes (Figure 10-1) are anti-iced by hot bleed air from the associated en-gine’s HP compressor. The intake lips formD-ducts, and a piccolo tube is located in eachD-duct. The hot air is discharged through thepiccolo tube holes to impinge on the duct skinand prevent ice formation. The air is then dis-charged overboard through louvered outlets.

switch also controlsply to the PT2TT2 pro

The switch is lever


Figure 10-2. Na



NO. 1 NACELLE AND RAM-AIRINLET ANTI-ICING

General

The anti-icing HP bleed-air supply from theNo. 1 engine is used to anti-ice the nacelle inletlip and the ECU heat exchanger ram-air inlet(Figure 10-2) in the No. 1 engine pylon.

Control

position. When turnthe PT2TT2 and to th

operated, engine picing valve. If the eoutlet pressure openmaintaining a relatpressure, which isram-air inlet lip andthe nacelle inlet lip

Figure 10-1. Nacelle Air Intakes

Indication and Warning

A dual light above the ENG 1 switch providesindication and warning for the No. 1 enginenacelle anti-icing. The light has a green andan amber lens. The indication and warningsystem is operated by a low-pressure switchand a high-pressure switch controlled by aPCB. The low-pressure switch controls thegreen light, which comes on and remains onwhen the ENG 1 switch is on a nd the pressurein the anti-icing line is at least 4 psi. Theamber light provides dual indication; if it ison steadily when the ENG 1 anti-icing switchis on and the green light is off, it indicates thatline pressure is less than 4 psi or that the anti-icing valve failed to open. If the amber lightis flashing when the ENG 1 switch is on, it in-dicates that line pressure is greate r than 90 psi,possibly caused by failure of the regulating fea-

ture of the anti-icing valve. The amber lightcomes on momentarily when the ENG 1 switchis turned on; it should go off when the greenlight comes on. The amber light also comes onflashing when the ENG 1 switch is turned off and the anti-icing valve fails to close. Thereis no indication or warning of operation ormalfunction of the PT2TT2 probe anti-icingsystem.

OperationFigure 10-4 in conjunction with Table 10-1depicts the operation of the No. 1 nacelle andram-air inlet anti-icing. The pressure-regulatingvalve is modulating the supply of HP bleed airto maintain a constant pressure in the anti-icingline and limits that pressure to 65.2 psi. The HP

air supply is applied directly to the ram-air inletlip and through a flow limiter to the piccolo tubein the nacelle intake lip. The anti-icing air is ex-hausted to the atmosphere through louvers on thenacelle and through drilled holes on the pylon.

NO. 2 NACELLE AND S-DUCT

ANTI-ICINGGeneral

Two separate areas of the No. 2 nacelle requireanti-icing: the nacel le air intake lip and the in-take S-duct.

The air source for the intake lip is suppliedfrom the No. 2 engine HP compressor. The airsource for S-duct anti-icing is LP air supple-mented by HP air, based on LP duct temper-ature, as described in Chapter 9, “Pneumatics,”and Chapter 11, “Air Conditioning.”

Normally, the air supply for the S-duct anti-icing is an HP/LP mix from the bleed-air ma n-ifold interconnecting all three engines; however,if the ISOLATION switch on the BLEED AIR

panel (Figure 10-3) is off (stripe vertical), theNo. 2 engine continues to supply air direct ly tothe bleed-air manifold, providing for S-ductanti-icing and cabin air-conditioning operation.

The anti-icing air distri bution for the No. 2 na-celle intake lip is identical to that for the No.1 nacelle intake lip. The hot air distribution forthe S-duct consists of four perforated lineswhich distribute the air to transverse ductsforming the double skin of the S- duct.

ControlThe No. 2 nacelle air intake and the S-duct anti-icing systems are control led by a PCB in con-

juncti on with a tw o-posi tion leve r-loc k switc hlabeled “ENG 2” located on the ANTI-ICE sec-tion of the overhead panel, in conjunction withthe PRV 2 switch on the BLEED AIR sectionof the overhead panel. In addition, the E NG 2



Figure

FLOW

LIMITER

LP PR

SW

T2 T2P T

PROBE

ECU HEAT

EXCHANGER

RAM-AIR INLET

ANTI-ICING

ENG1/ENG 3

SWITCH S

Off Not

On EOn E

On EOn E

Off NotOff Not

Off

Table 10-1.






FA

F

i g u r e 1 0 - 5 . N o . 2 N a c e l l e a n d S - D u c t A n t i - i c i n g O p e r a t i o n

I S O L

H E A T

A P U

N O R M

P R V 3

P R V 2

E N G 2

L P

L P

I S O L A T I O N

P A S S E N G E R

C R E W

B L E E D A I R

B A G

A U T O

O N O F F

H P 1

L E G E N D

H P B L E E D A I R

L P B L E E D A I R

H P / L P M I X

E L E C T R I C P O W E R

A M B I E N T

A U X I L I A R Y H P B L E E D A I R

C E N T E R

P E D E S T A L

E N G I N E

N A C E L L E

P R V

P I C C O L O T U B

E

M A N I F O L D

N O . 3 E N G I N E

C H E C K

V A L V E I S

O L A T I O N

V A L V E

F U L L O P E N I N G

S O L E N O I D

T O P A S S E N G E R

A I R C O N D I T I O N I N G S Y S T E M

T O

W I N G

S - D U C T

A N T I - I C E

V A L V E

H P

P R E S S U R E

S W I T C H

N O . 1 E N G I N E

L

P P R E S S U R E

S W I T C H E S

F L O

W

L I M I T

E R

A P U

P R I N T E D

C I R C U I T

B O A R D

N O . 2

E N G I N E

I N T A K E L I P

S - D U C T

;

;

;

;

;

;

N O . 2

E N G I N E

P R V

2

L P 2

S E N S O R

I S O L

T 2

T 2

P

T

P R O B E

H P 1

E N G 2

S W I T C H

P T 2 T T 2

S E N S O R

S - D U C T

A N T I - I C I N G

V A L V E

S - D U C T

P R E S S U R E

T ° L

P 2 / P R V 2

B L E E D A I R

A I R I N T A K E

L O W

P R E S S U R E

A I R I N T A K E

A N T I - I C I N G

P

R V

E N G 2

L I G H T

E N G 2

S W I T C

H

N o t e n e r g i z e d

E n e r g i z e d

E n e r g i z e d

O f f

O n

O n

C l o s e d

O p e n i n

g

O p e n

< 4 p s i

< 4 p s i

> 4 p s i

< 4 p s i

> 4 p s i

< 4 p s i

N o r m a l

I n i t i a l S t a t u s

A c t i v a t i o n

< 1 8 0 ° C

/ O p e n i n g

R e g u l a t i n g

O u t

A m b e r s t e a d y

G r e e n

< 9 0 p s i

< 9 0 p s i

< 1 8 0 °

C / O p e n

C l o s e d

< 9 0 p s i

O p e n i n g

T a b l e 1

0 - 2 . N O . 2 N A C E L L E A N D S - D U C T A N T I - I C I N G L O G I C S Y S T E M




switch also controls the electrical power sup-ply to the PT2TT2 probe in the No. 2 nacelle airintake.

Turning the ENG 2 on completes a circuit forthe PT2TT2 probe heat and applies electrical

power to the engine nacelle air intake pressure-regulating valve, which operates the same asthe No . 1 nace l l e a i r in t ake p ressure -regulating valve. The engine nacelle air intakepressure-regulating valve opens and suppliesHP air from the No. 2 engine HP compressorthrough a flow limiter to the piccolo tube inthe nacelle intake lip. This valve modulates tocontrol pressure exactly as described for the

No. 1 nacelle.

Simultaneously, a circuit is completed to theS-duct anti-icing valve, which opens to sup-ply bleed air through a flow limiter to the S-duct distribution system, as determined by LP2 duct temperature. If LP 2 duct temperature

provided as to determinethe nacelle o

Operatio

Figure 10-5 ipicts the opeduct anti-icinacelle lip isof the No. 2 evalve which constant preaddition limiThis air is cirthen dischargvered outlets

At the same tand suppliesThe mix ratiwill modula




p pis less than 192.5°C, a circuit is comple ted tothe full opening solenoid of PRV 2, and PRV2 opens fully and supplies HP 2 air to mixwith LP 2 air. If LP 2 duct temperature ishigher than 192.5°C, PRV 2 operates as a pres-sure-regulating valve, allowing LP air to main-tain a specific duct pressure.

Indication and WarningThe indication and warning systems associatedwith the No. 2 nacelle and S-duct anti-icingsystem consist of a PCB, a high-pressureswitch, two low-pressure switches, S-duct andair intake PRV valve position, PRV position,and a dual-lens light. The light is locatedabove the ENG 2 switch (Figure 10-3). Thegreen and amber lenses provide indication of operation corresponding to the ENG 2 switch

on the overhead panel. Light illumination iscontrolled according to data from the posi-

will modulaLP 2 duct. Fational char9 , “PneumConditionin

NO. 3 NA

GeneralThe No. 3 na

bleed air frosor. The airthrough a flothe air is disclouvered ou

Control

The system i

lock switch overhead pa

Indication and WarningThe indication and warning systems includea PCB, high-pressure switch, low-pressureswitch, and dual-lens light located above thecontrol switch. The green light, when on, in-

dicates operation, and the amber light indicatesmalfunctions. Operation and malfunction in-dication is identical to the description giveearlier for the No. 1 nacelle.

OperationFigure 10-6 in conjupicts the operation icing system. HP bengine HP compres

pressure-regulat ithrough a flow limthe nacelle lip. Folloduct, the air is discthrough louvered


ENG 1 ENG 2 EN

ANTI-ICE



PYLON

NO. 3 ENGINE

T2 T2P T

PROBE

PRINTEDBOA

The pressure-regulating anti-icing valve mod-ulates to maintain a relatively constant pres-sure in the anti-icing duct and limits thepressure to 65.2 psi.

WING LEADING-EDGESLATS ANTI-ICING

GENERAL

The sectional leading-edge slats and the fixedinboard leading-edge sections of each wingpanel are anti-iced by hot engine bleed-air.

Though system efficiency is predicated onbleed-air supplies from the No. 1 and No. 3 en-gines, the No. 2 engine bleed-air supply is alsonormally available to the wing slats and the in-board wing root fixed leading-edge sections of the wing unless the isolation valve is closed.

The slats and the fixed wing root leading edge

if the temperotherwise, itvalve, proviNo. 3 enginoperation of“Air Conditi

INDICAT

Indication anlens light loc3). The greenoperation. Tflashing indi

OPERAT

Figure 10-7 tion, indicatiing-edge slatswitch selecopens HP 1




The slats and the fixed wing root leading-edgesections are double-skinned and form a duct.

A manifold is installed in each leading-edgeslat. Hot air is supplied directly to the fixedinboard wing root sections and through tele-scoping tubes to the movable slat section man-ifolds on each wing.

CONTROL

The wing leading-edge slat anti-icing is con-trolled by a PCB in conjunction with a two-position lever-lock switch labeled “WING”on the ANTI-ICE section of the overheadpanel, and also in conjunction with the HP 1and PRV 3 switches on the BLEED AIR sec-tion of the overhead panel. When the switchis on (up), power is supplied to open the winganti-ice valve. The HP 1 valve opens if the LP1 bleed-air temperature is less than180°C;otherwise it remains closed. PRV 3 opens fully

opens. HP 1through its e

when the tembleed air is ltemperature 3 to resume HP/LP bleedducts to the leicing air is susections of thand right m

where it is dspective movtubes which cflow throughtion inset in Fthe operatioindication, awing slats an

1 0 - 8

F OR T RAI NI N

G P URP O S E S

ON

WING

N

MAIN LPBLEED AIR

(LP3)

ANTI-ICEPANEL

RIGHT WINGPRESSURESWITCH

192.5°C

192.5°C

PEDESTAL

POWER LEVER SWITCHES

FIXEDWING ROOT

LEADING EDGE

PRINTEDCIRCUITBOARD

MANIFOLD



R ev i s i on 3

LY

N

MAIN HPBLEED AIR

MAIN LP BLEEDAIR (LP1)

LEFT WING PRESSURESWITCH

TELESCOPICTUBES

Figure 10-7. Wing Leading-Edg


F OR T RAI NI N G P URP O S

LEFT RIGHT

WING ANTI-ICING WING WING

SWITCH VALVE PRESSURE PRESSURE T

Initial St

OFF Closed <4 psi <4 psi Activatio

On Opening <4 psi <4 psi <18On Open >4 psi >4 psi <On Open >4 psi >4 psi >2On Open >4 psi >4 psi <On Open >4 psi >4 psi <2On Open >4 psi >4 psi <

On Open >4psi >4 psi >2

On Open >4 psi >4 psi Eng


On Open >4 psi >4 psi Betweenand 200



On Partial Opening >4 psi >4 psi

Abnormal C



1 0 - 9

E S

ONL Y

NOTES1. The green light stays on for 23 seconds in the event of the LP temperature threshold being 2. If one of the engines (1or 3) is shut down (power lever angle less than 15°), the position of tprovided the air pressure is sufficient (increased power setting on the other engines).3. Air bleeding is nomal in the following conditions:

• LP 2 is less than 180°C, and HP 1 valve is open or PRV 3 is fully open.• If LP 2 is between 180 and 200°C, HP 1 valve can be open or closed or PRV 3 fully open• LP 2 is greater than 200°C, and HP valve is closed or PRV 3 is regulating.

On Partial Opening >4 psi >4 psiOn Open <4 psi >4 psiOn Open >4 psi <4 psi

On Open >4 psi >4 psi <180On Open >4 psi >4 psi

On Open >4 psi >4 psi >200

On Open >4 psi >4 psi

Off Closing >4 psi >4 psiOff Closed <4 psi <4 psi

Off Not closed >4 psi >4 psi

Off <4 psiOff >4 psi

Wing Switc

Abnormal C




ACTIVATION OF WING ANTI-ICING SY



; ; ; ; ; ;

; ; ; ; ; ;

; ; ; ; ; ;

; ; ; ; ; ;

; ; ; ; ; ;

; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

; ; ; ; ; ; ;

ACTIVATION OF WING ANTI-ICING AND BRAKE

Figure 10-7A. Wing Leading-Edge Slats Anti-icing and


Table 10-4. WING LEADING-EDGE SLATS ANTI-ICING HEATING LOGIC SYSTEM

SWITCHPOSITION

OPERATION OF THEWING ANTI-ICING

SYSTEM

OPERATION OF THEBRAKE HEATING

SYSTEM

Initial State (See Wing Anti-icing)

Activation of Brake Heating System

Abnormal Conditions

Reset to Initial Condition

WING Correct

Correct

Incorrect

Correct

Correct

Correct

Incorrect

Incorrect

Incorrect

WING-BRK

WING-BRK

WING-BRK

WING-BRK



WINDSHIELD ANTI-ICING AND SIDEWINDOW HEAT

GENERAL

The cockpit windows are categorized as pilot’sand copilot’s windshields, center windshield,pilot’s sliding DV window, copilot’s side window,and pilot’s and copilot’s aft side window.

Anti-icing and birdproofing is provided forthe pilot’s windshields. The system is designed

WINDSHIELD

GeneralThe windshield anttwo separate but idepilot’s system provpower for the pilot’the center glass. Tvides anti-icing eleglass and the right h

Each system consistemperature-sensinbutyl member loca

WING Correct

The electrical control power for each system isalso separate: the pilot’s system is powered frombus A1, and the copilot’s system from bus B2.

ControlA three-position control switch is providedfor each pilot on the WINDSHIELD sectionof the overhead switch panel (Figure 10-3).The switches are identified “PILOT” and“COPILOT,” and the positions are labeled“OFF,” “NORM,” and “MAX.” When theswitches are in the NORM position, the heat-ing element in each pilot’s windshield is in se-ries with the associated heating element inthe center glass. The associated controller di-rects main bus power through the normal heat-ing contactor to the appropriate glasses inresponse to the input signals from the tem-perature sensor in each pilot’s windshield.The temperature is maintained between 77

and 86°F.

The MAX position is used only when theNORM selection does not maintain the glassesfree of ice. Selecting MAX closes a high heatrelay connecting the pilot’s windshield glasselements in parallel with the associated cen-ter windshield element through a resistor. Theheating current now takes the line of least r e-sistance through the pilot’s windshield heat-ing elements, resulting in an increase in powerto maintain the pilot’s glass free of ice. Somecurrent still continues to flow through the as-sociated center windshield element but notsufficient to maintain the glass free from ice.The glass temperature range on the pilot’swindshield remains the same; however the cy-cling rate is almost doubled.

Indication and WarningAn amber light, labeled “XFR,” above thewindshield anti-icing control switches comeson if a temperature sensor short-circuits oropens, or if sensor resista nce is too high or low.

At the same time the affected controller au-tomatically transfers the system to the oper-ating controller so that both systems arecontrolled by a single controller, and normaloperation continues.

OperationFigure 10-8 depicts normal operation of the

windshield anti-icing system. Both switchesare at the NORM position, and the normalheat control relays are energized. The main andcenter glass heating elements are in serieswith each other. Moving a switch to MAXadds a nonheating resistor to the main glassheating element; consequently, the main heat-ing element has the lowest r esistance, and cy-cling time increases, maintaining the mainglass free of ice. The inputs from the temper-ature probes to the individual controllers cyclethe power on and off to maintain the glasstemperature range of 77 to 86°F.

SIDE WINDOW HEAT

GeneralThe pilot’s sliding DV window, the copilot’s

side window, and the left and right side win-dows are all heated to improve visibility andcockpit environmental conditions. The sys-tem includes two identical controllers: onefor the pilot’s DV window and copilot’s sidewindow, and one for the aft side windows.The copilot’s side window and DV controllerreceives inputs representing temperature froma sensor embedded in the copilot’s side win-dow; a similar sensor in the left aft side win-dow supplies the inputs to the aft windowcontroller.

ControlThe system is controlled by a single switch la-beled “SIDE” located to the right of the pilot’sand copilot’s windshield heat switches. Whenthis switch is on, the controllers cycle powerto the respective glasses to maintain the glasstemperature between approximately 77 and86°F.

OperationFigure 10-9 depicts the operation of the sidewindow heat. The SIDE switch is on, supply-ing power through the control relays to the re-spective glasses. The sensor inputs to thecontrollers determine the cycling of power tothe glasses to maintain the temperature with-ing the design range.



SLI

PILO

AFT WINDO

PILOT’S

CONTROLLER






AIR DATA SYST

ANTI-ICINGGENERAL

The air data anti-icing sypilot’s pitot probes, the ststall warning vanes, the standthe OAT sensor. Figure 10-10system components which

CONTROL

The air data system anti-icithree-position switches onof the overhead panel (Fswitches are labeled “PILO“COPILOT.” When the PILon (up), DC power is applie

ement in the left pitot probthe left and right static portin the left stall warning vathe element in the OAT senST-BY switch applies powthe standby pitot probe. TuLOT switch applies power tright pitot probe, the eleme

FA

X F R

P I L O T

C O P I L O T

M A X

N O R M

O F F

S I D E

W I N D S H I E L D

H E A T E R

H E A T E R

D V W I N D O W

R E L A Y

A F T S I D E W I N D O W

R E L A Y

L E F T A F T

W I N D O W

R I G H T A F T

W I N D O W

S E N S O R

P I L O T S S L I D I N G

D V W I N D O W

A F T W I N D O W C

O N T R O L L E R

( S A M E A S S I D E )

C O P I L O T S

S I D E W I N D O W

S L I D I N G

W I N D O W

C O N T A C T

C O N T R O L S T A G E

S E L F H O L D

R E G U L A T I N G

O P E N C I R C U I T

S H O R T

C I R C U I T

S I D E W I N D O W

H E A T C O N T

R O L L E R

A 2

B 1

+ + +

V R E F

1 V R E F

3 V R E F

2

C E N T E R

W I N D S H

I E L D

C O P I L O T S

W I N D S H I E L D

P I L O T S

W I N D S H I E L D

C O P I L O T S

S I D E W I N D O W

P I L O T S

S L I D I N G D V

P I L O T S

A F T W I N D O W

C O P I L O T S

A F T W I N D O W

L E G E N D

D C P O W E R

S E N S I N G

G R O U N D

Figure

LEFT PITOT HEAD STALL

STANDBY PITOT HEAD

F i g u r e 1 0 - 9 . D V a n d S i d e W i n d o w H

e a t O p e r a t i o n





PILOT

POWER

MONITORS

POWERMONITOR

ST-BY

PITOT

STALL WARNING VANES

PITOT PROBE

STATIC PORTS

COPILOT

STANDBY PITOT PROBE(LEFT SIDE ONLY)

OATSENSOR

LEFT



CMPTR 1

OIL 1

GEN 1

BAT 1

FUEL 1

LOFUEL 1

CMPTR 2

OIL 2

GEN 2

BUS TIED

FUEL 2

REV

UNLOCK

LOFUEL 2

CMPTR 3

OIL 3

GEN 3

BAT 2

FUEL 3

FUELING

LOFUEL 3

L. PITOT

AOAPROBE

HOTBAT

XTK 2OPEN*

AP

AILZERO

L. AO

ST BPITO

L. WHOVH

AUTSLAT

XTKCLOS

MISTR

AILFEE

LEGEND

DC POWER




;

;

;

;

; ;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

DISTRIBUTORVALVE

DISTRIBUTORVALVE

DUCT

DUCT

DEFOGNOZZLE

OUTERPANE

GLARESHIELDVENTILATION

PORTS

DEFOGNOZZLE

CONDITIONAIR

EFIS GRILLEBLOWERS



; ;

;

;

;

;

;

;

;

;

;

; ; ; ;

; ; ; ;

;

;

;

PLEXIGLASSPANE

INNERPANE

DEMISTING

DEMISTINGHOLE

HEATING ELEMENT

LEFT AFT WINDOW SEEN FROM IN

CONTROL

The wipers are controlled by three-positionswitches located on the WIPER sections of theoverhead switch panel (Figure 10-3). Theswitch positions are labeled “OFF,” “SLOW,”

and “FAST.” Each wiper system includes aDC motor and converter unit, fast and slow re-lays, and a travel-limit switch. When a WIPERswitch is moved to the FAST position, the fastand slow relays are both energized, and themotor rotates at maximum rpm. The converterunit converts the rotary motion of the motorto a reciprocating motion for the wiper arm.Selecting the switch to SLOW energizes only

the slow relay, and the motor operates at lowrpm. Moving the switch to OFF completes acircuit to the motor through an end-of-travelswitch which is opened by a cam when themotor reaches the stowed or parked positionof the wiper arm.

OPERATION

CONTROL

The water waste drDC power from bubreaker and operateerator power or ext

throughout the airp

OPERATION

Figure 10-16 depicanti-icing. When powcuit breaker is in, thewhen at least one geis supplying the bu

distributed in parall

LIMITATION

All limitations contaipertaining to ice andunder law regardless




Figure 10-14 depicts the operation of the wind-shield wipers. The pilot’s WIPER switch is atthe FAST position, and the fast and slow re-lays are energized, completing a circuit to themotor windings. The copilot’s switch has beenmoved to OFF, completing a circuit throughthe limit switch to the motor windings. Themotor operates until the limit switch is openedby the end of travel cam, at which time thewiper arm is at the fully stowed position.

WATER WASTE DRAINANTI-ICING

GENERAL

The water waste drains direct water waste into

g

R ev i s i on1

F OR T R

AI NI N G P URP O S E S

B3

B1A3A1X1

X2

3A

2B

1B

1B

1A

2B

15A

28V

BUS A2

PILOT

WIPER

FASTSLOWOFF

FASTSLOWOFF

COPILOT

WIPER

PILOT'S WIPER

OVERHEAD PANEL

RIGHT SLOSPEEDRELAY

SLOW

SLOW

FAST

OFF

OFF



1 0 - 1 9

ONL Y

B3B1A3

A1X1

X2

3A1A

15A

28V

BUS B2 COPILOT'S WIPER

LEFT SLOSPEED RELFAST

Figure 10-14. Windshi


DRAINAG

FRAME 25 SEEN FROM THE REAR

HEATINGELEMENTS

WASTE WATER



DRAINAGE

WASTE WATERDRAINAGE

DETAIL OF

HEATING ELEMENT (BEFORE SB

TAPED DRAIN HOSE


DRAINHEAT

ANTI-ICE

RELAY

ONE GENERATOOREXTERNAL POWAIRCRAFT BUS

5AA2

A1

X2

X128V

BUSA2

Figure 10-16. Water Waste Drain Anti-ici




1. The ECU ram-air inlet lip is:

A. Deiced by electrical power

B. Supplied anti-icing air from the

wing slat systemC. Anti-iced by HP bleed air from the

No. 1 engine

D. In common with the No. 2 engineS-duct

2. The engine nacelle anti-icing valvesare:

A. Electropneumatic pressure-regulating

B. Motor operated

C. Electrically operated by a PCB

D. Temperature modulated by ductsensors

3. If an engine nacelle anti-icing valvefails to close when the switch is turned

5. Electrical powis controlled b

A. Associated

B. Pilot’s pitoC. Standby pi

D. Copilot’s p

6. The MAX posiheat switch is u

A. XFR light c

B. NORM pos

ice formatiC. OAT is –60

D. Airplane isflight.

7. Demisting air drilled in the i

A. Pilots’ win

QUESTIONS



fails to close when the switch is turnedoff, the:

A. Green and amber lights remain on

B. Green light flashes

C. Amber light is on

D. Green light goes out, and the amberlight flashes

4. The anti-icing air supply for the No. 2nacelle S-duct is available, with theisolation valve open, from the:

A. No. 2 nacelle anti-icing duct

B. Bleed-air manifold

C. Nacelle anti-icing duct of eitherengine

D. No. 1 engine LP duct

B. DV windowC. Aft side wi

D. Center win

8. Power supply i

A. Standby pi

B. Static port

C. Stall warni

D. Oat sensor



LIMITATIONS ......................................................................................

QUESTIONS.........................................................................................




ILLUSTRATIONS

Figure Title

11-1 Overhead Switch Panel...............................................

11-2 Bleed-Air Operation—Anti-icing Off ........................

11-3 Bleed-Air Operation—Anti-icing On.........................

11-4 Bleed-Air Operation—APU.......................................

11-5 Heat Exchanger Ram-Air Inlet...................................

11-6A Turbofan Operation—Ground or Slow Flight ............

11-6B Turbofan Operation—Normal Flight..........................

11-7 Heat Exchanger Low-Pressure Air Inlet.....................

11-8 Heat Exchanger Exhaust Outlet ..................................

11-9 Turbocooler .................................................................




11-10 Normal Operation—Ground or Slow Flight (Cooling

11-11 Normal Operation—Flight (Heating)..........................

11-12 Water Separator (Typical) ...........................................

11-13 Distribution System.....................................................

11-13A Distribution System—Aircraft SN 163 and Subseque

11-14 Duct Interconnect Flap “COND” Lever......................

11-15 Footwarmer and Defogging Valve Lever (Typical).....

11-16 Temperature Controller ...............................................

11-17 Emergency Pressurization Controller..........................

11-18 Temperature Control Operation


AIR C



INTRODUCTION

This chapter describes the air-conditioning system of thThe air-conditioning system includes all of the following ing of the occupied areas of the cockpit and passenger cnose cone area, underfloor conditioning for the floor pancomponents at frame 25, toilet and baggage compartment cdefogging, and forced circulation in the area of the EFIS

GENERAL

Ozone catalyzers are installed in the passen-ger and crew ducts.

Independent electronic temperature controlsystems are provided for the crew and pas-senger system; each may be operated in fully

automatic or manual modes. On-ground con-ditioning is available without the use of ex-ternal control units.

AIR SOURCES

GENERAL

For descriptive purposes the air sources con-sist of LP, HP, and APU air. The LP source of air is supplied from the final stage of each en-gine’s LP compressor. The HP air source issupplied (when specifically required) fromeach engine’s HP compressor. Air is suppliedfrom the compressor of the APU to facilitateair-conditioning when the airplane is on the

d d h i i LP

CONTROL

The HP bleed-air vafer somewhat for th2 and No. 3 enginesthe No. 1 engine HP

position switch, labBLEED AIR sectiopanel (Figure 11-1).open or closed. Whe(up), the valve reWING switch labeICE section of the o(up) position and thless than 378°F (19

Once open, the HP fully open until eithWING anti-ice swiduct temperature exeither case the HP position.

When open, the HP

l f HP i h




ground and the engines are not running. LPbleed air is nonselective; since there are no LPbleed-air valves, it is available once an engineis started. The LP bleed air from the No. 1 andthe No. 3 engines is supplied directly to acommon distribution manifold. The No. 2 en-gine LP bleed air is supplied to the same man-ifold through a normally open isola tion valve.This manifolded supply is available to the

crew air-conditioning system. The LP bleed airfor passenger air conditioning is supplied bya branch duct directly from the LP compres-sor of the No. 2 engine, or from the No. 1 andthe No. 3 engines to the same branch ductthrough the normal open isolation valve. APUair can be supplied to both air-conditioningsystems when the airplane is on the ground.

The HP air source is a supplemental supply andthough selective its operation once selected is

volume of HP air thrthe LP air duct; this imum mixed bleed-

The HP bleed-air sthe No. 3 engine is regulating valve (PRvalves are electropncontrolled by associa

2” and “PRV 3” on tthe overhead switch

The PRV 2 valve opswitch on the BLE(up) and the “ENG 2panel is on (up) andis less than 378.5°Fduct temperature

(192 5°C) the PRVl i l




The PRV 3 valve is primarily controlled by thePRV 3 switch on the BLEED AIR section of the overhead switch panel. The valve will befully open whenever the WING anti-ice switchis on and duct temperature upstream of its jetpump nozzle is less than 374°F (190°C). It

modulates as a pressure regulator when ducttemperature exceeds 378°F (192.5°C). TheHP bleed air passing through the valve is in- jected into its LP duct, as described earlier.

The auxiliary air source is supplied from theAPU compressor through a bleed-air valve anda check valve. At engine speeds above 97%rpm, when APU bleed air is selected, the APU

surge valve closes. This supply is directed intothe LP/HP duct of the No. 2 engine and is there-fore available to the bleed-air manifold throughthe normally open isolation valve as a cockpitair supply; in addition, it is directly availableto the cabin air-conditioning valve bypassingthe manifold.

The APU bleed-air valve is directly controlled

by the two position switch labeled “APU ”

function of engine pPRV 2 and PRV 3 vmined by power lewhile the position otion of wing anti-icture. This valve is e

controlling temperaassociated LP duct

Figure 11-2 depictwith the anti-icing

Figure 11-3 depictwith the anti-icing

Figure 11-4 depictation.

For more informatifunctions, control, see Chapter 9, “Pnemanual.

INDICATION




by the two-position switch labeled APU,located on the BLEED AIR section of theoverhead switch panel, in conjunction with aproximity switch operated by each engine’spower lever and output signals representingrpm and EGT transmitted from the APU elec-tronic control unit (ECU). When the APUbleed-air switch is on, APU rpm is 97% orgreater, and all engine power levers are set to

less than 54° PLA, the APU bleed-air valveis open; however, it modulates toward closedin response to ECU signals representing APUEGT (T5). The APU bleed-air valve goes tothe closed position when any power lever isadvanced to a position representing 54° PLA,whether or not the engine is running. Failureof the valve to close in this condition will il-luminate the BLEED APU light after a 12-

second delay

A separate temperaprovided to detect omixed-air ducts dowpumps (venturis). Tature information twhich compares the635°F (335°C). The on operation of the w

systems. When thilight labeled “BLEEpanel (Appendix Bsystem may be idenPRV 3, HP 1, and PRThe light begins flswitch is turned offAPU,” also on the mif the APU bleed-a

any power lever is

R ev i s i on 3

F OR T R

AI NI N G P URP O S E S

O

N

MAIN HPBLEED PORT


CREW AIR-CONDITIONINGVALVE



PASSENGER CABINAIR-CONDITIONING VALVE


PRV3

AUXILIARY HPBLEED PORT

MAIN LPBLEED PORT

NO. 3ENGINE

LP3

BLEED-AIR

TEMPERATURE PROBE

OVERHEAT TE

ERATURE PRO

LP1 BLEED-AIR

TEMPERATUREPROBE



1 1 - 5

NL Y

NO. 1ENGINE

ISOLATION

MAIN HPBLEED PORT

MAIN LPBLEEDPORT


HP1BLEED VA


BLEED PORT

TEMPERATUREPROBE

Figure 11-2. Bleed-Air Op

1 1

- 6

F OR T RAI NI N G

P URP O S E S

ONL Y

MAIN HPBLEED PORT






ISOLATIOMAIN LP

BLEEDPORT

AIR INTAKE

ANTI-ICING

PRV3


MAIN LPBLEED PORT

NO. 3ENGINE

LP3

BLEED-AIR

TEMPERATURE PROBE

OVERHEAT TE

ERATURE PRO


PROBE



R

ev i s i on 3

Figure 11-3. Bleed-Air Op

NO. 1ENGINE

MAIN HPBLEED PORT

PORTAIR INTAKEANTI-ICING

HP1BLEED VA


BLEED PORT

R ev i s i on 3


O

NL

MAIN HPBLEED PORT







PRV3


MAIN LPBLEED PORT

NO. 3ENGINE

LP3

BLEED-AIR

TEMPERATURE PROBE

OVERHEAT TE

ERATURE PR




1 1 - 7

LY

Figure 11-4. Bleed-A

NO. 1ENGINE

ISOLATIO

MAIN HPBLEED PORT

MAIN LPBLEEDPORT


HP1BLEED V


BLEED PORT

PROBE


AIR CONDITIONING

GENERAL

Air conditioning consists of heating and cool-ing systems. Cooling functions are performedautomatically by an environmental controlunit (ECU). The bleed-air supply is directedto the ECU by two air-conditioning valves, onefor the crew and one for the passenger system.

The air-conditioning valves are controlled byassociated switches labeled “PASSENGER”and “CREW,” located on the BLEED AIR sec-tion of the overhead panel. The switches havethree positions: AUTO, ON, and OFF. TheON position drives the associated valve to theopen position. The OFF position drives theassociated valve to the closed position.Thefunctions of the AUTO position of the air-conditioning switches are determined by theairplane’s configuration. When the airplane ison the ground, placing the air-conditioningswitches switches to the AUTO position drivesthe associated valve open if all three power

heat exchangers is ris greater than 300 kinlet (Figure 11-5) leading edge. Whenborne and TAS is leair for the heat exch

bofan assembly, andgine pylon inlet. Its later in this chapteachieved by a turbooperating in conjunchanical water sepa



p plevers are at a position less than 54° PLA. Thevalves remain open until any power lever is ad-vanced to 54° PLA or more. This opening orclosing rate is at the normal operating rate of two to four seconds. When the switches are atthe AUTO position, both air-conditioningvalves close when takeoff power is applied. Atlift-off the valves begin opening at the normalrate during the first 0.7 second, ensured bytime-delay relays; then electrical sequencersassume control and this rate of opening de-creases. On later serial number aircraft, thetime-delay relays have been removed. The fullopening time in both cases is approximatelytwo minutes.

Heating is a function of compression, bleed-ing and mixing of the air supplies from the en

COMPONENT

Heat ExchangTwo two-pass air-to-

closed in a common primary and main. Treduces the temperato directing it to thecompressor. The funchanger is to removinduced by the ACMrecting the air mass

T rbofan

Figure 11-5. Heat

R ev i s i on 3


O

NL

ECUOVHT

NO. 1 ENGINEPYLONAIR INLET

PRIMARY HEATEXCHANGER T

INDICATIN

+4

SECONDS

• NOSEDOW

AN• BYPA

NOT

A1

COND'GCREW

DISC BRAKENOT APPLIEDLEGEND

BLEED AIR



1 1 - 8 A

LY

Figure 11-6A. Turbofan Operat

DEENEBRA

LOW PRESSUREAIR INLET FLAP

OPEN (GROUND)

HOT AIR FROMTEMPERATURE

CONTROL VALVES

TEMPERED AIR

BLEED AIR

ELECTRICAL POWER

COOLED AIR

AMBIENT AIR

STATIC AIR

1 1 - 8 B

F OR T RAI NI N G

P URP O S E S

ONL Y

NO. 1 ENGINEPYLON

AIR INLETPRIMARY HE

EXCHANGER

• NO

DO

• BY

NO

A1

CON

CRE

DISC BRA

APPL

LEGEND

TEMPERED AIR

BLEED AIR

COOLED AIR

AMBIENT AIRSTATIC AIR

+4

SECONDS



R ev i s i on 3 . 0 1

E

HOT AIR FRO

TEMPERATUCONTROL VAL

LOW PRESSURE

AIR INLET FLAP

CLOSED (FLIGHT)

ELECTRICAL POWER

STATIC AIR

Figure 11-6B. Turbofan Op


is to assist cooling when on the ground, or inflight, in any one of the following configurations:main landing gear extended, or slats extended, orTAS less than 300 knots. An electropneumaticbrake terminates operation of the turbofan inclean configuration flight when TAS is 300 knotsor greater. A bypass valve opens to bypass theturbofan during clean configuration flight of 300TAS or greater. The input signals for turbofan andbypass valve control, received from landing gearand slat position and from the ADCs (air datacomputers) representing TAS, are processed by aPCB and transmitted to the brake control valveand to the bypass valve.

OperationWhen the airplane is on the ground or in flightwith the main landing gear extended, or theslats extended, or TAS less than 300 knots, theturbofan brake solenoid control valve (Figure11-6A) is deenergized and closed, isolatingbleed-air pressure from the brake. The brake as-sembly then vents to the atmosphere and is re-leased. Hot bleed air is directed through theprimary heat exchanger, and since the turbofanb l i i d l d h d

Figure 1

Fig re 11 8



bypass valve is energized closed, the temperedair is directed to the turbine of the turbofan. Thepneumatic energy is converted to rotary me-chanical motion by the turbine and is used todrive the integral fan. The low pressure createdby the fan opens a low-pressure door on thelower side of the fuselage (Figure 11-7), per-mitting the fan to induce a large cooling airflowacross the heat exchangers and discharge itthrough a grilled outlet (Figure 11-8) alsolocated on the lower side of the fuselage. Thecooling by heat transfer at the heat exchangersand energy conversion by the turbine greatlyreduce bleed-air temperature.

The tempered air from the turbofan is directedto the turbocooler for further cooling. Thiscondition prevails during ground operation

pressure frothe isolationapply the brturbofan assthe heat exch

through the i1 engine pyl

Turbocoo

General

The turbocoturbocompre

the single sta

Figure 11-8.




Operation

Hot bleed air from the crew and passengertemperature regulating valves is tempered bythe primary heat exchanger. After passingthrough the turbofan during ground opera-

tion or during slow flight configuration

across the heat exchrected through a higtor to the turbocoolthe pneumatic enemotion and drives thwhile effecting both

sure decrease

Figure 11-9. Turbocooler

mixing ducts and two-way ducts, through acold air pressure-regulating valve (PRV), tosupply the final conditioned air for the crewand the passenger systems. The cold air PRVis fully automatic to maintain the gasper ductpressure at approximately 0.2 to 0.36 psi abovecabin pressure. The PRV is almost entirelyclosed at high altitude since the requirementfor cold air is at a minimum, thus maintain-ing a slight positive pressure and flow in thecold air (gasper) duct.

The efficiency of the turbocooler is deter-mined by the positions of the passenger andcrew temperature control valves. When in amaximum cooling mode, the major portion of the bleed air is directed through the cold sideof the temperature control valves, and the tur-bocooler is at maximum efficiency. This effi-ciency is reduced as the temperature controlvalves cycle toward heat; then the greater por-tion of the bleed air bypasses the turbocoolerand is directed to the passenger and crew mix-ing jet pumps.

Water Separators

exchanger. Wtemperature signal transmlates the antiing a volumeoutlet, increaperature andpoint (approincorporatin(turbine emestream of themergency aensures emefor turbine ain the case wis blocked inlight (ECU/Athe center inthe emergenton light is iis closed.

OVERHE

SYSTEMThe overhea




Water SeparatorsThe high-pressure and low-pressure afterseparators (Figure 11-12) are conventionalswirl-action units; they remove sufficientwater from the air to maintain a relativelycomfortable humidity level in the occupiedareas. The water removed by the se parators

is collected in sumps where it is exposed toa jet pump operated by a motive flow suppliedby the turbine outlet. The jet pump sprays thecollected water across the heat exchangers,increasing their cooling efficiency.

ANTI-ICING SYSTEM

Since the temperature drop across the turbo-

cooler may cause freezing at the low pressure

The overheasensor and anbleed-air ovsor is locateoutlet duct. Wor exceeds tdetection chnose gear is

tric valve is OVHT” on tcomes on. Ilgear and vafour secondsground whila signal to thof the APUlight, a magn

panel will in

1 1 - 1 2

F OR T RAI NI N G

P URP O S E S

ONL Y

LP WATER

SEPARATOR

ANTI-ICINGSENSOR

HP WATERSEPARATOR


DUCT

CREWCOLD AIR

CABIN

COLD AIROR

SUCTION

COLD AIRPRESSURE-

REGULATINGVALVE

COLD AIROR

SUCTION

CABIN


DUCT

LOW-PRESSURE

AIR INLET

ANTI-ICINGVALVE

BCCEV(S

ECUOVERHEATSENSOR

COMPRESSOR

TURBOCOOLER

TURBINE



R ev i s

i on4 . 0 1

LEGEND

BLEED AIRTURBOCOOLERCOMPRESSED AIR

CONDITIONED

TEMPERED AIR TURBOCOOLERTURBINE AIR RAM AIR


WATERINJECTOR

MAIN HEATEXCHANGER

PRIMARY HEATEXCHANGER

TURBEMERGANTI-IELEC

VAL(SB F90

Figure 11-10. Normal Operation—

F OR T RA


O

NL Y


LP WATERSEPARATOR


HP WATERSEPARATION


CIRCUIT

CREWCOLD AIR

CABIN

COLD AIROR

SUCTION

COLD AIRPRESSURE-

REGULATINGVALVE

CABIN

STATIC AIRINLET

ECUOVERHEATINTEMPERATU

SENSOR

COMPRESSOR

TURBO-COOLER

TURBINE

BCCE

V(

COLD AIROR

SUCTION



1 1 - 1 3

LEGEND

BLEED AIRTURBOCOOLERCOMPRESSED AIR

RECIRCULAT


CONDITIONE


CIRCUIT


WATERINJECTOR

MAIN HEATEXCHANGER

PRIMARY HEAT

EXCHANGER

ANTI-ICINGVALVE

TUREMERANTELE

VA(SB F

Figure 11-11. Normal Ope


DISTRIBUTION SYSTEM

GeneralSeparate conditioned air distribution systems

in the rear bulkheadof this air supply is

ger jet pump (Circlereceives motive flowpasses the cooling

COLDAIR

INLET

VORTEX

COALESCOR RELIEF VALVE

Figure 11-12. Water Separator (Typical)



Separate conditioned air distribution systems(Figure 11-13 on aircraft SNs 1 to 162 or Figure11-13A for aircraft SNs 163 and subsequent)are provided for the passenger and the crew sys-tems. The ducts for each system consist of conditioned air ducts and cold air ducts. Check valves are installed in all ducts entering through

the rear pressure bulkhead.

Passenger Ducts

The passenger conditioned air ducts are routedin the lower side of the left and right cabin con-soles. These ducts are interconnected; in ad-dition, they supply ducts which provide floorheat, toilet heat, and warm air distribution for

fli ht t l d i t l t

passes the cooling crew conditioned anected to allow eithsupply both ductson aircraft prior to lever (Figure 11-1“NORMAL” and “

copilot’s console.

Aircraft SN 163 anterconnect valve loFR 22 which enableconditioning systeof them fail. This imally closed by a terfly valve.



COPILOTGASPER

FLAPPERVALVE

PASSENGERDOOR

HEATING

TWO-WAYMANIFOLDS

INTER-CONNECT

FLAP(SNs PRIOR TO 163)

CREWCONDITIONED

AIR DUCT

CABINFLOORHEAT

PASSENGER TEMPERATURSENSOR/AMPLIFIER

EFISCOLD AIR PILOT

GASPER

CREWTEMPERATURE

SENSOR

FLAPPERVALVE

TWO-WAYDUCT

FRONT GROUNDCONDITIONING

FLAP

PASSENGERCONDITIONED

AIR DUCT

CABINFLOOR

HEATINGDUCTS

LAVATORYHEATING

PILOTS’FOOTWARMERS

NOSE CONEISOLATION

VALVE(SNs PRIOR TO 69)

EFISBLOWERS

NOSECONE

BLOWER

DEFOGOUTLETS


MANUALDISTRIBUTOR

CONTROL

AIR SOURCES:1. COOLED AIR FROM TURBOCOOLER2. BLEED AIR TO CREW JET PUMP

3. RECIRCULATED AIR IN TW4. BLEED AIR TO PASSENGE

GASPER PICK-OFF

THIRD CREW MEMBER


PASSENGERGASPERS

Figure 11-13. Distribution System




FALFALCON 900 PILOT TRAINING MANUAL



1

3

45

6

8

9

13

14

19

20

21

22

23

24

25

26

29

30

31

FLOW LIMITERS

HEAT EXCHANGER

TURBO COOLERHP WATER SEPARATOR

LP WATER SEPARATOR

ATOMIZERS

GASPER PRESSURE REGULATING VALVE

CABIN GASPERS

PILOT AND COPILOT GASPERS

COCKPIT VENTURI

RH “TWO-WAY” DUCT

FLAPPER VALVE

MANUAL COCKPIT AIR DISTRIBUTOR

CABIN VENTURI

LH “TWO-WAY” DUCT

FLAPPER VALVE

CABIN CONDITIONED AIR DUCT

INTERCONNECTION FLAP

CABIN SYSTEM OZONE CATALYSER

COCKPIT SYSTEM OZONE CATALYSER

(3HC)

(4HC)

(5HC)(6HC)

(22HC)

(23HC)

(24HC)

(25HC)

(3HN2)

(3HN3)

(22HN)

(23HN)

(43HN)

(44HN)

(25HP)

(3HQ)

(14HW)

COCKPIT TEMPERATURE SENSOR AMPLIFIER

COCKPIT TEMPERATURE REGULATING DUAL ELECT

COCKPIT DUCT SENSORCOCKPIT OVERHEATING THERMAL SWITCH

CABIN TEMPERATURE SENSOR AMPLIFIER

CABIN TEMPERATURE REGULATING DUAL ELECTRIC

CABIN DUCT SENSOR

COCKPIT OVERHEATING THERMAL SWITCH

CABIN AIR-CONDITIONNING ELECTRIC VALVE

COCKPIT AIR-CONDITIONNING ELECTRIC VALVE

TURBOFAN BYPASS ELECTRIC VALVE

TURBOFAN

DUCT SENSOR

TURBINE ANTI-ICING ELECTRIC VALVE

BAGGAGE COMPARTMENT DISTRIBUTION ELECTRIC

NOSE CONE BLOWER

ECU OVERHEAT TEMPERATURE PROBE

0 4 7 14 15 20

(3HQ)

DIFFUSERS

14 22 21

COCKPIT AIR-

CONDITIONING

CIRCUIT TOILET 13 1320 (22HC) (6HC) (5HC) 29 19P

REA

TOIL

HEAT

(24HC)(25HC)13FLOOR HEATING

MANIFOLD

26

CABIN

AIR-CONDITIONING

SYSTEM

2413COLD AIR25(3HC)22PILOT FEET

WARMING

MANIFOLD

GLARESHIELD

MANIFOLD

14

LEGEND

Figure 11-13A. Distribution System—Aircraft SN 163 and Subsequent






EFIS CoolingThe EFIS componenare cooled by an airf

duct, separately diiters to the pilot’s C i rcu la t io n o f thcontinuously operaglareshield. These from the DC bus breakers.

Battery CooliIn addition to theirChapter 2, “Electrimain batteries in thecooled on the grouCold air is supplied stream of the LP wrectly fed to the celectric valve.

With the aircraft on TER APU” switch valve is controlled byWhen the switch issition, the electric minate an amber ligin a fully OPEN po

Baggage ComConditioning

The baggage comthrough a branch duline (Figure 11-10)venturi, an isolatioControl is achieved

on the overhead swibeled “BAG.” Thebeled “NORM,” “His scribed across th

When the switch is isolation valves areto the venturi, wherair. The resulting floor level. This flothrough the isolationto the HEAT positioand flow to the bag

Crew DuctsThe crew conditioned air duct is routed alongthe right side of the fuselage and supplies con-

ditioned air to the entrance area; then it isrouted along the left and right sides of thecockpit to supply air to the aft cockpit windows,prior to SNs 116, to the footwarmer systems,and for forward window defogging. Each pilotcontrols the direction of air supply (for de-fogging and footwarmers) by levers (Figure 11-15) located on the outboard s ide of therespective instrument panels. A scribed whiteline above the lever slot has up and down arrowheads to determine the desired direction of flow. This lever is a distributor control andmay be positioned at any point between thetwo extremes.

Gasper DuctsThe cold (gasper) duct (Circle 1 in Figure 11-13) is a three-branch duct: one branch sup-plies the crew gaspers, and the other twobranches supply the gaspers on the left andright sides of the cabin. The cold air for this

duct is supplied directly from the ACM turbinethrough the low-pressure water separator. Anadditional branch from the ACM turbine routescold air through the gasper pressure regulator

(PRV) and divides into two ducts, which con-nect to the passenger and crew jet pumps andtwo-way recirculating ducts. These ducts aredual-function; when cooling demand is low, thegasper PRV is closed, and the jet pumps inducerecirculation through the ducts and distributethe recirculated air into the passenger and cr ewconditioned air ducts while some cold air stillcontinues to flow through the gaspers. On theother hand, when cooling demand is high, thePRV is open, and cold air is supplied throughthe recirculating ducts. The flappers at the out-let ends of these ducts are pressure- and/orsuction-operated as dictated by the cooling orheating demands of the system.

Figure 11-14. Duct Interconnect Flap“COND” Lever Figure 11-15. Footwarmer and

Defogging ValveLever (Typical)



cockpit. The inputs from these sensors aresupplied to the appropriate section of the dualcontroller and are used as anticipatory infor-mation, which prevents undershooting in thesystem.

ControllerThe temperature controller (Figure 11-16) is adual unit labeled “PASSENGER” and “CREW.”It contains separate circuitry for passenger andcrew temperature control and is mounted on thelower left side of the center instrument panel.Each controller consists of a mode selectorswitch with positions labeled “AUTO” and

“MANUAL”; the passenger mode selector hasan additional position labeled “REMOTE” topermit transfer of cabin temperature controlto an optional control panel located in any de-sired position in the cabin. The crew mode se-lector operates in conjunction with a switchlabeled “NORM” and “EMERG” located onthe PRESSURIZATION control panel (Figure11-17). This switch must be in the NORM po-

sition for normal operation of the passenger con-ditioning valve and crew temperature control.The functions of this switch are described in de-

UAL, the automatictem is inhibited, andto the manual tempcated below the auttor. This switch has“HOT,” and an unlwhich it is spring-loto the COLD or HOTselector is at MANous output signal toperature regulating

Separate valve posifor the passenger anvalves by indicatorsections. The indicatentiometers on thebrated; the needle position on the left position on the extrprovide valve positAUTO and MANUAPASSENGER controgage calibrated in dvide indication of c

Temperature C




tail in Chapter 12, “Pressurization,” in thistraining manual.

When the mode selector switches are at theAUTO position, the automatic temperaturecontrol system is armed. Automatic tempera-

ture selection is achieved by a rotary (rheostat)selector, which has approximately 180° of movement between the 9 o’clock and 3 o’clock positions. When the mode selector is at MAN-

Temperature C

The temperature cosenger and crew syssign, differ in opconsists of two buttlinked, and operate

a single motor. The meither the automaticontrol systems of In the AUTO mode,

by PCBs (in the hot direction) to 55% of fullyhot. The hot and cold extremes can be achievedin the MANUAL mode. However, followingpreconditioning on the ground using MAN-

UAL mode, the PCBs automatically cycle thevalves to 60% of fully hot when the mode se-lectors are retur ned to the AUTO positions, re-ducing air noise level in the cabin.

When the passenger temperature control valveis at the fully hot position, its cold butterflyvalve is not fully closed; therefore, a volumeof air is allowed to pass through the ECU tosupply gasper air. The crew temperature con-trol valve, in conjunction with the pressur-ization emergency switc h, is used to supply anemergency air source to the occupied areas.This operation is described in detail in Chapter12, “Pressurization,” of this training manual.

Each temperature control valve receives itsbleed-air supply from the associated air-con-

ditioning valve; then, as determ ined by the hotand cold butterfly valves in each temperaturecontrol valve assembly, the air mass througheach valve is directed into two ducts; one sup-plies the hot-air flow to the associated mixing

jet pump (ven turi ) in the pass enge r an d cr ewdistribution systems, and the other duct fromeach temperature control valve supplies an airmass to a common duct, which directs itthrough the ECU assembly for cooling.

INDICATION

A light on the warning panel labeled “COND’GOVHT” (Appendix B) comes on to indicatethat the temperature in either the passenger orcrew conditioned air duct is 95°C or greater.

The signals to this light are supplied from ther-mal switches located in the passenger and crewconditioned air duct.

OPERATION

In the automatic mode, the temperature con-trol systems (Figure 11-18) operate on theelectronic bridge principle. This bridge hasthree legs: one supplies the input representingthe actual duct temperature (from the ducttemperature sensor); the second supplies the

input representing the actual area tempera-ture (from the passenger or crew temperaturesensor); and the third is the desired or selectedtemperature which is the input signal from

the automatic temperature (rheostat) selectorresulting from crew selection. These signalsare compared by the bridge circuitry, and theerror is determined, which results in one of twopulsating output signals: a positive (hot) ornegative (cold) signal to the appropriate valvemotor. The magnitude of the pulses is deter-mined by the error difference; each pulse isshorter than the preceding one until the bridgecircuit is again balanced. Once balanced, anychange in duct or area temperatures or repo-sitioning of the automatic temperature selec-tor results in a repeat performance.

Moving the mode selector to MANUAL com-pletely inhibits the automatic temperature con-trol system, and the temperature control valveposition is a sole function of the COLD and

HOT positions of the manual temperature se-lector switch. As long as the switch is held ineither position, the associated valve receives acontinuous signal in the cold or hot directions,as selected. When using manual temperaturecontrol, all changes should be made in small in-crements, allowing the system to stabilize be-tween selections. Reference should also be madeto the associated valve position indicator.

NORMAL OPERATION—FLIGHT (HEATING)

During normal flight in the medium to highpower range, the bleed-air supply is as shownin Figure 11-2. LP bleed air is supplied to themanifold from the LP compressor of the No. 1engine. The No. 3 engine supplies the HP/LPmixed air from its compressors to the manif old,as determined by the a ssociated PRV. The No.2 engine provides an HP/LP mixed air supplyto the manifold through the normally open iso-lation valve, in addition to a direct supply to thepassenger air-conditioning valve; both sup-plies are as dictated by the No. 2 engine PRV.The air from the manifold is directed to thecrew air-conditioning valve.




CONDOVH

CREWSENSOR

AMPLIFIER

CABIN CONTROLS (OPTIO

COCKPIT

INTER-CONNECT

FLAP

AU



The passenger and crew air-conditioningvalves are as shown in Figure 11-11 since itis assumed that the PASSENGER and CREWswitches (Figure 11-1) are at ei ther the AUTOor the ON positions. The demand for heat-ing as determined by the temperature con-trollers positions the passenger and crew

temperature control valves so that the greatervolume of bleed air is passing through the hotport of the valves and is supplied directly to therespective jet pumps. The remaining flowpasses through the cold ports of the valves intoa common duct. This airflow is directed throughthe primary heat exchanger and through theopen turbofan bypass valve (clean flight con-dition and TAS greater than 300 knots) to theturbocooler turbine. The resulting outlet tem-perature from the turbine is a function of tur-bocooler efficiency at any given time.

This airflow is directed to the crew and pas-senger gaspers. Figure 11-19 depicts the finalconditioning and distribution in the occupiedareas. The low pressure at the jet pumps in-duces recirculation through the two-way ducts,determining the conditioned air temperaturein the passenger and crew conditioned airducts The gasper airflow is determined by

temperature contrnormally open air-c11-10 depicts the assumes that TAS isfore, the turbofan bvalve is closed. The are positioned by th

major volume of blthe cold ports intoprimary heat exchanis closed, the turbofaefficiency, resultindrop across the uniair to the turbocooheat exchanger, andThe turbocooler is oficiency, resulting ilet. The water separmodulate toward opperature sensor dowarator. The warm scroll prevents ice farator. The hot air hot ports of the teinduces mixing at tthe distribution depithere is a high demaPRV is open the hig




ducts. The gasper airflow is determined bythe gasper PRV.

NORMAL OPERATION—

FLIGHT (COOLING)Operation in the lower altitude range in cleanconfiguration flight increase the demand forcooling rather than heating; in addition, me-teorological conditions may require the use of sur face ant i - ic ing systems. Under suchconditions the bleed-air supply may be asshown in Figure 11-3. All engines in this case

l i HP/LP i d i h if ld

PRV is open, the higtwo-way recirculatvalves close, resultinthese ducts as well a




COPILOTGASPER

FLAPPERVALVE

PASSENGERDOOR

HEATING

TWO-WAYMANIFOLDS

PASSENGERGASPERS

CREWCONDITIONED

AIR DUCT

CABINFLOORHEAT

PASSENGER TEMPERASENSOR/AMPLIFIE

EFISCOLD AIR

PILOTGASPER

CREWTEMPERATURE

SENSOR

FLAPPERVALVE

TWO-WAYDUCT


FLAP


AIR DUCT

CABINFLOOR

HEATINGDUCTS

L


NOSE CONEISOLATION

VALVE(PRIOR TO SNs 69)

EFISBLOWERS

NOSECONE

BLOWER

DEFOGOUTLETS


MANUALDISTRIBUTOR

CONTROL


3. RECIRCULATED AIR 4. BLEED AIR TO PASS

LEGEND

HOT AIR

COOLED AIR

CONDITIONED AIR (HEATED)

RECIRCULATING AIR


GASPER PICK-OFF

THIRD CREW MEMBER

INTER-CONNECT


Figure 11-19. Distribution System Normal Operation—Flight (Heat






FAL

COPILOTGASPER

FLAPPERVALVE

PASSENGERDOOR

HEATING

TWO-WAYMANIFOLDS

PASSENGERGASPERS

CREWCONDITIONED

AIR DUCT

CABINFLOORHEAT

PASSENGER TEMPERATURESENSOR/AMPLIFIER

CTHESW

EFISCOLD AIR

PILOTGASPER

CREWTEMPERATURE

SENSOR

FLAPPERVALVE

TWO-WAYDUCT


FLAP


AIR DUCT

CABINFLOOR

HEATING

DUCTS

LAVATORYHEATING

P


EFIS

BLOWERS

NOSECONE

BLOWER

DEFOGOUTLETS


MANUALDISTRIBUTOR

CONTROL


3. COLD AIR FROM GASPER PRESSUR4. HOT AIR TO PASSENGER JET PUMP

LEGENDHOT AIR

COLD AIR CONDITIONED AIR (COOLED)


GASPER PICK-OFF

THIRD CREW MEMBER

INTER-CONNECT


NOSE CONEISOLATION

VALVE(PRIOR TO SN 69)

Figure 11-20. Distribution System Normal Operation—Flight (Cooling)



GROUND OPERATION

Except for the air source, ground and flight op-eration of the air-conditioning system is sim-ilar. Figure 11-4 depicts ground operationwith the APU running and supplying bleed

air. The APU bleed-air supply is a function of the APU bleed-air valve, as determined by theAPU switch on the BLEED AIR section of the overhead switch panel (Figure 11-1) andproximity switches in each power lever quad-rant. In addition, the valve modulates under theinfluence of the APU ECU signals initiated byT5 inputs to the ECU. The APU bleed-air valvecloses when any power lever is advanced to or

beyond 54° PLA, whether or not the engine isrunning. An ECU overheat causes the APU toautomatically shut down.

The APU bleed air is supplied directly to thepassenger air-conditioning valve and to thecrew air-conditioning valve through the nor-mally open isolation valve. From this pointon, the operation of the air conditioning is as

described for slow or dirty configurationflight in Figure 11-10, in conjunction withFigures 11-19 and 11-20.




LIMITATIONS

All limitations contained in the AFM pertain-ing to the air-conditioning system and relatedsystems are binding under law, and they mustbe observed regardless of the type of operation.

1. HP bleed air from the No. 1 engine isavailable when the HP 1 switch is inAUTO and the:

A. WING anti-ice switch is on.

B. All power levers are less than 54°PLA.

C. S-duct anti-ice switch is on.

D. Selections in A or C are made.

2. When the isolation switch on the BLEEDAIR panel is at the ISOL position:

A. No. 2 engine bleed air is shut off.

B. No. 1 and No. 3 engine bleed air isshut off.

C. No. 2 engine bleed air is isolatedfrom the manifold.

D. No. 2 engine bleed air is isolatedfrom the wing anti-icing system.

3. Moving the isolation switch to theISOL position when the APU bleed airis selected:

A. Causes the APU to flame out

B I l t th APU bl d i f th

6. The conditionigage compartm

A. Gasper duc

B. Recirculati

C. Crew hot-aD. Passenger h

7. Cooling air forplied from the

A. Footwarme

B. Crew gaspe

C. Nose cone

D. EFIS blowe

8. The interconnecopilot’s consothe:

A. Crew and p

B. Left and ri

C. Passenger

cone ductD. Passenger

ducts


QUESTIONS



B. Isolates the APU bleed air from thecrew ducts

C. Causes the APU bleed-air valve toclose

D. Isolates APU bleed air from thepassenger ducts

4. If the ECU OVHT light comes on dur-ing ground air conditioning using theAPU bleed-air, the:

A. Turbofan bypass valve opens.

B. Air-conditioning valves close.

C. Turbofan brake releases.

D. APU automatically shuts down.

9. The temperatuture controller

A. Passenger t

B. Crew duct

C. Passenger dD. Crew temp

10. The AUTO poperature modewhen the:

A. CREW swipanel is at

CHAPTER 12

PRESSURIZATION

CONTENTS

INTRODUCTION......................................................................

GENERAL .................................................................................

PRESSURIZED AREAS............................................................

AIR SOURCES ..........................................................................

CABIN PRESSURIZATION......................................................

General................................................................................

Major Components .............................................................


Operation ............................................................................




Protection............................................................................

Depressurization .................................................................

BAGGAGE COMPARTMENT PRESSURIZATION..............

General................................................................................

Air Sources .........................................................................

Control ................................................................................

Protection............................................................................

Control............................................................................................

Protection........................................................................................

Indication........................................................................................

Operation ........................................................................................

LIMITATIONS......................................................................................

QUESTIONS.........................................................................................




ILLUSTRATIONS

Figure Title

12-1 Pressurized Areas ........................................................

12-2 Emergency Pressurization Control Panel....................

12-3 Air Source—Emergency Pressurization......................

12-4 Automatic Controller ..................................................

12-4A Principle of Automatic Cabin Pressurization —PROG

12-5 Outflow Valves............................................................

12-6 Triple Indicator............................................................

12-7 HORN SIL Switch ......................................................

12-8 Overhead Switch Panel ..............................................

12-9 Pressurization Operation—Auto Mode Ground .........

12-10 Pressurization Operation—Auto Mode Flight ............

12-11 Pressurization Operation—Manual Mode Flight........




12-12 Baggage Compartment Pressurization Operation.......

12-13 Nose Cone Isolation Valve Lever (SNs Prior to 70) ...

12-14 Nose Cone Ventilation/Pressurization Operation........

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;


PRE



INTRODUCTION

This chapter describes the pressurization system of the F

values, relative to pressure and flow rates, are used only fActual values must be obtained from the approved manwith the airplane.

GENERAL

Air conditioning and pressurization are syn-onymous terms. The term air conditioning ap-

pressure chanerating enve

AIR SOURCES

A normal and an emergency air source areused for pressurization. The normal source of air is supplied by the LP compressors of thethree engines, automatically supplemented asrequired by HP bleed air. For more informa-

tion on the normal air sources see Chapter 9,“Pneumat ics” and Chap te r 11 , “Ai rConditioning,” in this training manual.

Figure 12-3 depictization air source. Tvalve is closed; thetemperature controlits hot port is fully rected to the cockpThe only condition

mixing of ambient cby the jet pump. Thigher altitudes mahowever, when alti


Figure 12-1. Pressurized Areas

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ; ; ; ; ; ; ;

; ;

; ;

; ;

; ;

; ;

LEGEND

NOSE CONE

OCCUPIED AREAS

BAGGAGE AREA



An emergency source of air is required if theECU malfunctions. This airflow is controlledby a two-position switch on the emergencypressurization control panel (Figure 12-2).The switch positions are labeled “NORM”

and “EMERG.” Moving this switch to theEMERG position simultaneously closes thecabin air-conditioning valve, drives the cock-pit temperature control valve to the fully hotposition, and arms the cockpit manual tem-perature control switch. Airflow to the ECUis now shut off.

ture will increase imay be controlledCOLD position of perature control swalso reduce the airf

areas and conseque

CABINPRESSURI

GENERAL

any isobaric range. The maximum positivedifferential is also automatically limited to 9.6psi (662 mb). Negative pressure in the pres-surized area is prevented by a low differen-tial limit of 0.29 psi (20 mb). In case of lossof inflow air the maximum cabin altitude islimited to approximately 14,500 feet. A 100%

redundancy is incorporated for all of thesesafety features. The manual pressurizationcontrol system is completely independent of electrical power.

MAJOR COMPONENTS

GeneralThe major components of the pressurizationsystem include the following:

• Autom

• Manua

• Emerg

• Main o

• Emerg

• Jet pum

Automat

The automatcated on thestrument panA1, is a digimemory, capsurization symat ic modoperation.


HP WATER

SEPARATION

ANTI-ICING

DUCT SENSOR

LP WATER

SEPARATOR

TURBINE

TURBO-

COOLER

COMPRESSOR

ECU OVERHEATINGTEMPERATURE

SENSOR

ANTI-ICING

BATTERY COMPARTMEN

COLD AIR SUPPL

ELECTRICAL VALV

(SB F-900-125



PASSENGER

PASSENGER

COLD AIR

COLD AIR

OR

SUCTION

COLD AIR

OR SUCTION COLD AIR

PRESSURE-

REGULATING

VALVE

CABIN

COLD AIR

CREW COLD

AIR

CREW-AIR

CONDITIONING

CIRCUIT

VALVE

CABINPRESSURIZED

The face of the automatic controller has twoliquid crystal displays (LCD): one labeled“LAND ELV” and the other “QNH.” A three-position mode select switch on the upper right

of the face is labeled “PROG” (program), “FL”(flight level), and “LDG” (landing) and isused to select the automatic and semiauto-matic modes, in conjunction with a three-

i i i h h l f f h f

destination or landiby the associated ±of –1,500 to +14,00

The two-position swthe face is labeled “m(inches of mercury)

sired digital displaysure in the QNH winswitch on the lowefield barometric pre

A push-button switlower center of the lidity of the selectethe mode select swi

pushing the TEST sin the altitude displathe test display wiously the QNH dispor in. Hg window insystem can be usedit is inhibited by thlanding gear touchdure be detected in th

must be retested in the test is now validusing these two mo


Figure 12-4. Automatic Controller



position switch on the upper left of the face.

The switch has + and – positions and is springloaded to the center position. When the modeselect switch is at PROG, the associated ±

switch will change the digital display in in-creasing or decreasing increments when heldat the + or – positions. The digital displaychange in response to switch selection beginsat a slow rate and then increases in proportionto the time the switch is held.

For ease of display setting, move the switchintermittently in the desired direction. Since

The display windotroller going blankmalfunction of thematic modes.

The “LAND Etomatic controPROG or LDGbut it will notmode is selecte

NOTE

The ± switch for the QNH displaychanges it in increments of 0.01 in.Hg between 28.00 and 31.00 in. Hgwhen in. Hg is selected, and in in-crements of 1 mbar between 950 and

1,050 mb when mb is selected.

The automatic controller senses three param-eters in the PROG mode: absolute cabin pres-sure, airplane altitude, and airplane rate of change. The cabin pressure is sensed througha port on the body of the unit. Altitude and rateof change information is supplied from ADC1 and ADC 2. The preprogrammed function is

inhibited in the FL and LDG modes, and thesystem then operates from the only availableparameter, absolute cabin pressure.

In the PROG mode the automatic controller isconstantly monitoring the actual cabin altitude,the destination field elevation, the landinggear position, power lever position, and ratesof climb and descent, comparing these values

in the computer’s memory, and then produc-ing output signals which ensure that the cabinaltitude will approach the theoretical cabinaltitude at a theoretical rate which will al-

The FL modein the PROGdication. Thinto the contrminimum calevel and ca

The pilot mflight level ential systempressure chathe airplanethe pilot muand set the Q

The lanserted ifield. Ta preprtouchdo

When oimportathe LDG




altitude at a theoretical rate which will always remain within the established limits of +650 feet per minute in climb and –450 feetper minute in descent (Figure 12-4A).

The system will automatically maintain thelowest possible cabin altitude relative to theairplane’s altitude and cabin pressure differ-ential. Rates of pressure change will auto-matically be maintained as low as possiblefor all flight profiles and will always remainwithin the programmed limits.

The system also has the capability of initiatinga prepressurization program during takeoff,

scent wsufficiebelow lthe maxThe LD

mode seequal toting thedisplayo n t h esetting be usedabove 8


THIS PAGE INTENTIONALLY LEFT BLAN




C

T

D

E

F

G

H




EmergencyPressurization Control

The emergency pressurization control panel(Figure 12-2) includes a two-position mode se-lector switch labeled “AUTO” and “MAN,” atwo-position air source selector switch la-

beled “NORM” and “EMERG,” a two-positionguarded switch labeled “DUMP,” and a rotarymanual controller labeled “DN” and “UP.”

With the mode selector switch at AUTO, theair source selector switch at NORM, theDUMP switch at the guarded position, andthe manual controller rotated (counterclock-wise) to the DN position, the pressurizationsystem is preset for automatic operation.

Manual Controller

Placing the mode selector switch to the MANposition inhibits automatic operation. A cabinrate of descent will also be indicated on thetriple indicator. The manual controller per-

mits the pilot to control the rate of climb, upto a maximum of +1,500 feet per minute whenrotated to the UP position or, +2,500 feet perminute when moved beyond the UP positionto the hard stop. The maximum rate of de-

Main Outflow The main outflow matic poppet valvethe rear pressure bsists of a control chactuator chamber a

ble diaphragm connA spring in the contfail-safe closed pos

The pressure in themined by a quadramotor in response tautomatic controlleopens two nozzles:

to the control chambward close, and thecontrol chamber toreducing control ppoppet toward open

The main outflow vincludes a cabin alvalve. The function

is to operate in respautomatic controllecabin at a rate whigrammed cabin alticlimb and descent



to the hard stop. The maximum rate of descent is a preset value of –1,500 feet per minutewhen the manual controller is set to the DN po-sition. A constant cabin pressure may beachieved by adjusting the manual controller

until the cabin rate indicator stabilizes at 0.

The manual controller must be at theDN position when operating in thePROG, FL, or LDG modes. If it is re-

climb and descentmatic controller.

Emergency O

The emergency outflar in design and opevalve. It is normallbut it does have a solclosing for takeoffbased on the differesure and cabin prepneumatic relay.

CAUTION


EMERGENCYPRESSURIZATION

CONTROLPANEL

FWD

Figure 12-5. Outflow Valves



The control chambers of both valves are in-terconnected by a sense line; therefore, dur-ing automatic operation the emergency outflowvalve is slaved to the main outflow valve, andduring emergency operation the main outflowvalve is slaved to the emergency outflow valve.

Jet PumpA jet pump supplied with a motive flow from aline supplied by No. 1 and No. 2 engine’s HPbleed-air lines, or from the APU bleed air when

INDICAT

Cabin altituferential, and

all indicated12-6) locatedter instrumeis calibratedThe scale is cand 10,000 asweep needl

The differen

A cabin pressure rate-of-change scale is lo-

cated on the left side of the dial. It is calibrated

in feet per minute times 1,000 from 0 to 2, UP

and DN. The prevailing rate is indicated by a

needle.

A light labeled “CABIN” (Appendix B) on the

warning panel will come on in conjunction

with a continuous synthetic voice (“CABIN”)

OPERATION

General

Operation of the pre

described in three m

auto mode flight, an

During the preflight

firmed that the manu

mal/green) DN po

switch is at AUTO, th

is at NORM, and th

guarded position (F

The flight planned latered in the automat

in feet in the LCD. T

and the LCDs react

fying the validity o

matic operation (Fig

running, the HP 1, P

are set at auto, the PA

conditioning switche(Figure 12-8), and th


Figure 12-6. Triple Indicator



y ( )

either when the airplane is weight on wheels and

the TEST switch on the automatic controller is

pushed or when the airplane is in flight in any

pressurization mode and cabin altitude exceeds

10,000 feet. The aural warning may be silencedby the HORN SIL switch on the center pedestal

(Figure 12-7).

Advancing the power levers to takeoff powerwith weight on wheels will complete circuitsto rapidly close the cabin and cockpit air-conditioning valves, isolating all bleed airfrom the occupied areas. The HP1 valve willbe closed; the PRV2 and PRV3 valves willalso close for approximately 18 seconds.

Simultaneously, the master controller willtransmit an output signal to the main outflowvalve torque motor, which will increase con-trol pressure and close the valve. A circuit isalso completed to the solenoid valve on theemergency outflow valve. The energizedsolenoid admits cabin pressure to the controlchamber of the emergency outflow valve forrapid closing.

Figure 12-9 depicts pressurization in the automode ground configuration. Bleed air is off andthe cabin is sealed until after lift-off.

Auto Mode Flight

When the landing gear proximity switches as-sume the flight position, circuits are com-

pleted to slowly open the cabin and cockpitair-conditioning valves, permitting airflow tothe occupied areas. The automatic controllerassumes the PROG mode (Figure 12-10) . Itsreaction will be determined by one of the fol

cabin altitureaches the pthe minimumthroughout ttude, the lanbient differe

No further throughout t

Before landi(QNH) is to

When the airpower leverwill initiate

mode, and thtion at a max

To ensuin the cments, fand bag

These fnected tof the reopeningand equ




reaction will be determined by one of the fol-lowing conditions.

If takeoff field elevation is higher than theprogrammed landing field elevation, the out-

put from the auto controller will position thetorque motor so that the cabin pressure jet isopen and the control pressure jet is closed.Control pressure will increase, inducing theoutflow valve toward closed, and cabin pres-sure will increase at a programmed rate untilit is equal to the programmed landing field el-evation.

and equambien

Referring to

sure is detertion and maapproximate

The control pthe main outby the commtroller. Whensure as a res

1 2 - 1 0

F OR T RAI NI N G P

URP O S E S

ONL Y

; ;

;

;

;

;

;

;

;

;

;

™

AUTO

MAN

NORM

DUMP

CABINPRESS

CABINALTITUDELIMITER

RELIEFVALVE

28 VDCBUS A1

FILTER

TORQUEMOTOR

QUADRANT

CABIN

LEGEND

CABIN PRESSURE

NEGATIVE PRESSURE

MODIFIED CONTROL PRESSURE

AMBIENT PRESSURE

ELECTRICAL POWER

ENGINE BLEED

CHECKVALVE



R ev

i s i on 3

;

NO. 1 AND NO. 2 ENGINEOR APU BLEED AIR

INTERIOR

EXTERIOR MAIN OUTFLOWVALVE

Figure 12-9. Pressurization Op

F OR T RAI


ON

L Y

R ev i s i on 3

AUTO

MAN

NORM

DUMP

CABINPRESS

28 VDCBUS A1

; ;

;

;

; ;

™


RELIEFVALVE

FILTER

TORQUEMOTOR

QUADRANTCHECKVALVE

LEGEND

CABIN PRESSURE

NEGATIVE PRESSURE

AMBIENT PRESSURE

MODIFIED CONTROL P

ELECTRICAL POWER

ENGINE BLEED



1 2 - 1 1

; ;


CABININTERIOR


VALVE

;

;

;

Figure 12-10. Pressurization O

Since the control chambers of both outflowvalves are interconnected, any change in con-trol pressure in the main valve control cham-ber wil l also be ref lected in the controlchamber of the emergency outflow valve, and,consequently, it remains slaved to the main out-flow valve.

If the pressurization test prior to takeoff revealsa malfunctioning PROG mode, the systemmust be tested in the FL and LDG modes. If these tests are valid, the crew must insert theflight level information into the controllercomputer. The ground prepressurization modewill be inoperative.

After takeoff the FL mode will achieve andmaintain the minimum lowest cabin altituderelative to the flight planned airplane altitudeof the isobaric range selected; for example, if the flight planned altitude is 25,300 feet, theFL program will maintain a sea level cabin ata differential of 9.3 psi. However, if this flightplanned altitude is exceeded, the system willautomatically transition to a new isobaric

range, and cabin altitude will increase at aprogrammed rate. When the airplane reaches50,000 feet (which is the isobaric limit of thesystem), the cabin altitude will be approxi-mately 8,000 feet at a differential of 9.3 psi.

manual mode contrlector switch to MAtating the manual countil the rate needle

Manual Mode

Figure 12-11 depicWhen the mode selall electrical powertroller. The manual ctering valve. One ppressure, the secondpump, and the third pmatic relay on the e

The control pressurrelay is determined ual control knob. Asclockwise, the contmatic relay is beinbellows is contracticabin pressure. The a metering port, exber of the emergenc

pressure of the jet p

As control pressurvalve is moved towsure, increasing the Th t f bi




If the airplane altitude exceeds 50,000 feet, thepressurization system will transition to dif-ferential control and maintain 9.3 psid. As

descent is initiated, the pilot must insert thelanding field elevation into the controller com-puter and also insert the reported QNH. No fur-ther selections are necessary through thedescent and landing.

When the airplane touches down and with thepower levers at idle, the automatic controllerwill initiate the controlled depressurization

The rates of cabinrectly proportionatthe control knob is mtrol knob should be

increments to avoidby rapid and/or larg

When operatinization contromonitor the sy

CA

F OR T RAI


ON

L Y

R ev i s i on 3

AUTO

MAN

NORM

DUMP

CABINPRESS

28 VDCBUS A1

LEGEND

CABIN PRESSURE

NEGATIVE PRESSURE

MODIFIED CONTROL P

AMBIENT PRESSURE

ELECTRICAL POWER

ENGINE BLEED

; ;

;

;

;

;

;

™


RELIEFVALVE

FILTER

TORQUEMOTOR

QUADRANT CHECKVALVE



1 2 - 1 3

; ;


CABININTERIOR


;

;

;

Figure 12-11. Pressurization Op

As explained earlier the main outflow valve isslaved to the emergency outflow valve duringmanual mode operation. A flow limiter in theslaving line will maintain the integrity of theemergency valve if the slaving line ruptures.The flow limiter also dampens any surges fromaffecting both valves.

PROTECTION

The protection systems consist of normal dif-ferential limiting, maximum differential lim-iting, negative differential prevention, andcabin altitude limiting. All except the normaldifferential limiting are incorporated in eachoutflow valve (Figure 12-10), and each is in-

dependent of the other.

The normal differential limit of 9.3 psi is afunction of the PROG mode in the controllercomputer. This limit is automatically main-tained during all normal flight conducted inthe PROG mode.

The maximum differential of 9.6 psi is a func-

tion of a relief valve in each outflow valve. Itconsists of a bellows-operated valve. The bel-lows is exposed internally to cabin pressure andexternally to static atmosphere. When the dif-ferential between cabin pressure and staticatmosphere reaches or exceeds 9 6 psi the

The low differentialture from the effecpressure. This condample, if the airplceeds the cabin ratchamber of each oucabin pressure whic

orifice. This pressumain diaphragm anance diaphragm. Thdiaphragm is exposure; consequentlproaches to within pressure, the balanceupward, opening thcabin and atmosph

will now descend aplane with accompfort. To ensure an imcondition, a one-waflow control chambmitting control pres

DEPRESSURI

The airplane may bwhen conditions waguarded switch on ttion control panel (Fposition. Power wiltomatic controller a




atmosphere reaches or exceeds 9.6 psi, thebellows expands, opening the valve and con-necting the control chamber to the atmosphere.The drop in control pressure induces the pop-

pet toward open, increasing the cabin airflowto the atmosphere, thus limiting the differen-tial to 9.6 psi.

Each outflow valve includes an altitude limiter,consisting of a valve connected to a sealed bel-lows evacuated to respond to a pressure equiv-alent to approximately 14,500 feet. The bellowsis exposed externally to cabin pressure. If air-

tomatic controller ato the torque motor The torque motor wto close the cabin pr

trol chamber of theopen the control prlow pressure. Contrrapidly, and the popSince the emergencto the main valve, iand rapid decompreimum cabin altitudecabin altitude to 14,

Since rapid decompress ion maycause severe discomfort, it shouldnot be selected until the cabin/am-bient differential is first reduced, un-

less a serious emergency dictatesotherwise.

BAGGAGECOMPARTMENTPRESSURIZATION

GENERAL

The baggage compartment located aft of thepassenger compartment is normally pressur-ized and maintained at approximately the samedifferential as the occupied areas. However,it can be isolated completely from the pres-surization system.

AIR SOURCES

The baggage compartment has two electri-cally controlled valves: one is a normally open

When the BAtion, the baglation valvesis controlled

Placing the Bwill close th

ply of heatiment. Pressthe open iso

If the switchthe heat valvthe baggagelated.

PROTECPositive and installed in heat valve fais moved to Ivalves will to-cabin pre

When the banegative preby the negatwill open apartment to


WARNING



cally controlled valves: one is a normally openisolation valve which connects the passengercabin to the baggage compartment for pres-surization, and the other, also normally open,

supplies air from the cockpit hot air line to amixing jet pump for baggage compartmentheating; therefore, there are two air suppliesto the baggage compartment.

CONTROL

Pressurization and heating for the baggagecompartment is controlled by a three-position

INDICAT

A light on tISOL” (AppBAG switch the isolationposition.

The operatiolation valvemaintenance


3



NOSE CONEVENTILATION/ PRESSURIZATION

is to ensure a positivcone for cooling oequipment.

AIR SOURCES

Figure 12-12. Baggage Compartment Pressurization

CONTROL

The ventilating air source is induced into thelower side of the nose cone by a fan powereddirectly from the DC system, through a relaycontrolled by a proximity switch. This airflows upward through the electronic equipment

and exhausts through a standpipe into the nosewheel well and then through louvers to theatmosphere.

The exhaust standpipe incorporates a butter-fly valve operated by a pneumatic actuatorand a spring. Cabin pressure is applied to theactuator piston and is opposed by the springand nose cone pressure. When cabin/atmo-

sphere differential pressure is low, such as onthe ground or during low-altitude flight, thespring maintains the exhaust butterfly fullyopen, and a probe on the butterfly shaft causesthe proximity switch to conduct, closing theblower relay and powering the blower.

The cabin airflow to the nose cone is con-trolled by a mechanically operated, normally

open isolation valve (installed on aircraft SNsprior to 70). The valve is controlled by a two-position lever (Figure 12-13) on the copilot’sconsole. The quadrant slot is labeled “NORM”and “CLOSED.” The supply line directs the airto the lower portion of the cone and incorpo-

PROTEC

A pressure-rcone bulkhein case the cThis valve w

bient differe

INDICAT

The temperament is sen


Figure 1



p prates a flow limiter to protect the pressurizedcabin in case of radome rupture.

After takeoff and when cabin/atmosphere dif-ferential is between 1.74 and 2.9 psi, the dif-ferential across the pneumatic actuator beginsto close the exhaust butterfly, shutting downthe fan, restricting the outflow, and increas-ing nose cone pressure. This process will con-tinue until the butterfly is fully closed at acabin/atmosphere differential of 3.04 to 4.35psi. Nose cone/ambient differential is now ap-

ment is senturns on the Non the warnilack of venti

OPERAT

Figure 12-1tion/pressuroperating moclosed, and calibrated hvalve on the


Figure 12-14. Nose Cone Ventilation/Pressurization

LEGEND

CABIN PRESSURE

AMBIENT AIR

EXHAUST AIR

FLAPBUTTERFLY

VALVEPRESSURE

RELIEF VALVE

ELECTRICBLOWER

PROXIMITYSWITCH

CABINPRESSURE

CABINPRESSURE

THERMALSWITCH

FLOW LIMITER

EVACUATIONDUCT

PNUEMATICACTUATOR

NOSEGEAR WELL

NOSE CONE

NOSE CONEISOLATION VALVE(SNs PRIOR TO 69)

M



LIMITATIONS

The limitations contained in the AFM per-taining to pneumatics, air conditioning, andpressurization are binding under law regard-less of the type of operation.

Figure 12 14. Nose Cone Ventilation/Pressurization


1. The emergency source of air for pres-surization is supplied from the:

A. Cabin air hot duct

B. Turbofan

C. Cockpit air hot ductD. No. 1 engine anti-icing duct

2. For normal operation of the pressuriza-tion system the manual controller mustbe positioned:

A. Midway between UP and DN

positionsB. At the DN position

C. To produce zero rate of change

D. To the up hard stop

3. When operating on emergency pressur-ization, cabin temperature is a function

of:A. Engine power

B. Alt itude

C. The cockpit air jet pump

D. All of the above

4. A solenflow va

A. Raptake

B. Dumener

C. Slowhigh

D. Lim9.6

5. The cablimit ca

A. A d

B. The

C. 14,5

D. A ramin

QUESTIONS



CHAPTER 13

HYDRAULIC POWER SY

CONTENTS

INTRODUCTION......................................................................

GENERAL .................................................................................

DESCRIPTION OF MAIN COMPONENTS............................

Pressurized Bootstrap Reservoirs .......................................

Engine-Driven Pumps.........................................................

Pressure-Holding Valve ......................................................

Accumulators......................................................................

Filters ..................................................................................

Standby Pump.....................................................................

Hydraulic System Control and Display Panel....................

NO. 1 HYDRAULIC SYSTEM.................................................




NO. 2 HYDRAULIC SYSTEM.................................................

AUXILIARY HYDRAULIC SYSTEM ....................................

In-Flight Operation.............................................................

On-Ground Operation.........................................................

LIMITATIONS...........................................................................

QUESTIONS .............................................................................

ILLUSTRATIONS

Figure Title

13-1 Hydraulic System Block Diagram..............................

13-2 Hydraulic Reservoir (Typical) ....................................

13-3 Reservoir Fluid Level Indication................................

13-4 Pressure-Holding Valve ..............................................

13-5 Reservoir Servicing ....................................................

13-6 Engine-Driven Pump Operation ................................

13-7 Pressure-Holding Valve Schematic ............................

13-8 Filter (Typical)............................................................

13-9 Standby Hydraulic Pump............................................

13-10 Hydraulic System Control and Display Panel ............

13-11 Standby Pump Selector ..............................................

13-12 No. 1 Hydraulic System..............................................

13-13 No. 1 System Components ........................................

13-14 No. 2 Hydraulic System ............................................




y y

13-15 No. 2 System Components ........................................

13-16 No. 1 and No. 2 Hydraulic Systems Operating ..........

13-17 Standby Pump Pressurizing No. 2 System ................

13-18 Standby Pump Pressurizing No. 1 System ................

TABLE

INTRODUCTION


HYDRAULIC POW

;

;

;

;

;



The Falcon 900 hydraulic systems consist of two indepesystems which cannot be interconnected and an auxiliary

draulic fluid spec ification MIL-H-5606 (NATO codes H

Main hydraulic power is supplied by three self-regulating pumps driven by the correspondingengine accessory gearbox. Each pump deliversa maximum flow of 7.6 U.S. gpm (28.65 1/min)and provides a regulated pressure output of

Selection of selector loca

The hydraulion the hydrau

GENERAL

Figure 13-1 presents a block diagram of hydrau-lic power distribution.

DESCRIPTION OF MAINCOMPONENTS

PRESSURIZED BOOTSTRAPRESERVOIRS

Each system has its own reservoir. The capac-ity of the No. 1 system reservoir is 2.38 U.S.gallons (9 liters); the capacity of the No. 2 sys-tem reservoir is 1.58 U.S. gallons (6 liters).Figure 13-2 shows a typical reservoir.

Each reservoir is pressurized by its corre-sponding hydraulic system. System pressureacting on a small piston that is mechanically con-nected to a larger (pressurization) piston imposespressure on the fluid in the reservoir. The dif-ference in piston diameters results in a low pres-sure of approximately 40 psi imposed on the

reservoir fluid. Reservoir overpressurization isprevented by a relief valve. Entrapped air isautomatically purged through a venting valve.

Each reservoir features a direct-reading fluidlevel indicator with two scales labeled ‘PRES

Reservoir SerServicing (topping uensures the correctsystem operation. Pthe applicable hydepressurized and complied with.

No. 1 System

1. Gear down andclosed

2. S la t s r e tr acte

3. Reservoir depred but ton ovalve (Figure pressurized th

4. Check the nitcumulators:

a. System ac(103.5 bar

b. Pressure1,700 psi (

No. 2 System

1. Slats and airbr

2. Reservoir deprfor No. 1 syste

3 Check the nitr




level indicator with two scales labeled PRES-SURIZED” and “UNPRESSURIZED” (Figure13-3).

Each scale includes two red ranges for low andexcessive fluid, and a green range optimumlevel at a temperature between –40 and 90° Cwhen the system has been unpressurized for sev-eral hours.

A quantity transmitter electrically actuates aHYDR QTY i di t i th k it

3. Check the nitrand pressure-has step 4 for N

4. Check the nitrsystem accumu

a. Thrust revpsi (103.5

b. Parking brpsi (70 bar

c. Airbrake plator—1 00


g t in

P1

P2

P3

NO. 1 SYSTEM NO

NO. 1RESERVOIR R

AUXILIARY HYDRAULIC POWER

MAIN HYDRAULIC POWER

STANDBY PUMP SELECTION(IN FLIGHT POSITION)

(OTHER POSITION—GROUND TEST)

STANDBYPUMP

ELEVATOR SERVOACTUATOR

AILERON SERVOACTUATOR

RUDDER SERVOACTUATOR

ELEVATOR ARTHUR

INBOARD SLATSOUTBOARD SLATS

LANDING GEAR AND DOORS

FLAPS

AIRBRAKES

OUTBOARD SLATS(EMERGENCY MODE)

NOSEWHEEL STEERING

EMERGENCY BRAKES

PARKING BRAKES

THRUST REVERSER

LEGEND

NORMAL BRAKES ANDANTI-SKID

AILERON ARTHUR QAIRCRAFT < 165



FOR TRAINING PURPOSES ONLY Revision 3

LEGEND

NO. 1 PRESSURE

NO. 2 PRESSURE

AUXILIARY PRESSURE

SUPPLY FLUID

Figure 13-1. Hydraulic System Block Diagram



VOLUME(LITERS) VOLUME(LITERS)

LEVEL

INDICATOR

NO. 1 RESERVOIR NO. 2 RESERVOIR

LEVELINDICATO

90°C (194°F)

90°C (194°F)

90°C (194°F)90°C (194°F)

20°C (68°F)

20°C (68°F)

20°C (68°F)20°C (68°F)–40°C (–40°F)

–40°C (-40°F)

–40°C (–40°F)

–40°C (–º40°F)

9

8

7

6

6

5

4

3

2

1

0

5

4

3

2

1

0

U N P R E S S U R I Z

E D


U

N P R E S S U R I Z E D





Figure 13-3. Reservoir Fluid Level Indication


0

1

2 3

45

32

NO. 1 HYDRAULIC FILLING VAL



TEST L R

NO. 2 HYDRAULIC SYSTEM

FILLING VALVE

FWD

30

Pump OperationRotation of the barrel (Figure 13-6) displacesthe sliding pistons, which draw fluid and pro-vide pressure output according to swash plateposition. At maximum output, the swashplate is at maximum tilt.

When discharge pressure is greater than refer-ence spring tension, the compensation spoolis pos itioned to regulate pressure applied to

the control pistonstroke of the pist oswash p la te ang l

As discharge pressusation spool is raiseallowing control pThe swash plate splate to the maximton stroke and pum


BARREL

CONTROL

PISTON

REFERENCE

SPRING

COMPENSATION

SPOOL



PRESSURE-HOLDING VALVE

The pressure-holding valve (Figure 13-7)consists of a relief valve, accumulator, check valve, and depressurizing button.

The check valve in the pressure-holdingvalve prevents accumulator depletion whenthe engine-driven pumps cease operation.The depressurizing button, when depressed,enables the hydraulic reservoir and systemto be depressurized for proper servicing.

The pressurtaining a hea

ACCUMU

Each systemmulator to da

instantaneouaccumulatorand a pressurtwo addition


TO

RESERVOIRD

B



RETURNACCUMULATOR


;

;

;

;

;

;

;

;

;

;

;

;

;

;

;

CLOGGING

DETECTOR

CLOGGING

DETECTOR

INLET

Figure 13-8. Filter (Typical)



power for the parking brake system and thethrust reversers.

NOTE

After engine shutdown or hydrau-l ic fai lure, hydraul ic pressure inthe system accumulator drops to

zero. Check valves maintain hy-draul ic pressure in the brake andthrust reverser accumulators.

FILTERS

Each main system incorporates filters equippedwith clogging detectors (Figure 13-8). In the

event a filter element becomes clogged beyondan acceptable level, the detector protrudes.

STANDB

In the eventa pressure drdriven standvide backupThe pump, Dcontrolled bOFF, ON, aswitch, locadisplay panesition, the ption in the presented un




PUMP 1

PUMP 3

PUMP 2

ST/BY

PUMP

TEST L R

ASKID

ON#1

HYDR

1/1 4

HYDR

1/1 4

HYDRAULIC SYSTEMCONTROL AND DISPLAYPANEL

Fluid quantity and pressure for each main sys-tem are read on the HYDR indicators (the leftindicator is for the No. 1 system). The PUMP 1,

PUMP 2, or PUMP 3 annunciator will come onwhen pressure from the respective pump dropsbelow 1,500 psi (103.5 bars). All coloredannunciators are shown in Appendix B. TheST/BY PUMP annunciator illuminates if thestandby pump operates continuously (longerthan 60 seconds) or if the standby pump selectorin the rear compartment (Figure 13-11) is not inthe IN FLIGHT position.

NO. 1 HYDRAULICSYSTEM

The No. 1 hydraulic system is pressurized by twoself-regulating pumps driven by the No. 1 andNo. 3 engines. Both pumps draw fluid from the

No. 1 reservoir (Figure 13-12).

The pressure output charges the No. 1 syplies pressure to bartor, and rudder servoslats, normal brakes,in the system is readSURE indicator on tThe PUMP 2 annuncgine not operating). cessive pressure engine-driven pump

If the pump on the Nthe remaining pumpsubsystem operation

When the engines apressure bleeds off, sure-holding valve surization.

External hydraulic pthe system for maint

Most of the No. 1 sysin the left hydraulic c





g t in

;

;

;

;

y

y

y

y

;

;

;

y

y

y ; y

;

;

;

y

y

y

;

;

y

y

;

;

y

y

;

;

y

y

; y

;

;

y

y

;

;

y

y

P1

; y

; y

LEGEND


RETURN

SUPPLY FLUID

NITROGEN

ELECTRICAL


QUICK DISCONNECT

NO. 1 RESERVOIR


PRESSURE-HOLDINGVALVE

SYPR

TR

ACCUMULATO




; y


CHECK VALVE

RELIEF VALVE




g t in

;

;

;

;

y

y

y

y

;

;

;

y

y

y ; y

;

;

y

y

;

;

y

y

; y

;

;

y

y

;

;

y

y

P1

; y

; y

LEGEND


RETURN

SUPPLY FLUID

NITROGEN

ELECTRICAL


QUICK DISCONNECT

PRESSURE

SWITCH

;

;

;

y

y

y

;

;

y

y




; y


CHECK VALVE

RELIEF VALVE




g t in

;

;

;

;

y

y

y

y

;

;

;

y

y

y ; y

;

;

y

y

;

;

y

y

; y

;

;

y

y

;

;

y

y

P1

; y

; y

LEGEND



RETURN

SUPPLY FLUID

NITROGEN

ELECTRICAL


QUICK DISCONNECT

PRESSURESWITCH


SYSTEPRES

TRANS

ACCUMULATOR

PRESSURE-HVALVE

NO. 1 RESERVOIR

;

;

;

y

y

y

;

;

y

y




; y


CHECK VALVE

RELIEF VALVE

Figure 1




g t in

;

;

;

;

y

y

y

y

;

;

;

y

y

y ; y

;

;

y

y

;

;

y

y

; y

;

;

y

y

;

;

y

y

P1

; y

; y

LEGEND

AUXILIARY PRESSURE

RETURN

SUPPLY FLUID

NITROGEN

ELECTRICAL


QUICK DISCONNECT

FILTER WITH SELF SEALING VALVE

;

;

;

y

y

y

;

;

y

y




; y


CHECK VALVERELIEF VALVE

Figure



;

;

;

;

y

y

y

y

;

;

;

y

y

y ;

;

y

y

;

;

y

y

;

;

y

y

; y

; y

; y

P1

;

;

y

y

; y

LEGEND

AUXILIARY PRESSURE

RETURN

SUPPLY FLUID

NITROGEN

ELECTRICAL


QUICK DISCONNECT

FILTER WITH SELF-SEALING VALVE

NO. 1 RESERVOIR

;

;

;

;

y

y

y

y ; y




; y

AND CLOGGING INDICATOR

CHECK VALVE

RELIEF VALVE

Figure 13-18.

1. The hydraulic reservoirs are pressurizedby:

A. Regulated bleed air

B. System pressure applied to a

small piston in the reservoirC. System pressure applied directly

to a large piston in the reservoir

D. An internal spring

2. When either of the main hydraulic systemsis depressurized, reservoir pressurization:

A. Dissipates

B. DecreasesC. Is maintained by the pressure-

holding valve

D. Is maintained and cannot bereleased

3. An indication of reservoir fluid level ispresented:

A. Only on the fluid level indicatoron the reservoir

B. On the indicator on the reservoirand on the indicator in thecockpit

C. Only on the indicator in thecockpit

5. The sta

A. Drathe flui

reseB. Drathe fluirese

C. Canmai

D. Canthe N

6. In flighin AUTin the Ipump a

A. Presdrop

B. Theof th

C. Presdrop

D. Presdropairb0 po


QUESTIONS



D. By a sight glass on the reservoir

4. The standby pump can be used to:

A. Pressurize the No. 1 system onthe ground.

B. Pressurize the No. 2 system onthe ground.

C. Pressurize the No. 2 system inflight

p

7. One of tvicing i

A. Maiin th

B. Ens

C. Relefromh

8. If the pump on the No. 1 or No. 3 enginefails:

A. The remaining pump will supplyfull pressure for actuation of all the as-sociated subsystems.

B. The No. 1 system will be com-pletely inoperative.

C. The standby pump automaticallysupplies pressure to the No. 1system.

D. The associated subsystems will besupplied with one-half the normalpressure.

9. The primary flight control servoactuatorsare pressurized by:

A. The No. 1 hydraulic system only

B. Both the No. 1 and No. 2 hy-draulic systems

C. The No. 2 hydraulic system only

D. The standby pump only

10. The engine accefrom damage inby:

A. A shear secdrive shaft

B. A pressure-system

C. A slip clutcgearbox

D. A bypass va




CHAPTER 14

LANDING GEAR AND B

CONTENTS

INTRODUCTION......................................................................GENERAL .................................................................................

Landing Gear ......................................................................

Brakes .................................................................................

Nosewheel Steering ............................................................

LANDING GEAR......................................................................

Main Gear ...........................................................................

Nose Gear ...........................................................................

Landing Gear Components .................................................

Controls, Indications, and Warnings...................................

Landing Gear Operation .....................................................

Wheels and Tires ................................................................

BRAKES....................................................................................

General




General................................................................................

Controls and Indications.....................................................

Operation ............................................................................

NOSEWHEEL STEERING.......................................................

General................................................................................

LIMITATIONS......................................................................................

Tires ................................................................................................

Brakes .............................................................................................

QUESTIONS.........................................................................................





ILLUSTRATIONS

Figure Title

14-1 Landing Gear and Doors.............................................

14-2 Main Landing Gear .....................................................

14-3 Main Gear Doors.........................................................

14-4 Hydraulic Emergency (Gear Pull) Gear Extension Ha

14-5 Nose Landing Gear and Doors....................................

14-6 Landing Gear and Door Operation..............................

14-7 Normal Landing Gear Control Handle, Position Indicand Circuit Breakers....................................................

14-8 Mechanical Gear Manual Unlocking Controls ...........

14-9 Landing Gear Position Indicator Panel.......................

14-10 Gear Position Indications............................................

14-11 HORN SIL Pushbutton ...............................................

14-12 Displays and Warnings during Landing Gear Retracti

14-13 Displays and Warnings during Landing Gear Extensio

14-14 Landing Gear Retraction.............................................

14-15 Landing Gear Extension .............................................



g

14-16 Gear Emergency Hydraulic Extension........................

14-17 Position and Warning Indications—Emergency Hydrand Gravity Extensions ...............................................

14-18 Main and Nosewheels .................................................

14-23 Normal Braking with Antiskid (Airplanes without SB F-900

14-24 Normal Braking with Antiskid (Airplanes with SB F-900-42

14-25 Tachometer Generator (Typical) ............................................

14-26 Antiskid System Test Indications...........................................

14-27 Emergency Brake Operation..................................................

14-28 Emergency/Parking Brakes....................................................

14-29 Parking Brake Accumulator Nitrogen Gage ..........................

14-30 Steering Actuator ....................................................................

14-31 Steering Handwheel...............................................................

14-32 Nosewheel Steering System...................................................

14-33 Scissors Disconnect Pin .........................................................

TABLE

Table Title

14-1 Functions Controlled by Flight/Ground Detection Proximity




INTRODUCTION

The Falcon 900 has a retractable tricycle landing gear congears and a dual-wheel nose gear. The main gears retracwell in the fuselage. The nose gear retracts forward into t


LANDING GEAR



enclosed by doors when the gear is retracted.Three hydraulic braking systems (normal, emergency, anbrake units in all main gear wheels. Antiskid is available

The nose gear is equipped with an electrohydraulic steerinshimmy feature.


The system includes main gear and door actua-tors, a nose gear actuator and telescopic lock-ing tube, gear and door uplock boxes, andproximity switches.

The normal landing gear handle is used for gearextension and retraction. Proximity switchessequence the opening of the main gear inner

doors, extension or retraction of the gear, andclosing of the doors.

The landing gear is mechanically and hydrauli-cally downlocked, pressure being maintainedon the gear actuators after extension. Thegears and main gear inner doors are uplockedby gear and door uplock boxes without hy-draulic assistance.

A hydromechanicalinner door opening asion. The main geagears are mechanhydraulic pressure actuators. Normal care deenergized dur

The landing gearsFigure 14-1.

BRAKES

Normal and emergeby either pilot. Norantiskid is poweresystem and is differretraction, automatmain wheel rotation

LH AND RH

FAIRING DOOR



GEAR DOORS

FAIRING

MAIN DOORS

Emergency (No. 2) braking without antiskid,powered by the No. 2 hydraulic system, isalso differential and progressive. As in normalbraking, either pilot can apply the brakes inthis mode.

Parking brakes are powered by the No. 2 hy-draulic system or by pressure from a parkingbrake accumulator. Braking is activated by aparking brake handle (not with the toe ped-als) and is progressive but not differential.Antiskid protection is not provided in thismode of braking.

NOSEWHEEL STEERING

Nosewheel steeri ng, powered by the No. 2 hy-draulic system, is provided by a rack-and-pin-ion actuator on the nose gear. A handwheel onthe left console provides electrical signals tooperate a control valve on the steering actua-tor. Feedback signals from a potentiometernulls steering signals as the selected positionis obtained.

The nosewheels are self-centering at lift-off, anda shutoff valve closes to prevent inadvertentn o sewh ee l s t ee r in g o p era t io n in f l ig h t .Antishimmy is provided within the steeringactuator.

LANDING GEAR

MAIN GEAR

Each main gear leg (Figure 14-2) consists pri-marily of a shock strut barrel housing a shock

absorber (lower unit) serviced with hydraulicfluid and nitrogen. The shock absorber mountsthe axle, wheels, and brake assemblies and isconnected to the barrel with scissor links.

Each main gear also features two flight/groundproximity switches and two tachometergenerators (wheel speed transducers) for thebrake antiskid system, one driven by eachmain wheel. When extended, the main gear isdownlocked by an integral lock in the gearactuator and by constantly applied hydraulicpressure. When retract ed, the gear is uplockedb y mech an ica l u p lo ck u n i t s th a t a rehydraulically unlocked in the normal operat-ing mode and mechanically unlocked in theemergency extension mode.

Main Gear Doors

Each main gear is enclosed at retraction with amain and fairing door (Figures 14-2 and 14-3).




Figure 14-3. Main Gear Doors

DOACTU

MAINDOOR



The fairing door is mechanically connected toand actuated by the gear. The electrically se-quenced main door is hydraulically actuatedto the closed position with the gear extendedor retracted in the normal mode and opensduring gear transit.

The main doors can be opened on the ground

by levers on the uplocks for maintenance orinspection, or, if the No. 1 hydraulic systemis pressurized, by pulling the hydraulic emer-gency gear extension handle labeled “GEARPULL” (Figure 14-4).

Pressure is applied to the doorlocks and thedoor actuators, releasing the locks and ex-tending the doors. To close the doors, push in

(stow) the GEAR PULL handle. This ener-gizes the normal sequencing circuits and ahydraulic selector valve, directing pressureto close the doors.

If hydraulic pressure is not available, eachdoor can be opened manually. Actuate the un-locking lever on the doorlock. Manually movethe door open, evacuating fluid from the door

cylinder. Close the door by moving it to theclosed position, mechanically engaging thedoorlock.

NOSE GEAR

The nose gear leg (Figure 14-5) includes anouter shock strut barrel housing a shock ab-sorber serviced with hydraulic fluid and ni

also functions as a dis hydraulically unand incorporates thindication proximit




sorber serviced with hydraulic fluid and ni-trogen. A rotatable inner barrel is controlledby the steering system.

The inner barrel and shock absorber are con-nected with scissor links. A hydraulically actu-ated rack-and-pinion steering actuator integralwith the inner barrel provides motion throughthe scissor links for nosewheel steering.


GEARUPLOCK

BOXFWD

ACTUATINGCYLINDER

FORWARDDOORS

TELESCOPIC

LOCKINGTUBE

UPLOCK



UPLOCK

ROLLER

Nose Gear DoorsFour nose gear doors (Figure 14-5) enclose thenose gear at retraction. The upper aft door is me-chanically connected to the strut barrel and tothe airplane structure. The lower aft door isconnected to the scissors. Both doors move toenclose the rear area of the gear at retraction.

The two forward doors are mechanically actu-ated by rollers on the torque link lower arm,closing the doors at retraction. Each doormechanism includes a proximity switch for thered nose gear light.

Figure 14-6 shows the landing gear extendedand retracted configurations and the nose gear

forward doors opened and closed.

LANDING GEARCOMPONENTS

Main Gear Bracing CylinderThe main gear cylinders are secured to thewing structure and to the main gear strut bar-

rel. Hydraulic extension or retraction of thecylinder piston raises or lowers the gear. Thecylinders also function as side braces with thegear extended and contain an internal mech-anism for locking the gear in the extendedposition. They are equipped with two prox-imity switches: one controls the gear operat-ing sequence, and the other provides gearposition indications and warnings.

locked down by an iand hydraulically ununit is equipped witthat control gear op

Main Gear Do

One door actuator i

attached to the airplaThe door is opened ation of the actuator, switch controlling ge

Flight/Ground D

The flight/ground dflight or on-ground

nose gear shock abious circuits accordsition is detected bon each gear.

The various functionthe detection systemproximity switcheTable 14-1.

Normal Gear OpSequence and In

Hydraulically sequing of the main geproximity switchesor downlock unit. Swcontrol the red MO




p g

Nose Gear Actuating Cylinder

The nose gear actuating cylinder is secured tothe airplane structure and to the nose gearstrut barrel. It hydraulically raises or lowersthe nose gear, but does not contain a downlock mechanism or proximity switches.

landing gear.

The doors wilthe main and nolocked or down




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UPLOCKROLLERS

ACTUATINGCYLINDER

NOSE GEARDOOR MECHANISM

PROXIMITYSWITCHES

SPRING

NOSE LANDING GEAR

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TELESCOPICLOCKING

TUBE

GEARUPLOCKBOX

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Figure 14-6. Landing Gear and Door Operation




Table 14-1. FUNCTIONS CONTROLLED BY FLIGHTDETECTION PROXIMITY SWITCHES

Authorizes on the ground:• One of the two switches

• Nosewheel steering

• Both switches• Auto radar forced standby• Omega• Flight recorder (option)• Public address (option)• ENG 2 FAIL warning• MFD• Normal antiskid

operation (SB 54)• Batteries cold air

cooling (SB-125)Authorizes during flight:

• One of the two switches• Copilot’s Davtron clock

One of the two proximity switches One

Combined with

Authorizes on the ground:• Normal antiskid operation• APU operation• Baggage compartment door (external) c• Battery ventilation• Ground idle for all three engines• Standby hydraulic pump cycling with sele• Takeoff warning indication• Test of:

• ADC 1 and 2• EFIS 1 and 2

• FGC 1 and 2• FMS 1 and 2• Utility monitor light• Pressure refueling• Lighting of main entrance door steps (SB

Cancels on the ground:• Automatic electrical load shedding

Controls on the ground:• Outflow valve opening

Inhibits on the ground:• Locking of the elevator Arthur unit auto h• Standby horizon operation

Authorizes in flight:• Air-conditioning valves (cabin and cockp• Pilot’s Davtron clock• Standby horizon (one of the two proximit

Prevents in flight:• Normal engine start• Thrust reverser deployment• Pressurization test

NOSE GEAR

LEFT MAIN GEAR1 (A1) 2 (A1)

LEFT MAIN GEAR1 (A1) 2 (A1)

1 (B1/A1) 2 (A1) 1 (B

1 (B1/A1) 2 (A1)LENOSE GEAR SHOCK



Prevents on the ground:• Displacement of the normal landing gear control handle

Ground and flight:• Provides slat relay and landing gear relay monitoring indication

Inhi• One of the twoproximity switches(1)

Combined with

One of the twoproximity switches(1)

Ste

Main Gear

Normal gear operation is controlled by the doorproximity switches. The gear will not extend orretract until the door actuators are fullyextended.

Nose Gear DoorsThe nose gear doors are not sequenced; they aremechanically actuated by nose gear movement.

Nose Gear

The nose gear telescopic locking tube and up-lock, in conjunction with the main gear prox-imity switches, control main gear dooroperation. The doors will not close until all threegears are either uplocked or downlocked.

CONTROLS, INDICATIONS,AND WARNINGS

Controls

Normal Landing GearControl Handle

The normal landing gear control handle (Fig-ure 14-7) controls a six-contact switch. Four of the contacts control the operation of the innerdoors and extension and retraction of the gear.

The fifth contact triggers the warning voiceand the red flashing light in the gear controlhandle. The sixth contact energizes the anti-skid system when the gear handle is down.

A red safety catch can be applied to lock the gearhandle in the down position. The catch is thumbedopen prior to placing the handle up or down. Inaircraft 157 and subsequent the safety catchis removed.

Inadvertent retraction on the ground is pre-vented by a spring-loaded locking pin thatholds the handle down. Extension of the lef tmain and nose shock absorbers at lift-off automatical ly releases the spring-loadedlocking pin, allowing handle movement. Thelocking pin can be overridden at any time bydepressing the red pushbutton above the con-trol handle. In aircraft 157 and subsequent thered pushbutton is removed.

The control handle incorporates a red light thatflashes when the gear is in transit or fails to lock in the retracted or extended position.

Hydraulic Emergency GearExtension Handle

This red control handle (Figure 14-4), locatedbeside the normal gear control handle, is labeled“GEAR PULL.” Pulling the handle out me-chanically actuates a hydraulic selector valve inthe hydraulic emergency extension system to di-rect No. 1 hydraulic system pressure for unlock-ing of the gears and doors and for gear extensiononly. A microswitch on the selector valve tripsto interrupt electrical power to the gear sequenc-ing system. Stowing the handle restores electri-cal sequencing and repositions the selector valveprior to normal gear operation.

Mechanical EmergencyExtension Controls

Three independent mechanical control handles(Figure 14-8) provide for manual unloc king of

the gears and the main gear inner doors.

The main gear handles are located in coveredrecesses in the floor on either side of thecenter pedestal. Each handle is mechanicallyconnected by a Teleforce cable to the corre-sponding gear door lock and gear uplock box.The nose gear handle is located in a recess onthe left side of the center pedestal, held byspring-loaded claws, or a spring on later aircraft.

Pulling the handles mechanically releases themain gear doors and the gear uplocks forfreefall emergency extension.

Indications

Landing Gear PositionIndicator Panel

This panel, labeled “LAND ING GEAR,” is lo-cated on the lower portion of the configurationpanel (Figure 14-9).

The panel includes six colored lights and aTEST pushbutton. The two red rectangularleft and right MOVING lights are controlledb y m a i n g e a r d o o r u p l o c k m e c h a n i s m




;

NCONT


IGA

5

L/G

CONTROL

IIML

5

ST BY

PUMP

LIDW

2.5

HYDR1

INDIC

ICC

5

ANTISKID

HYDR

IDG

5

L/G

INDIC

RIDW

2.5

HYDR2INDIC

IGH

5

NOSE

WHL

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SAFETYCATCH

OVERRIDEPUSHBUTTON

A/C <157

Figure 1




MAIN GEAUPLOCK UN

(LEFT ANDRIGHT)



proximity switches. When illuminated, theyindicate that the main gear doors are notlocked closed. During normal retraction orextension, the red lights indicate that thegear is in transit, extinguishing when gear se-quencing is properly completed and the doorsare up and locked.

The main gear red lights illuminate any timethe associated gear door is not closed. Thenose gear red light illuminates when the nosegear is not locked in either the extended or re-tracted position. It also illuminates if the nosegear is locked down and either of the two for-ward nose gear doors is not fully open.

The three green NOSE, LH, and RH lights are

controlled by proximity switches on the nosetelescopic locking tube and on the main gearbracing cylinders. Illumination indicates thatthe corresponding gear is locked down.

The green main gear lights are controlled by oneof two proximity switches on the main gearbracing cylinders and illuminate when the gearsare locked down. The green nose gear light is

controlled by one of the two proximity switcheson the nose gear telescopic tube and illuminateswhen the nose gear is locked down.

The indicator lights are powered through thelanding gear indication 1DG circuit breaker onthe primary B1 bus. If that bus is not powered,or if the circuit breaker is open, the lights areautomatically powered through the landing

t l 1GA i it b k th i

Warning VFlashing R

An audible vof the follow

• Airspehandleany ge

the powcan bepushb

• Flaps any geof the pspeed

Airspeeboth the

Figure 14-12plays and getraction and




gear control 1GA circuit breaker on the pri-mary A1 bus. Figure 14-10 shows light dis-plays during normal operations.

Warnings

Handle Light

A d li h i h l h dl



R

ev i s i on 3 . 0 1


ONL Y

NORMAL GEAR

EXTENSIONCONTROL

BEFORE GEAREXTENSION

*WHEN THE SLAT/FLAP CONTROL IS SET TO 40°,THE WARNING VOICE CANNOT BE SILENCED.

FLASLIG

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LANDING GEAR

NOSE MAIN

HORNSILENCE

WARNINGVOICE

*

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*



1 4 - 1 3

Figure 14-13. Displays and Warning

LANDING GEAR OPERATION

Normal Retraction (Figure 14-14)

At lift-off, extension of the left main andnose gear shock absorber s r e l eases thesolenoid-operated lockpin inhibiting gearretraction on the ground.

Moving the gear control handle up modifiesthe status of the six-contact switch, resultingin the following sequence of events:

1. The red light in the control handleflashes.

2. The extension solenoid in the gear elec-tric selector valve deenergizes, ventingthe pressure on the extension side of themain gear bracing cylinders to return.

3. The door-opening solenoid in the elec-tric selector valve energizes (providedthe gear is down and locked), directingpressure to unlock and open the maingear doors, as indicated by illumina-tion of the left and right red MOVING

lights.4. The gear retract solenoid in the elec-

tric selector valve energizes as soonas the main gear doors are fully open,directing pressure to unlock the gears,as evidenced by the green NOSE, LH,and RH lights going out.

NOTE

During gear retraction the brakes

(the left andout). The geaergizes, ventto return.

8. The door cloventing presto return.

9. All lights opanel are no

Normal ExtensWith the airplane itrol handle in the upthe handle and the wspeed decreases to

power levers has beePLA power settingabove 160 knots or levers above 73° Pings to cease and threset even if the voHORN SIL pushbu

Moving the gear co

fies the status of thesults in the followin

1. The brake anprovided theis on.

2. The red ligflashes, indidle and gea

voice activat




During gear retraction, the brakesare automatically applied to stopwheel rotation prior to retraction intothe wheel well.

5. The red center MOVING light comes onas soon as the green nose gear light goes

voice activatknots, any ppower settinnot down an

3. The door opelectric selerecting presthe main gea




Figure 14-14. Landing Gear Retraction

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VALVE


SELECTOR VALVE

HYDRAULIC UNIT

DOOR ESELECT

RETRACTION OPENINGEXTENSION

AUTOMATICWHEEL

BRAKING

GEAR

UPLOCKBOX

GEARUPLOCK

UNIT

DOORUPLOCK

BOX

MAIN GEARBRACING

CYLINDER

LEFT MAIN NOSE

RETRACTION

RETRACTION

NOSE GEARACTUATOR

RETRACTION

TELESCOPICLOCKING

TUBE

OPENING

LEGEND

NO. 1 SYSTEM PR

GEAR RETRACT

DOOR OPEN

RETURN

MECHANICAL

RESTRICTOR

SHUTTLE VALVE

DOORACTUATOR

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FAL


VALVE


SELECTOR VALVE

HYDRAULIC UNIT

DOOR ELECTRICSELECTOR VALVE

CLOSIRETRACTION OPENINGEXTENSION

GEAR

UPLOCKBOX

GEARUPLOCK

UNIT

DOORUPLOCK

BOX

MAIN GEARBRACING

CYLINDER

LEFT MAIN NOSE

EXTENSION

EXTENSION

NOSE GEARACTUATOR

EXTENSION

TELESCOPICLOCKING

TUBE

OPENING

LEGEND


DOOR OPEN, GEAR EXTEN

RETURN

MECHANICAL

RESTRICTOR

SHUTTLE VALVE

DOORACTUATOR

Figure 14-15. Landing Gear Extension



5. At full extension, the gears mechani-cally lock down. The green NOSE, LH,and RH lights come on. At the same time,the nose gear is locked down, both for-ward doors are fully open, and the cen-ter MOVING light goes out. The gearextension solenoid remains energized,maintaining pressure on the main gearbracing cylinders and nose gear actuator.

6. The flashing red light in the controlhandle extinguishes, indicating that allthree gears are locked down.

7. The door solenoid in the electric se-lector valve energizes, directing pres-sure to close the main gear doors. Whenthe doors are fully locked closed, theleft and right MOVING lights go out.

8. Pressure to the door actuators is re-lieved to return.

At touchdown, compression of the left mainand nose gear shock absorbers deenergizesthe locking pin solenoid, allowing the lockingpin to engage the control handle lock in thedown position.

Hydraulic EmergencyExtensionThe emergency extension system is to be usedwhen extension in the normal mode is unsuc-cessful. The control handle should be placedin the down position (the red handle light willbe flashing). The warning voice will be heard

if airspeed is below 160 knots and any one of the power levers is at a reduced power setting

actuaopen

4. PressbracigearPressand b

5. The and ron (th

6. The gcome

7. The flashsound

MechaniExtensio

The mechanisystem is usfail. No elecquired. The nbe in the dow

ing). The waspeed is belois at a redupower is ava

To activate t

1. Pull posit

rect f2 With




p ythe power levers is at a reduced power setting.

Pulling the GEAR PULL handle out results inthe following:

1. Electrical power to the landing gearcontrol circuit is interrupted.

2 Th l l i i i d

2. Withsuccegear uThis mboth mainlight

5. The main gear doors remain open, andthe left and right MOVING lights re-main on (the center MOVING light goesout when the nose gear is downlockedand both forward doors are fully open).

6. The green NOSE, LH, and RH lightsare on, indicating that the gear is downand locked.

Figure 14-17 presents gear position and warn-ing indications observed during emergencyhydraulic and gravity extensions of the gear.

WHEELS AND TIRES

Wheels

The airplane uses split-type wheels (Figure 14-18) on both the nose and the main gears. A sealbetween the wheel halves prevents nitrogen pres-sure leakage. Each wheel is equipped with an in-flation valve. Driving tenons in the main gearwheel drive the rotating discs of the brake unit.

Tires

The type of tire used on the airplane is Miche-lin AIR X radial. Main gear tires are 29 x 7.7x 15-inch; the nose gear utilizes 17.5 x 5.75x 8-inch tires. Installation of radial tires on onemain gear and conventional tires on the othermain gear is forbidden.

Tire inflation is shown in Figure 14-20.

C C

• Flat spots thforcing ply or

• Nylon rings the bottom of

BRAKES

GENERAL

The airplane has thrscribed below. Theing” means that the can be applied indepgear. The terms “pring” relate to increa

braking controls ar

Normal BrakinThe normal (No. 1) btial and progressiveBrake application iactuating the tips opedals). Pressure is

draulic system.

Emergency Br

Emergency (No. 2)No. 2 hydraulic syand progressive butprotection. This mowith the brake pedaas for normal braki




Checking Tire Condition

Check for tire slippage on the wheel. If the let-ter B and the red triangle are not in alignment,the tire has slipped and tire pressure must bechecked. Tires must also be checked for cuts,tears, and cracks.

Emergency/PaThis braking mode idraulic system or bhandle on the forwaates emergenc bra





VALVE


SELECTOR VALVE

HYDRAULIC UNIT

DOOR SELECT

GEAR

UPLOCKBOXGEAR

UPLOCKUNIT

DOORUPLOCK

BOX

MAIN GEARBRACING

CYLINDER

LEFT MAIN NOSE

EXTENSION

EXTENSION

NOSE GEARACTUATOR

EXTENSION

TELESCOPICLOCKING

TUBE

OPENING

LEGEND

NO. 1 SYSTEM PR

DOOR OPEN, GEA

RETURN

MECHANICAL

RESTRICTOR

SHUTTLE VALVE

DOORACTUATOR

Figure 14-16. Gear Emergency Hydraulic Extension






FAL

F i g u r e 1 4 - 1 7 .

P o s i t i o n a n d W a r n i n g I n d i c a t i o n s —

E m e r g e n c y H y d r a u l i c a n d G r a v i t y E x t e n s i o n s

T E S T

L H

N O S E

L A N D I N G G E A R

M O V I N G

R H

T E S T

L H

N O S E


M O V I N G

R H

T E S T

L H

N O S E


M O V I N G

R H

L A N D I N G

G E A R

N O S E

M A I N

H O R

N

S I L E N

C E

N O S E G E A R

E M E R G E N C Y

C O N T R O L

M A I N G E A R

E M E R G E N C Y

C O N T R O L

W A R N I N G

V O I C E

C O N F I G U R A T I O N

P A N E L

S L A T / F L A P

C O N T R O L

4 0 °

*

* W H E N T H E S L A T / F L A P C O N T R O L I S S E T T O 4 0 ° ,

T H E W A R N I N G V O I C E C A N N O T B E S I L E N C E D

W I T H A L L G E A R

N O T L O C K E D D O W N .

I A S <

1 6 0 K T

R E D U C E D

P O W E R

* *

GEA R UN L OCK PU SH PU LL

C L E A N

7 °

F L A P S

+

S L A T S

2 0

° F

L A P S

+ S L A T S

4 0

° F

L A P S

+

S L A T S

F L A S H I N G

F L A S H I N G

4 0 °

*

C L E A N

7 °

F L A P S

+

S L A T S

2 0

° F

L A P S

+ S L A T S

4 0

° F

L A P S

+

S L A T S 4

0 °

*

C L E A N

7 °

F L A P S

+

S L A T S

2 0

° F

L A P S

+ S L A T S

4 0

° F

L A P S

+

S L A T S

DRIVING TENONS

FUSESCREW

THERMALSHIELD

FUSE




INTENTIONALLY LEFT BLAN



TIRES

NOSE GEAR

MAIN GEARMTOW- 45,500 LB

MAIN GEARMTOW- 46,500 LB

9.8 BAR (142 PSI

12.8 BAR (185.5 PS

13.1 BAR (190 PS

NOSE GEARINFLATING PRESSURE

10° C(50° F)

BAR (PSI)

PRESSU

TIRE INFLATING PRESSURAMBIENT TEMPERATURE LESS TH

AIRPLANE ON JAC

20° C(68° F)

9.8 BAR(142 PSI)

10.2 BAR(148 PSI)

10.15 BAR(147 PSI)

10.55 BAR(153 PSI)

10.5 B(152

10.9 B(158

30° C(86° F)

40° (104°

MAIN GEARINFLATING PRESSUREBAR (PSI)

12.8 BAR

(185.5 PSI)

13.3 BAR(193 PSI)

13.24 BAR(192 PSI)

13.8 BAR(200 PSI)

13.7 B(198.7

14.25 (206.5




10° C(50° F)

EQUIPMENT:• PRESSURE GAGE• ADJUSTABLE NITROGEN SOURCE• LEAK DETECTION SOLUTION, MIL-L-25567

PROCEDURE:

20° C(68° F)

30° C(86° F)

40° (104°

NOTE:

Brake UnitsEach main gear wheel incorporates a compos-ite carbon brake unit (Figure 14-21) consist-ing of a hydraulic unit with two sets of fivepistons, each set pressurized by a differentsystem, and a set of carbon discs.

The brake ukeyed to andfour nonrotaforces the disto provide bfeatures returance adjustmvide a visual


BACP

BRAKE DISC

INDICATOR



CONTROLS AND INDICATIONS

Controls

Brake Pedals

The pilot’s and copilot’s brake pedals, spring-loaded to neutral, are mechanically intercon-nected by flexible control cables. Depressing

either set of pedals actuates the normal andemergency brake control valves. Since either leftpedal controls the brakes on the left main gearand the right pedals control braking on the rightmain gear, braking is differential.

PARK BRAKE Handle

This spring-loaded yellow handle is located in

the center section of the instrument panel (Fig-ure 14-22), beneath the glareshield. It has twolocking detented positions which can be un-locked by the red UNLOCK PUSH button inthe center of the handle.

The handle can be pulled to the first detent, butthe UNLOCK PUSH button must be depressedprior to movement to the second detent. To

return the handle to the pushed-in (stowed) po-sition from either detent, depress the UNLOCKPUSH button, and move the handle forward torelease the brakes.

BRAKE Selector Switch

This three-position switch, located on the hy-draulic control panel (Figure 14-22), selects

the braking mode. The switch is powered bybus A1 through the ST-BY PUMP circuitbreaker The upper face of the bevel shaped

IndicationsThe normal brakingdent on gear handletion, is powered bthrough the ANTI the overhead panel

When illuminated,

located on the hydraindicate that brakisponding gear has ibars). As the brakesout at a decreasingbars). The lights opbraking and will be isystem when it is sw

The amber #2 P BK ing lights panel, illflashing. It comes when the No. 2 hydthe emergency or paing pressure of 260the brake lines. It g145 psi in the lines

Flashing, the light the parking brake acbetween 1,102 and below) and that onetion remains.

OPERATION

Normal BrakinNormal braking is in




breaker. The upper face of the bevel-shapedswitch is red, and the lower face is green. Withthe switch in the #1 ON (upper) position, thegreen face is visible, and normal braking withantiskid is selected. The No. 1 hydraulic sys-tem pressurizes this braking mode. In the cen-

Normal braking is inpressing the brake pelector switch in the #selector valve is deesystem pressure is agency brake control


COPILOT’SPEDALS

PILOT’SPEDALS

PARKING BRAKEHANDLE

WARNING PANEL

PARK

UNLOCK

PUSH BRA

KE

BRIGHT

DIM

#2 P BK



L RPUMP 1

TEST

ASKID

1 / 4

1

3

4

HYDR

#1

#2

ON

OFF

PUMP 3

Metered braking pressure is directed throughantiskid servovalves to the brake units. Thebrakes on either main gear are applied inde-pendently by depressing the appropriate pedalto provide differential braking.

The pressure metered to the brakes is 2,102 to2,349 psi (145 to 162 bars) for airplanes SNs2 to 74 (except those with SB F900-42 incor-porated) or 1,595 to 1,711 psi (110 to 118bars) for airplanes SNs 75 and subsequent.Indication of pressure in the brakes is pro-vided by illumination of the green L and Rlights at 232 psi.

During retraction, gear upline pressure is appliedto both slide valves in the brake control valve toprovide a pressure of about 360 psi (25 bars) forbrake application during gear retraction into thewheel well. In the event of total electricalfailure, normal braking without antiskid isavailable from the No. 1 hydraulic system,regardless of BRAKE selector switch position.

AntiskidThe antiskid system prevents locking of themain gear wheels during heavy braking bymodulating the pressure at the brake units.The system is operative only with the BRAKEselector switch in the #1 ON position and thenormal gear control handle down.

The system includes an electronic control box,

an antiskid servo valve for each main gear, andsix tachometer generators. Four of the tachom-

• Memorized deceleration

With the shock absothe wheels not rotatitor switch in the #1box commands thedump braking pressdepressed.

When the shock abtouchdown, the maidump command is csure is restored. Thcauses the servo vapressure according treceived from the tamain gear wheels.

As the nosewheels electronic control servo valves to moaccording to airplanwith a programmeprovides braking prvery low speeds.




eters are driven by the main gear wheels andprovide wheel speed signals to the controlbox. Two tachometer generators driven by thenosewheels provide airplane rolling speed in-formation. A typical tachometer generator isshown in Figure 14 25


FlightS in

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;

;

;

;

;

;

;

#

;

;

;

;

BRAKE

ASKID

1 / 4

1 / 2

0

#1

#2

#1

ON

OFF

OFF

;

;

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;

;

H

NORMALCHAMBER

PILOT’S PEDALS

FLEXI

AUTOMATICBRAKING

NORMAL CONTROL

HANDLE

GEAREXTENDED

HYDRAULIC

PANEL

+

BRAKES

NOSEWHEELTACHOMETERS

#2 P BK

NO. 1SYSTEM

NORMAL SELECTOR VALVE(DEENERGIZED OPEN)

WHEELTACHOMETER

GENERATOR

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ANTISKIDELECTRONIC

CONTROLBOX

;

Figure 14-23. N




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g

1 4

- 2 8

P

F

PARK BRAKE HANDLE ACCUMULATOR

#2 P BK



;

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PARKBRAKE

DISTRIBUTOR

NO. 2SYSTEM

LEGEND

NORMAL PRESSURE

EMERGENCY PRESSURE

;

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; ;

Memory speeds for control box operation areprovided by one of the following:

• In flight, a 60 meter-per-second, fixed-speed reference until the nosewheels areon the ground (reference speed)

• Outboard or inboard main wheel speedinputs (whichever is lower)

When the nosewheels contact the runway, the60-meter-per-second reference speed is re-moved. Since the nosewheel rpm is greaterthan that of the main wheels, the control boxprocesses the new parameter to provide adeceleration rate that allows a maximumg-force of 0.6, controlled by the nosewheels.

In the event of system failure or hydroplaning

of both nosewheels, the control box referencespeed is provided by main wheel memoryspeed to allow a 0.3-g braking force.

To prevent low-speed vibration in the event of heavy braking below 40 knots, braking forceis limited to the 0.3-g speed deceleration rate.

In the event of a main wheel tire blowout, the

burst tire discriminator system will completelyrelease braking pressure on both wheels of the gear leg. After a 0.5-second time delaythe control box cancels the brake release order.From then on, the antiskid control box willfunction using rotation speeds from the un-damaged wheel.

Dual Braking SystemOn airplanes SNs 75 and subsequent andthose with SB F900-42, a dual braking sys-

The aircraft stwo nose lan

Electrical poSKID circuiply for syste

Correct operitored by thethe mainten

A discrepancleft and righter generatosystem failua red indicat

System T

Antiskid systthe ground BRAKE seleposition.

Ground Tes

With No. 1

able, perform

1. Set th

2. Deprthe gdraulpress(Figu

3. Deprdraulmaint




, g ytem is incorporated which assures greater ef-ficiency during pretakeoff full-power brakingor during runup with brake pedals operation(Figure 14-24).

The l

4. Releaseconcorre


L R

L R

IN-FLIGHT TESTLIGHTS

AFTER A DELAY

OF 0.8 SECONDLIGHTS ON FOR1 TO 2 SECONDS

THEN OFF

DEPRESS TESTBUTTON FOR0.5 SECOND

DEPRESS TBUTTON 1.5 SECO

ACTION

DEPRESS BRAKE PEDALSFOR DURATION OF TEST



3. Lights illuminate for approximately1 to 2 seconds and then go off.

4. Release the brake pedals; the lightsremain out.

Emergency Braking

With the No. 2 hydraulic system pressurized,emergency braking without antiskid is avail-

able when the brake selector switch is in the#2 OFF position.

Selection of the #2 OFF posit ion of theBRAKE selector switch completes circuitry toenergize the emergency brake control valve tothe open position and the normal brake con-trol valve to the closed position (Figure 14-27).This directs No. 2 system pressure or parking

brake accumulator pressure to the emergencychamber of the normal/emergency brake con-trol valve.

Depression of either set of brake pedals movesthe slide valves to meter emergency brakingpressure in direct proportion to applied pedalforce (progressive braking). Braking pressurepasses through shuttle valves and is applied to

a second set of pistons in the brake units, by-passing the antiskid servo valves. Metered pres-sure to the brakes is 1,100 to 1,203 psi (76 to85 bars) for aircraft SNs 2 to 74 (except thosewith SB F900-42 incorporated) or 1,000 to1,100 psi (69 to 76 bars) for aircraft SNs 75 andsubsequent. Indication of pressure being appliedto the brakes is provided by illumination of the#2 P BK light when pressure exceeds 261 psi.

As in normal braking, the brakes on eithermain gear are applied independently by

(Figure 14-2in direct propgressive bra

Braking presand is applieously (no di

The PARK Bpositions. Pumeters 595 ting the airpwithout lockdle to the se2,103 psi (1during enginwhen requir

Antiskimode. Eblowou

The #2 P BKoperation wit

The nitrogenaccumulatorport on the riaccumulatorrepeated app




main gear are applied independently bydepressing the appropriate pedal to providedifferential braking.

Emergency/Parking Brakes


FlightSafet international

; ;

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;

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;

;

;

;

;

;

;

;

;

TEST

ASKI

#1 O

L

;

;

;

;

BRAKE

ASKID

1 / 4

1 / 2

0

#1

#2

#1

ON

OFF

OFF

;

;

;

;

;

;

;

ANTISKIDELECTRONIC

CONTROLBOX

;

;

HYDRAUL

EMCH

NORMALCHAMBER

PILOT’S PEDALS

FLEXIBLE CAB

AUTOMATICBRAKING


GEAREXTENDED

HYDRAULIC

PANEL

+

+

#2 P BK

NO. 1SYSTEM

NORMAL SELECTOR VALVE(ENERGIZED CLOSED)

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Figure





FlightS in

; ;

; ;

;

;

;

;

;

;

;

#

;

;

;

;

BRAKE

ASKID

1 / 4

1 / 2

0

#1

#2

#1

ON

OFF

OFF

;

;

;

;

;

;

;

ANTISKIDELECTRONIC

CONTROLBOX

;

;

H

SV

PILOT’S PEDALS

FLEXI

NORMAL CONTROL

HANDLE

GEAREXTENDED

HYDRAULIC

PANEL

+

+

#2 P BK

NO. 1SYSTEM

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prior to checking the nitrogen charge. The ac-cumulator precharge should be approximately1,000 psi (70 bars).

NOSEWHEEL STEERING

GENERAL

The hydraulically powered, electrically con-trolled nosewheel steering system providesfor deflection of the nosewheels up to 60°either side of center for steering during take-off, taxiing, and landing. Electrohydraulicinpu t s to the sys tem must be manua l lyinitiated; however, antishimmy and nose gearcentering occur automatically.

STEERING SYSTEM

The steering system includes a handwheel,slaving unit containing a torque motor-actuatedcontrol valve, rack-and-pinion steering actua-tor, spring accumulator, and valves controllingpressure to the steering actuator. Also included

in the system are a solenoid-operated selectorvalve, a slide valve selector, a feedback poten-tiometer, and an electronic control box. Powerfor the electric control box is from bus B2through the NOSE WHL circuit breaker.

The steering actuator (Figure 14-30) is inte-gral with the barrel of the nose gear strut.Hydraulic pressure applied to either end of

the actuator piston moves the rack-and-pinionassembly to transmit motion through the scis-sors for steering.

A slide valve selectby gear retraction, ping in flight. The vgear extended and cbegins to retract.

The feedback potewheel position signtrol box to cancel ththe wheels reach th

The steering handwcated on the left crotated, the handwpotentiometer to pthrough the electronmotor on the steerin

The handwheel musing and energizing tRotation of the handsteering. When releato neutral, and the stenergized closed.




A solenoid-operated selector valve is electri-cally opened to direct pressure to the steeringassembly when the handwheel is depressedand rotated and is spring-loaded closed to pre-

ANTISHIMMY SYSTEM

With hydraulic power not supplied to the steer-ing system, nosewheel shimmy is dampenedby the antishimmy system consisting of thesteering actuator, two restrictor valves, and aspring accumulator. Damping occurs whenmovement of fluid from one chamber of the

steering actuator to the other chamber is forcedthrough the restrictors.

OS G

NOSEWHOPERAT

Nosewheel swheel (Figu120° in eithewheel rotatio8° from cenmovement, th

The handwhThis energizto direct pre(Figure 14-3

If the airplanselector hasopen to dire

Handwheel rsend electriccontrol box positioning otrol valve. Tmines flow tin the selecteactuator tranto deflect the


Figure 14-31. Steering Handwheel

LEFT CONSOLE



NOSE GEARCENTERING SYSTEM

The nose gear centering system consists of i hi h h k b b



;

DEPRESSED:

UNLOCKINGENERGIZATIONOF STEERING

SYSTEM

RIGHT 120° LEFT

CONT

WITH

HANDWHEEL

LOCKED

SELEC

PROXIMITSWITCHE

5A

BUS

B2

+

+

INFLIGHT

INFLIGHT

ROTATED:

WHEELDEFLECTION

CONTROL

Figure




When the nosewheels have reached the anglecommanded by the steering signal input, thefeedback potentiometer cancels the steeringsignal and steering ceases.

The scissors should be disconnected to allowfree swiveling of the nosewheels duringtowing (Figure 14-33). During the preflight in-spection, ensure the pin is reinstalled prior to

flight.

LIMITA

TIRES

The nosewhetires.

BRAKESThe brake kfoot-pounds




1. Normal landing gear and door opera-tion is:

A. Hydraulically controlled

B. Electrically controlled

C. Hydraulically actuated

D. Electrically controlled and hy-draulically actuated

2. The nose gear telescopic locking tube:

A. Is hydraulically locked with thegear extended and mechanicallyreleased for retraction

B. Mechanically locks the nose gear

down and is hydraulicallyreleased

C. B and D

D. Functions as a drag brace

3. Antiskid braking is available:

A. Only with the normal brakingsystem

B. With the normal and emergencybraking systems

C. With the normal, emergency, andparking brake systems

D. With any braking mode, providedthe ASKID switch is on

4. When extended, the main gears arelocked down by:

A. Manually installed downlock pins

B. Locking mechanisms integral withthe main gear bracing cylinders

5. When retractedlocked by:

A. The main g

B. Constantlypressure

C. MechanicaD. Overcenter

mechanism

6. The main gear

A. Hydraulicacally seque

B. Electricall

cally sequeC. Electrically

sequenced

D. Mechanicafrom the in

7. On the groundbe opened for

A. With the NB. Manually

C. Electricall

D. A or B

8. When illuminalights indicate

A. No pressur

the brake uB. Pressure is

brake units

C An overpre


QUESTIONS



the main gear bracing cylindersand hydraulic pressure

C. Downlock mechanisms that areseparate from the main gear bracingcylinder

C. An overprein the brak

D. Minimum b

9. The antiskid system:A. Receives inputs only from the

tachometer generators on the maingear wheels

B. Receives inputs from the tachome-ter generators on the main gearwheels and the nose gear wheels

C. Prevents main gear wheel lockup

during heavy brakingD. B and C

10. Pushing in on the nosewheel steeringhandwheel:

A. Electrically opens the steeringselector valve

B. Electrically opens the steering

selector valve and the slideselector valve

C. Unlocks the handwheel for rotationand electrically opens the steeringselector valve

D. None of the above





CHAPTER 15FLIGHT CONTROL

CONTENTS

INTRODUCTION......................................................................

GENERAL .................................................................................

PRIMARY FLIGHT CONTROLS.............................................

General................................................................................

Ailerons ..............................................................................

Rudder ................................................................................

Elevator...............................................................................

Horizontal Stabilizer...........................................................

SECONDARY FLIGHT CONTROLS.......................................

General ...............................................................................

Trailing-Edge Flaps ............................................................

Leading-Edge Slats.............................................................

Airbrakes ............................................................................

LIMITATIONS...........................................................................

Secondary Flight Controls..................................................



Airbrakes ............................................................................

QUESTIONS .............................................................................

ILLUSTRATIONS

Figure Title

15-1 Flight Control Surfaces ...............................................

15-2 Primary Flight Control Servoactuator (Typical) .........

15-3 Aileron Control System ..............................................

15-4 Aileron Emergency Actuator Buttons and AILERON

15-5 Arthur Q Unit..............................................................

15-6 Aileron Control System Block Diagram.....................

15-7 Aileron Trim Position Indicator..................................

15-8 Rudder Control System...............................................

15-9 Rudder Control System Block Diagram .....................

15-10 Elevator Control System.............................................

15-11 Elevator Arthur System...............................................

15-12 Horizontal Stabilizer Trim System .............................

15-13 Elevator and Horizontal Stabilizer Block Diagram ....

15-14 Horizontal Stabilizer Controls and Indicators ............

15-15 Mach Trim Pushbutton ...............................................

15-16 Flaps and Airbrakes ....................................................

15-17 Flap Hydraulic Motor and Reduction Gearing...........

15-18 Flap Control and Indicators ........................................




15-19 Flap System Operation ...............................................

15-20 Leading-Edge Slats.....................................................


15-24 Slats Controls and Indications, in Flight—Handle out of CLE

15-25 Automatic Extension of Outboard Slats—Handle in CLEAN

15-26 Automatic Retraction of Inboard Slats (Extended with Contr

15-27 Automatic Slat Control Block Diagram .................................

15-28 STALL 1 and STALL 2 Test Buttons .....................................

15-29 Emergency Extension of Outboard Slats ................................

15-30 AIR BRAKE Handle ..............................................................

15-31 Airbrakes System....................................................................

TABLE

Table Title

15-1 AOA Vane Angle versus Airplane AOA.................................



INTRODUCTION

The primary flight controls on the Falcon 900 airplane arebe manually actuated in the event both hydraulic systems

Aileron and rudder trim is hydraulic, init iated by electricing is provided but is not dependent on hydraulic pump p

ated, autopilot servo-controlled yaw damper prevents stabilizer trim is provided electrically.

The secondary flight controls consist of trailing-edge flapbrakes They are electrically controlled and hydraulicall


FLIGH



brakes. They are electrically controlled and hydraulicall

GENERAL

The primary flight control system is fullyboosted and nonreversible. Control inputsfrom the cockpit are transmitted through asystem of push-pull rods and bellcranks to thedual hydraulic servoactuators powering theailerons, rudder, and elevators. Failure of aninput control linkage results in the affected ser-voactuator returning to neutral.

In the event of a total hydraulic failure, the pri-mary flight controls may be operated man-ually but with some loss of efficiency. In thiscase, the servoactuators mechanically transmitcontrol inputs to the control surfaces, and air-speed should be limited for flight to 260 KIAS.

Aileron and rudder trim are controlled by dual-rocker switches located on the trim control

box on the center petuators reposition ttors, moving the eprovide aileron andtrim tabs on the airp

Horizontal stabiliztrolled by dual-rockwheels or by an eme

trim control box. Thby two 28-VDC mgency) to provide p

The electrically concan be actuated by thsystem, the airbrakby the No. 2 systemleading-edge slats ar


AILERON

LEADING-EDGESLATS

FLAPS



AIRBRAKES

Figure 15-1. Flight Control Surfaces

PRIMARY FLIGHTCONTROLS

GENERAL

Servoactuators

Hydraulic servoactuators in each primary con-trol axis position primary flight control sur-faces in response to control inputs. Figure15-1 shows both primary and secondary con-trol surfaces.

The servoactuators (Figure 15-2) consist of twoindependent barrel and piston assemblies oper-ating in unison. One barrel is powered by theNo. 1 hydraulic system, the other barrel by theNo. 2 system.

Each servoactuator incorporates the following:

• A piston dividing the barrel into twochambers. The piston rods are connectedto the airframe.

• A slide valve equipped with neutral re-

turn springs and end-of-travel stops• An input arm common to both barrels

and connected to control linkage fromthe cockpit. This input arm controls themovement of both slide valves.

• An auxiliary arm integral with the inputarm and connected to the auxiliary feelunit (AFU) to center the controls.

• A bypass valve that interconnects thetwo chambers in the event of hydraulicfailure. A calibrated heating orificemaintains the hydraulic fluid to a satis-factory level.

• A check valve in the pressure line toprevent backflow if high aerodynamicload overcomes hydraulic force.

Movement of the control input linkage movesthe slide valve, directing pressure to displacethe servoactuator barrel, which is connectedto the control surface. In addition to moving

the control surface, movement of the barrelcounters the relative movement of the slidevalve to provide follow-up action.

If one hydraulic system fail s, the bypass valvein the corresponding barrel interconnects thetwo chambers. Therefore, the depressurizedbarrel offers minimal resistance to the re-

maining operative barrel.

If both hydraulic systems fail, return springsmove the slide valves to neutral. Control inputsthen mechanicall y move the barrels to providemanual deflection of the control surface.

Flutter and Gust Damping

Automatic gust damping of the rudder andelevator movement on the ground and pre-vention of surface flutter in f light in the eventof hydraulic failure are provided by a cali-brated valve (Figure 15-2). After engine shut-down, the elevators move slowly to the fullydown position.

Artificial Feel System (Springs)Hydraulically boosted flight controls do notprovide aerodynamic load “feel” to the pilot.Therefore, spring-loaded artificial feel units(AFUs) are incorporated in the primary flightcontrol linkages prior to the servoactuators.The load feel provided is in direct proportionto control input movement and the resultantspring compression. Compression or extensionof the AFU produces a force that must be over-come by the crew or the autopilot.

The force pattern of the AFU varies with theamount of control input. Force required tocompress the AFU increases directly to theamount of control input. There are two levelsof resistance in the aileron and r udder AFUsand three levels in the elevator AFU. The rud-

der AFU has a breakout resistance that mustbe overcome prior to leaving neutral.




FIXED POINT

CHECKVALVE

HEATINGORIFICE

INPUT ARM

TO AUXILIARY A

TO AIRFRAME

NOTE: SINCE BOTH BARRE ONLY ONE IS SHOW

Fig



R ev i s i on4

1 5 - 4

F OR T RAI NI N G P URP

O S E S

ONL Y

Figure 15-3. Ailero

;

;

PILOT’S CONTROLWHEEL

COPILOT’S CONTROLWHEEL

PRESSURESEAL BOX

;

;

;

;

;

VARIABLEBELLCRANK(ARTHUR Q UNIT)

AFU

AILERONTRIM ACTUATOR




Auxiliary AFUA spring-loaded auxiliary AFU is connectedto the airframe and to the auxiliary arm of each servoactuator (Figure 15-2). If controlinput linkage fails, these AFUs provide a cen-tering force to the servoactuator slide valves.

AILERONS

General

Aileron deflection is initiated by moving thepilot’s or copilot’s control wheel. Control in-puts are through a rigid linkage composed of fixed and adjustable push-pull rods hinged onbellcranks. The linkage is routed along theright side of the fuselage, where it is separated

into left and right paths (Figure 15-3).

From this point to the servoactuators, linkagesare routed forward of the wing front spars. Inthe left wing, an electrically actuated emer-gency aileron actuator can also drive theservoactuator. In the emergency mode, it willcontrol the left aileron if control linkage failsor seizes.

NOTE

If the left servoactuator is unable tomove, the emergency unit will op-erate the right servoactuator in theopposite direction of control switchmovement through the control push-pull rods.

Aileron Operation

Movement of the input arm positions slide

AFU. If the becomes disslide valvesmaining ailetrol for fligh

In the eventemergency abe utilized to

to position ttrolled flightred emergenAILERON trtons are sepac lude depsimultaneou

The position

is indicated ZERO warn



valves in the servoactuator (as during normaloperation), causing the barrels to move andreposition the aileron. As the desired posi-

actuated by a switch in the actuator. The lightis on when the actuator leaves the neutralposition.

Arthur Variable BellcranksVariable bellcranks are insta lled in the aileronand elevator control linkages to vary the feelforce of the flight controls. The feel force varies

with changes in the IAS (ailerons) and withthe position of the horizontal stabilizer (pitch).

As airspeed increases, the pivot point of theaileron bellcrank, commonly referred to as an“Arthur Q unit,” changes to effectively in-crease the spring tension of the AFU (Figure15-5). Horizontal stabilizer nosedown position(high-speed Arthur) also results in an increase

of the tension of the elevator AFU.

This, in turn, requires that more force be appliedto the controls for a given control input. Theaileron Arthur Q unit is pressurized by the No. 2hydraulic system for aircraft lower than SN 165or electrically controlled on SN 165 andsubsequent. The elevator Arthur unit ispressurized by the No. 1 hydraulic system.

Arthur Unit Principlesof Operation

The Arthur unit consists of an actuating cylin-der articulating around a pivot point. The rodsliding within the cylinder is connected to themain AFU at one end and at the other end tothe trim actuator or to an elevator control rod

connected to the main linkage (Figure 15-5 and15-11). The position of the sliding rod, andtherefore the pivot point, depends on the hy-draulic pressure in the cylinder.

2 inputs or ADC 1 oThis variation is aelectric motor.

In the event either hspring-loaded slidinspeed, “controls eastoring is provided by

Aileron Arthur QFor aircraft lower thunit is monitored binput from the ADCof ADC 1 failure) wpotentiometer on thpositioned by an IASstandby pitot and st

For aircraft SN 165unit is monitored bt iomete r on the Afrom ADC 1, or onfrom ADC 2.

If a difference of apists at low speed, or

amber AIL FEEL wFor aircraft SN 165is provided after a also illuminates if thoperative (all colorein Appendix B).

If the No. 2 hydraullower than SN 165,

curs on all subsequautomatically reverttrol forces during lanever, at high speeds




For aircraft lower than SN 165, the aileron Qunit cylinder pressure is controlled by an ori-

Aileron TrimAil i i l

R ev i s i on 3

1 5 - 7


ONL Y

Figure 15-5. A

HIGH SPEED MODE

MAIN AFU

ARTHURQ UNIT

PIVOT

POINT

AILERONTRIMACTUATOR

INPUT COCKPWHEEL

NOTE: POWERED BY NO. 2 HYDRAULIC

STATIC

STANDBY

PITOT




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STANDBY PITOT-STATIC SYSTEM

TRIMPOSITIONINDICATOR

TRIMPOSITION

INDICATOR

ADC 2 FOREMERGENCY MODE

ARTHURQ UNIT

MAINAFU

TRIM CONTROLBOX

CONNECTINGROD

WARNING PANEL

TRIMCONTROL

BOX

NO. 2HYDRAULICSYSTEM

ADC 1

SERVOMOTOR

AUTO-PILOT

EMERGENACTUATO

±7°

AIRCRAFT PRIOR TO SN 165

TRIM±12°30'

;

;

;

;

;

;

TRIM POSITION

ARTHURQ UNIT

MAINAFU

ADC 1 A/C ≥179

ADC 2 A/C 165 TO 178

TRIM±12° 30'

SERVOMOTOR

AUTO-PILOT

ADC 1 A/C 165 TO 178ADC 2 A/C ≥179

EMERACTU

±



;

TRIM POSITIONINDICATOR

(2DQ)TRIM

CONTROL

CONNECTINGROD

WARNING PANEL

NOTERunaway aileron trim can be stoppedby applying opposite trim and thendisengaging the cont rol c i rcui tbreaker. Trim will remain inoperative.

Movement of the AILERON trim switch ap-plies 28 VDC to the trim actuator. Linear mo-

tion from the actuator displaces the input armon the servoactuator, thus hydraulically repo-sitioning the ailerons. Maximum trim deflec-tion of the aileron is 12° 30’. Maximum ailerontrim is indicated by a reading of 50 on thetrim indicator.

Aileron position is shown on the AIL positionindicator on the instrument panel. Do not at-

tempt aileron trim when hydraulic pressure isnot available.

A block diagram of the aileron control system isshown in Figure 15-6. The AIL position in-dicator is shown in Figure 15-7, the AILERONtrim switch in Figure 15-4.

RUDDER

General

Positioning of the rudder is initiated by mov-ing the pilot’s or copilot’s rudder pedals. Controlis through a rigid linkage composed of fixed andadjustable push-pull rods hinged on bellcranks.The linkage is routed along the right s ide of the

fuselage, leaving the pressurized sectionthrough a pressure sealing box (Figure 15-8).From that point, it is routed through the verti-cal stabilizer to the servoactuator.

the servoacting (barrel) As the desirvalves return

The main Alinkage throtrim actuatoin parallel to


Figure



Rudder Operation

O id f h i ll d di

artificial feeopposite serv

dd

R ev i s i on 3

1 5 - 1 0

F OR T RAI NI N G P URP

O S E S

ONL Y

Figure 15-8. Rudde

COPILOT’S PEDAL

PILOT’S PEDAL

SERVOYAW

DAMPER

AUXI




0L R

0

50 50

;

;

S

ACT

;

-2

02

4

6

8

10

AFT

FWD;

;

NORMAL

UP

TAILPLANE AIRBRAKES

0

1

DOWN

AILERON E

MERG

TRIMC

CONNECTINGROD

AUXA

MAINAFU

AUTO-PILOT

COPILOT’SRUDDERPEDALS

PILOT’SRUDDER

PEDALS

YAWDAMPER

TRIMACTUATOR

;

;

;

;

040 40

L R

AILRUDDER

;

;

;

;

;

;

RUDDER



TRIMPOSITION

INDICATORTRIM CONTROL BOX






;

;

;

;

;

y

y

y

y

y

; ;

;

; ;

y y

y

y y

;

;

y

y

PITCHFEEL

ARTHUR UNIT

BUSA1

NO. 1HYDRAULIC

SYSTEM

;@;@

F i g u r e 1 5 - 1 0 .

E

l e v a t o r C o n t r o l S y s t e m

2 0 °

1 6 °

P I L O T ’ S C O N T R O L

C O L U M N

C O P I L O T ’ S

C O N T R O L

C O L U M N

I N P U T

A R M

S E R V O A C T U A T O R

A U X I L I A R Y

A F U

A R T H U R U N I T

S E R V O V

A L V E

P R E S S U R E

S E A L B O X

A U T O P I L O T

Q U A D R A N T

M A I N

A F U

A R T H U R

U N I T

E L E V A T O R A R T H U R

C O N T R O L B O X



R ev

i s i on 3

1 5 - 1 5

F OR T RAI NI N G

P URP O S E S

ONL Y

Figure 15-13. Elevator and Horiz

TRIMPOSITION

INDICATORWARNING PANEL

WARNING PANEL

TRIM

CONTROLBOX

XTK 2

OPEN

XTK 2

CLOSEDAIL

ZEROAIL

FEEL

AP MISTRIM

NOSECONE OVHT

BLEEDAPU

PITCHFEEL

BAG

ACCESS

MACHTRIM

BAGISOL

REARDOORS

T/OCONFIG.

AUTOPILOT

NORMALDUAL-ROCKER

SWITCH

NORMALDUAL-ROCKER

SWITCH

NHYDRAU

CONNECTINGROD

– 4IAS GREATER

IAS LESS TH–10

AIRPLAGROU

+ONE POWE

BEYON+

HORIZONTALOUT OF – 4° TO

RELAYBOX

SERV

AUTOPILOTAUTOTRIM

MACHTRIM

AUTOPILOT

SERVOMOTOR

AR

ACT




DOWN

NORMAL

UP

RUDDER

TAILPLANE AIRBRAKES

0

1AILERON EMERG

XTK 2OPEN

XTK 2CLOSED

AILZERO

AILFEEL

AP MISTRIM

NOSECONE OVHT

BLEEDAPU

PITCHFEEL

MACHTRIM

BAGISOL

BAGACCESS

T/OCONFIG.

;

;

;

;

;

;

PILOT’S

DUAL-ROCKERSWITCH

EMERGENCYPITCHTRIM SWITC

NORMAL CONCIRCUIT BREA

HORIZONTALSTABILIZERACTUATOR

TRIMPOSITION INDICATOR

WARNING PANEL

TRIM CONTROL BOX

L R

40 400

AIL

L R0 0

50 50

RUDDER

NOSEDN

NOSEUP

AFT

FWD

-2

0

2

4

68

10



Horizontal Stabilizer Operationwith the Mach Trim System

To augment stability in the pitch axis at highMach numbers, the Mach trim system modu-lates the stabilizer position as the Mach num-ber varies. This system is normally activatedprior to takeoff and becomes operative whenthe Mach number is above .775. The pitch

trim command inputs are supplied by the Machtrim control box in parallel with the normalmanual control inputs.

With the Mach trim system engaged, the nor-mal trim system can be used at any time to adjust stabil izer position. When the normal pitchtrim switch is released, the Mach trim systemresumes its automatic operation.

NOTE

Mach trim is overridden by the auto-pilot or stabilizer trim normal control.

The Mach trim system is activated by the MTRIM pushbutton on the autopilot controller(Figure 15-15). The amber MACH TRIM warn-ing light is illuminated when the system is de-energized or when a system fault is detected.

SECONDARY FLIGHTCONTROLS

GENERALThe secondary flight controls include electricallycontrolled and hydraulically actuated trailing-edge flaps, leading-edge slats, and airbrakes.The flaps are actuated by an electrically con-

TRAILINGGeneral

The trailing-of an inboasegment on mounted for


Figure 15



trolled hydraulic motor driving torque tubes andj k Th l di d l di

The flaps are controlled by a combinationslat/flap handle on the slat/flap control box.Handle movement provides circuitry to a con-trol valve on the flap hydraulic motor reduc-tion gear (Figure 15-17). Reduction gearingdriven by the motor provides motive force toactuate the flaps through torque tubes and jackscrews.

Overtravel of the flaps is prevented by a spring-loaded brake within the motor.

Hydraulic power isdraulic system or bpump if the No. 2 hposition indicationtor on the instrumetection is also prov

Flap Operatio

The flaps are controslat/flap control bo


LEFT-WINGTRANSMISSION

REDUCER

RETURN

HYDRAULIC MOTOR

HYDRAULIC M

MICROSWITCH BO

SOLENOIDSELECTOR VALVE

PRESSURENO. 2

HYDRAULIC

SYSTEM

FLAP RETRACTIONSOLENOID



FLAP EXTENSION

The handle is spring-loaded and detented and

must be pulled upward prior to movement toany position. Since the handle also controls theleading-edge slats, selection of any flap po-sition causes the slats to extend.

The flaps can be extended by moving theslat/flap handle to the 7° FLAPS+SLATS po-sition. This energizes the flap control valve,and the hydraulic motor operates to drive theflaps (Figure 15-19).

As the flaps reach the selected 7° position, thecontrol valve is deenergized, and pressure isremoved from the hydraulic motor. The brak-ing device in the motor is hydraulically oper-ated during operat ion and then released,holding the flaps in position. Further exten-sion can be made by positioning the handle to

the next position.

Flap Indication System

A flaps asymmetry detection circuit comparesflaps position information from position trans-

mitters on each outboard flap. If asymmetry

greater than 5° (flaps near takeoff position) or8° (flaps near landing configuration) exists, adifferential comparator triggers the amberFLAP ASYM warning light. The flap motorcontrol circuit breaker is also automaticallydeenergized to remove hydraulic pressurefrom the motor, and the motor brake is ap-plied to hold the flaps in position.

With the airplane on the ground, the flaps ex-tended beyond 22°, and at least one powerlever advanced beyond the 84° power leverangle, the “no takeoff” voice sounds, and thered T/O CONFIG warning light comes on.

There is no provision made for emergency op-eration of the flaps. However, the standby hy-draulic pump can be used to actuate the flaps

in the event the No. 2 hydraulic system pumpfails for reasons other than line rupture and lossof fluid. Regardless of the pressure source,the flaps control handle is used for flap oper-ation, and asymmetry protection and positionindication are the same as previously described.




POINTER

Figure 15-18. Flap Control and Indicators

SLAT/FLAP CONTROL

CONFIGURATION PA

AIRBRAKE

MOVING

LANDING GEAR

NOSELH

TEST

RH

SLATS

FLAPS

40DN

XTOP

AZE

A

NOCONE

HOBA

BAACC

AUDIWARNI

“NO TAKE

CLEAN

7° FLAPS+SLATS

20° FLAPS+SLATS

40° FLAPS+SLATS

E

S






LEADING-EDGE SL

General

The leading-edge slats exteleading edge of the wing (Fwing incorporates one inleading edge) and one outb

Slat operation is electricallydraulically actuated. Threeach outboard slat. One actuextend force; two are douential) units. Each inboardone differential actuator. continuously pressurized toin the retracted position.

NOTE

The single extend acoutboard section is foperation.

Operation of the inboard anseparate; each has its own valve. Slat operation is elec

causing the outboard slats ting extension. Logic circuboard slat retraction withflaps extended, and retractislats with the inboard slats operate in three modes: nand emergency.

If hydraulic pressure failsslats are held retracted bymaintain residual pressure

Control and Indicat

The slats are normally cohandle on the slat/flap co15-20). Moving the handle

lects full extension of the edge flaps, as selected). Emonly of the outboard slatwith the guarded EMERswitch adjacent to the norm

Slat position indication is pl ig h t resemb l in g a s l a t

FA

SLAT/FLAP CONTROL BOX

SLATPOSITION

INDICATOR

SLAT/FLAP

CONFIGURATION PANEL

EMERGENCYSLATS

(OUTBOARD)

CENTERCIRCUIT-

BREAKER PANEL

WARNING PANEL

XTK 2OPEN

XTK 2CLOSED

AILZERO

AILFEEL

AP MISTRIM

NOSECONE OVHT BLEEDAPU

PITCHFEEL

MACHTRIM

BAGISOL

REARDOORS

T/OCONFIG.

R AOAL AOA

L PITOT ST BYPITOT

L WHLOVHT

HOTBAT

R PITOT

R WHLOVHT

FLAPASYM

BAGACCESS

;

;

;

;

;

;

;

;

;

;

;

AUTOSLATS

OUTBOARDSLAT

EMERGENCYACTUATOR

SLAT

ACTUATOR

INBOARDSLAT

ANGLE-OF-ATTACKVANES

;

;

;

;

;

;

;

;

;

;

FLT CONTROL

RH AUTOSLAT

21CF10

SLATINDIC

LHAUTOSLAT

AOAINDIC

(FOR A/C ≥ 162)

CLEAN

7° FLAPS+SLATS

20° FLAPS+SLATS

40° FLAPS+SLATS

EMERG

SLATS

AIRBRAKE

MOVING

LANDING GEAR

NOSELH

TEST

RH

SLATS;

;

;

;

;

;

; ;

0

FLAPS

7

2040DN

UP

;

5

5

Figure 15-20. Leading-Edge Slats



R ev i s i on4

1 5 - 2 1

F OR T RAI NI N G

P URP O S E S

ONL Y

Figure 15-21. Norm

SLATS/FLAPCONTROL BOX

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

LEGEND

NO. 1 PRESSURE

ADC 1IAS LESS THAN

265 KT

AOA

LESSTHAN 23°LEFT ANGLE-

OF-ATTACKVANE

INBOARDVALVEBOX

AUTOMATIDISTRIBUTO

INBOARDSLAT

SOLENOIDSELECTOR

VALVE

OUTBOARSLAT

SOLENOIDSELECTOR

VALVE

EXTEND RESTRIC

LEFT OUTBOARDLEFT

INBOARD

R E R E R E



The outboard selector valve is dual-solenoidactuated. One solenoid is powered by bus A1through L/H AUTO SLAT breaker. The othersolenoid receives power from the battery busthrough the SLATS circuit breaker located onthe main electrical box (6PA). Either solenoidwill actuate the selector valve. The inboardsolenoid is powered only through the L/HAUTO SLAT circuit breaker.

When the outboard slats have been extendednormally via the slat/flap handle and there isa failure of the main left and right buses, theoutboard slats are kept extended by electricalsupply from the battery bus.

Slat control and indications in flight with thecontrol handle in CLEAN and out of CLEAN

position are shown in Figures 15-23 and 15-24.

Automatic Operation(Clean Stall)

Automatic operation of the slats is armed byIAS (from the ADC), is controlled by angleof attack (AOA), and is operative up to 265knots, Airplane AOA is detected by two angle-

of-attack vanes, one on either side of the for-ward fuselage (Figure 15-22). The left systemis armed by ADC 1; the right system by ADC2. Each vane is heated through the PITOTswitch on the overhead panel.

The left vane is poSLAT circuit breakeby the R/H AUTO S

On the ground, autohibited by the mainimity switches. Thethe left main gear; tmain gear.

In flight, with the sand IAS below 265 kmal actuators supplitem pressure automaslats when the left AOA. Similarly, if thairplane AOA, No. tem pressure suppl

gency outboard slaextending the outboatracted when airplathose values (Figure

If the slats have beecontrol handle out othe inboard slats auairplane AOA exce

The outboard slatsthe airplane AOA dethe inboard selectoand the inboard sla

A block diagram ofsystem is shown in F


Table 15-1. AOAVER

AOA VANE ANGL

14°

16°

23°

34°



system is shown in F


CONTROL

1

AIRPLANEANGLE OFATTACK

INITIALCONFIGURATION

AOA11°

(INCREASING)

AIRPLANE AOA

SLATPOSITI

CLEAN

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

AIRPLANE IN CLEANCONFIGURATION,SLAT/FLAP HANDLEIN CLEAN

• EXTENSION OF OUTBOARD SLATS

• RED SLAT-MOVING LIGHT ON

• AUTOMATIC IGNITION, IGN LIGHT ON

CLEAN

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

A

CROSS-OUTBO


;

;

A

CROSS-OUTBO


;

;

INCREASING

AOA = 11°

IGN

DECREASING

AOA = 11°

IGN

1

2



• AUDIO WARNING


CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

CONTROL 7° FLAPS +SLATS

AIRPLANEAOA

SLATSPOSITION

A

B

CROSS-SECTIOUTBOARD S

CROSS-SECTION A

INBOARD SLAT

;

;

;

A

B

CROSS-SECTIOUTBOARD S


;

;

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

AOA < 16°5

A

B


;

;

;

INCREASING

AOA = 16°5

IGN

INCREASING

AOA = 23°

1

2

3

; ;




AOA<16°5

INITIALCONFIGUR-

ATION

AIRPLANE ANGLE-OF-ATTA

SLAT/ FLAPCONTROLOUT OFCLEAN

• ALL SLATS EXTENDED

• GREEN LIGHT STEADY

• ALL SLATS EXTENDED

• AUDIO WARNING ON

• GREEN LIGHTSTEADY

• RETRACTIONOF INBOARDSLATS

• AUTOMATIC IGNITION

• IGN LIGHT ON

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

A

CROSOUT


;

;

;

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

A

CROSOUT

CROSS-SECTION A

INBOARD SLAT

;

;

;

CONTROL 7° FLAPS +SLATS AIRPLANEAOA SLATSPOSITIO

AOA≥16°5

(INCREASING)

AOA≥23°

(INCREASING)

DECREASING

AOA = 23°

IGN

DECREASING

AOA = 16°5

4

5

1 2 3

; ;

; ;



STEADY IGN LIGHT ON

R ev i s i on 3

1 5 - 2 6

F

OR T RAI NI N G P URP O S E S

ONL Y

Figure 15-25. Automatic Extension o

OUT OFCLEAN

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

LEGEND

NO. 1 PRESSURE

ADC 1IAS LESS THAN265 KT

AOAABOVE


VANE

INBOARDVALVE

BOX

AUTOMATICDISTRIBUTO

INBOARDSLAT

SOLENOIDSELECTOR

VALVE

OUTBOARDSLAT

SOLENOIDSELECTOR

VALVE

EXTEND

NO. 2 PRE

RESTRIC

LEFT OUTBOARDLEFT

INBOARD

R E R E R E

AOA 11°



R ev i s i on 3

1 5 - 2 7

F OR T RAI NI N G

P URP O S E S

ONL Y

Figure 15-26. Automatic Retraction of Inbo

OUT OFCLEAN

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

LEGEND

NO. 1 PRESSURE

ADC 1IAS LESS THAN

265 KT

AOAABOVE


VANE

INBOARDVALVE

BOX

AUTOMATICDISTRIBUTO

INBOARDSLATSOLENOIDSELECTOR

VALVE

OUTBOARDSLATSOLENOIDSELECTOR

VALVE

EXTEND

NO. 2 PR

RESTRIC

LEFT OUTBOARDLEFT

INBOARD

R E R E R E

AOA 11°




If the slats are retracted at the time of testing,the outboard slats automatically extend, theIGN indicator lights on the overhead panelcome on, the audio warning is heard, and thegreen light flashes. If the slats are extended, theinboard slats automatically retract, the audiowarning sounds, and the green light flashes.

Emergency Extension ofthe Outboard Slats

If normal slat control or the No. 1 hydraulic

system fai ls, the guarded EMERGENCYSLATS switch located on the pedestal close tothe slat/flap handle (aircraft SNs 1 to 161) oron the slat/flap control box (aircraft SNs 162and subsequent) can be used to extend the out-board slats (see Figure 15-18).

As the outbslats movingilluminates, tended. The gand flashes aoperation is iand does not

On the ga b ov e slats noconfigu

AIRBRAK

GeneralThe airplanconsisting oposition of eels are actuatuators and thThe No. 2 hbrake systemas a backup

Extension ahandle on thelectrical sigvalves, one tuators and opanel actuatairbrakes is:

• Inboar

• Center

• Outbo

ControlTh i b k

Figure 15-28. STALL 1 and STALL 2Test Buttons



R ev i s i on 3

1 5 - 3 0


ONL Y

Figure 15-29. Emergency Ex

CLEAN

7° FLAPS-SLATS

20° FLAPS-SLATS

40° FLAPS-SLATS

LEGEND

NO. 2 PRESSURE

ADC 1IAS LESS THAN

265 KT

AOALESS


VANE

INBOARDVALVEBOX

AUTOMATICDISTRIBUTO

INBOARDSLAT

SOLENOIDSELECTOR

VALVE

OUTBOARDSLAT

SOLENOIDSELECTOR

VALVE

EXTEND RESTRIC

EMERGENCYACTUATOR

LEFT OUTBOARDLEFT

INBOARD

R E R E R E

RETURN



The handle has three positions labeled “0,”“1,” and “2,” each incorporating a ball-lock de-tent. In the 0 position, all surfaces are re-tracted (flush). Position 1 results in extension

of only the center panels; position 2 causes allpanels to extend.

Indications

System indication consists of two lights. Thered T/O CONFIG light on the warning panelilluminates when the airplane is on the ground,the inboard or center airbrakes are not re-

tracted, and one or more power levers are ad-vanced beyond the 84° PLA setting.

The amber AIR BRAKE light on the configura-tion panel comes on with the AIR BRAKE han-dle in position 1 or 2 and the airbrakes extended.It will flash when the airbrakes automatically re-

deenergizedthe airbrakesis applied toactuator pist(Figure 15-3

Moving the hthe center aNo. 2 systemcenter airbrathe larger arthe center aiflection of 5light comes

Moving the henergizes thselector valvside of the fboard and inlarger area ofboard surfacfaces to 68°. remains enerextended, anmains on.

Conversely, position deenselector valvboard surfac

Handle movethe center airsurfaces retrBRAKE light

Automat

If either AOcreasing abodeenergized,by the consta


Figure 15-30. AIR BRAKE Handle






FA

MOVING

LANDING GEAR

TEST

SLATS

;

;

;

;

;

;

UP

;

;

;

;

FLAPS

07

2040DN

;

;

;

;

;

;

;

;

;

;

;

;

;

;

XTK 2OPEN

XTK 2CLOSED

AILZERO

AILFEEL

A P M IS TR IM

NOSECONE OVHT

BLEEDAPU

PITCHFEEL

MACHTRIM

BAGISOL

BAGACCESS

T/OCONFIG.

NOSELH RH

DOWN

NORMAL

UP

RUDDER

TAILPLANE

AILERON EMERG

;

;

;

;

;

;

;

;

;

;

;

;

;

;

AIRBRAKES

0

1

CONFIGURATIONPANEL

TRIM/AIRBRAKECONTROL BOX

MASTERWARNING

PANEL

AUDIOWARNING

“NO TAKEOFF”

AIRPLANE ON GROUND+

ONE POWER LEVERBEYOND 84°

+AIRBRAKES NOT

RETRACTED

CONFIGURATION PANEL

MASTER WARNING PANEL

TRIM/AIRBRAKECONTROL BOX

AIRBRAKE IN EXTENDED POSITION

RIGHTANGLE-OF-ATTACK

VANE

LEFTANGLE-OF-ATTACK

VANE

BLINKER

RELAYS

AIRPLANEAOA 16.5°

AIRPLANEAOA

GREATER THAN16.5°

POSITION 1

POSITION 2

AIRPLANEAOA 16.5°

LEGEND

NO. 2 PRESSURE

RETRACT PRESSURE

ELECTRICAL

RESTRICTOR

RETURN

OUTBOARDAND INBOARD

AIRBRAKESOLENOID

SELECTOR VALVE

AIRBRAKE

CENTERAIRBRAKESOLENOIDSELECTOR

VALVE

AIRBRAKES-RETRACTEDPRESSURE-

HOLDINGVALVE

α

;

Figure 15-31. Airbrakes System



LIMITATIONSFull application of rudder and aileron con-trols and maneuvers that involve angles of at-tack near the stall must be confined to speedsbelow VA.

SECONDARY FLIGHTCONTROLS

High Lift Devices Operating orExtended Limit Speeds (VFE)

7° Flaps + slats.................................. 200 KIAS

20° Flaps + slats................................ 190 KIAS

40° Flaps + slats................................ 180 KIAS

AIRBRAKESAirbrakes must not be extended in flight within300 feet AGL.




1. Arthur units:

A. Function as variable bellcranks

B. Increase force required to satisfyAFU spring tension

C. Are installed in the aileron and ele-vator control linkages

D. All the above

2. The auxiliary AFU:

A. Provides load feel to the pilot

B. Provides a centering force to the ser-voactuator slide valves

C. Provides load feel to the pilot if themain AFU fails


3. The primary flight control servoactua-tors are:

A. Supplied with hydraulic pressurefrom the No. 1 system only

B. Suppl ied with hydraulic pressurefrom the No. 2 system only

C. Dual actuators receiving pressurefrom both hydraulic systems

D. Single actuators receiving pressurefrom both hydraulic systems

4. Electrically controlled and hydraulicallyactuated control surface trim is providedfor the:

A. Rudder only

B. Aileron and rudder

C. Aileron only

D. Aileron, rudder, and elevator

5. Pitch trim is accomplished with:

A El t i t d i i th h i

6. Automatic retoccurs:

A. When bothplane angle

B. As airspeed

C. When bothplane angle

D. When eithplane angle

7. Automatic exslats occurs:

A. As the airp

B. At high ang

C. At low angD. At high IA

8. The leading-edthe retracted p

A. Constantlysure on thactuators

B. Airload

C. Ball locks w

D. A spring-lo

9. The airbrakes a

A. Electrically

B. Automatica

C. Electricall

lically actuD. None of the


QUESTIONS



CHAPTER 16AVIONICS

CONTENTS

INTRODUCTION......................................................................

GENERAL .................................................................................

PITOT-STATIC SYSTEM..........................................................

General................................................................................

Pitot Pressure Probes ..........................................................

Static Pressure Ports............................................................

Temperature Sensor ............................................................

Engine Total Pressure Sensing............................................

PITOT-STATIC/AIR DATA INSTRUMENTS..........................

General................................................................................

Mach Speed Indicators........................................................

Altimeters............................................................................

Vertical Speed Indicators....................................................

Standby Mach Speed Indicator ..........................................

Standby Altimeter ...............................................................

ID 802 Annunciator ...........................................................

Cabin Differential Pressure Indicator .................................




Aileron Q and Elevator Arthur Unit Control..................................

Landing Gear Monitoring ..............................................................

Horizontal Stabilizer......................................................................

Cabin Pressure Control ..................................................................

Turbo Cooling Unit Control ..........................................................

AIR DATA COMPUTERS (ADC) ........................................................

General............................................................................................

ADC Monitoring ............................................................................

ADC Testing ..................................................................................

ADC Switching ..............................................................................

ANGLE-OF-ATTACK SYSTEM .........................................................

COMMUNICATIONS...........................................................................

General............................................................................................

Audio Panels..................................................................................

Control Heads ................................................................................

Radio Magnetic Indicator..............................................................

ELECTRICAL POWER SOURCES ....................................................

QUESTIONS .........................................................................................




ILLUSTRATIONS

Figure Title

16-1 Pitot-Static System......................................................

16-2 Pitot Probes.................................................................

16-3 Static Ports (Typical) ..................................................

16-4 Ram-Air Temperature Sensor .....................................

16-5 PT2 /TT2 Probe ..............................................................

16-6 Pitot-Static/Air Data Instruments Locations .............

16-7 Electric Mach Speed Indicator ...................................

16-8 Altimeter .....................................................................16-9 Vertical Speed Indicator..............................................

16-10 Standby Mach Speed Indicator ...................................

16-11 Standby Altimeter.......................................................

16-12 ID 802 Annunciator.....................................................

16-13 Cabin Differential Pressure Indicator ..........................

16-14 Altitude Preselect/Alert Characteristics .....................

16-15 Air Data System .........................................................

16-16 Angle-of-Attack System ............................................

16-17 Radio-Navigation Equipment Locations.....................16-18 Audio Panel (Typical).................................................

16-19 Radio and Intercom Connections ................................

16-20 Radio Magnetic Indicator ...........................................




INTRODUCTION

The Falcon 900 avionics consist of the pitot-static and aflight guidance and control system, and other associattions systems. This chapter includes the standard avioniexception of the automatic flight control system (AFC

will be handed out separately by the instructor. The ussupplements in the AFM and vendor manuals for addimation on specific systems not included in this chapter

GENERAL




Electrical power sources for the avionics are

listed at the end of this chapter under ElectricalPower Sources. Refer to Chapter 2, “Electri-cal Power Systems,” for additional information.

PITOT-STATIC SYSTEM

GENERAL

The pitot-static system serves the computerso f a S p e r r y a i r d a t a sy s t e m , t h e c a b i npressurization system, the stall warning andidentification system, a standby altimeter, astandby Mach speed indicator, and the aileronQ and elevator Arthur units.

Figure 16-1 shows the Falcon 900 pitot-static

system configured with the electric pilot’sand copilot’s Mach speed indicators.

PITOT PRESSURE PROBES

Three identical pitot probes supply total pres-sure information to the three systems (pilot,copilot, and standby). The probes supply sep-arate, independent pressures as depicted inFigure 16-1. The pilot’s and copilot’s pitotprobes (Figure 16-2) are located on each sideof the lower fuselage nose; the standby sys-tem probe is mounted directly below thepilot’s probe.

Each probe is equipped with an anti-icingheating resistor. Heating is controlled by threeswi tches on the overhead panel labeled“PILOT,” “COPILOT,” and “ST-BY.” ThePILOT switch controls heating of the left pitotprobe, the left and right static pressure ports(pilot’s and copilot’s), the left angle-of-at-tack sensor, and the temperature probe. TheCOPILOT switch controls heating of the right

Two drain ports are ltion of each tube.

STATIC PRES

The pilot’s and cop(Figure 16-3) are flof the airplane noseof each independenlot’s) are interconneffect. A flange at thto prevent port posiare very slight, espec


Figure 16-

Fli htS




FlightS in


ADC 1

STANDBY STATIC

STANDBY PITOT

PILOT PITOT

COPILOT STATIC

PILOT STATIC

MACH SPEEDINDICATOR 1

(OPTION)

HOR STABCRUISE STOP

ID 802 ANNUNCIATOR 1(TAT, SAT, TAS)

AILERON ARTHUR Q

MONITORING A/C SNs UP TO 178 ELECTRICALTIMETER1

SLATMONITORING

VERTICAL SPEEDINDICATOR 1

SLATCONTACTS

VMO /MMO

WARNING

LANDING GEARCONTACT

ALTITUDEWARNING

ATC 1 CONTROL PANEL

TRIPLE INDICATOR(CABIN ALT, ∆ P, AND

RATE-OF-CLIMB)

STANDBY MACHSPEED INDICATOR

AILERON ARTHUR CONTROLA/C SNs 179 AND HIGHER

Fli htS f t






PILOT’S VERTICAL

SPEED INDICATOR VMO /MMO

TESTPUSHBUTTONS

STANDBY

MACH SPEEDINDICATOR

* PILOT’S AND COPILOT’S ELECTRIC MACH SPEED INDICAT

ADC 1 ADC 2

;

;

;

RESET

20°c SAT 25°C TAT 175 KTAS

L AFCS FAIL

R AFCS MASTER

AP ENGAGED

L AFCS R AFCS

BRT

NAVIGATION

ADC1

XXX2.5

BUS A1

;

;

;

;

;

;

;

;

;

;

;

RIGHT CIR

;

;

;

;

;

PILOT’SALTIMETER

STANDBYALTIMETER

PILOT’S MACHSPEED INDICATOR*

PILOT’S ID 802 ANNUNCIATORLEFT CIRCUIT-BREAKER PANEL

PILOT’S EADI

019

46

37

28

5

00 0010

120

300

100

350

400

8060

6

4

5140

250160

200

180

1000FTPERMIN

VERTICAL SPEEDUP

DOWN

14

14

6

2

2

5

0

5

019

46

37

28

5

ALT

28 7 20

BARO

100

300

80

350

40060

120

250

140

200

230160

180

6 6

IASKT

180

160

120

180

100

80

LNAV HDG VASL

ASEL

300020

10

20

10

000

G

S

SPERRY

X

IRS

Figure 16-6.



TEMPERATURE SENSOR

The ram-air temperature sensor (Figure 16-4) is located on the right lower side of the air-plane nose section and is heated for anti-icing.The temperature information goes to both air

data computers.

ENGINE TOTAL PRESSURESENSING

PT2 total pressure is measured at each engine airintake (Figure 16-5) and supplies the respectiveengine fuel computer. Refer to Chapter 7,“Powerplant,” for additional information.

PITOT-STATIC/AIR DATAINSTRUMENTS

MACH SP

Standard-inMach speed pilot’s and cceive input f

1 and ADC


Figure 16-3. Static Ports (Typical) Figure 16-4



Indicated airspeed is displayed by a pointerread against a circular airspeed scale whichranges from 60 to 400 knots. A striped bar-ber-pole pointer is set to display VMO . Machdisplay is provided in a window in the rangefrom 0.1 to 0.99 in one-hundredth Mach in-

crements. The indicator has two built-in mon-itoring circuits—one for IAS and the other forVMO . The OFF flag comes into view in theevent of power supply failure or IAS dataerror; the VMO flag comes into view in caseof a VMO error. The flags are not visible dur-ing normal operation.

ALTIMETERSThe pilot’s and copilot’s altimeters receive in-puts from ADC 1 and ADC 2, respectively.They display altitude by means of a four-drum counter and a pointer (Figure 16-8).The extreme right (20-foot) drum displays

Barometric pressuBARO knob. The displayed on two comillibars and the rmercury. The corre30.99 inches Hg (9

tude warning light icorner of the instrufrom the preset altituremote controller. the counter to indicply or faulty instrum

VERTICAL SP

The pilots’ verticaare electrical and rtheir respective Atwo semicircular scfeet/minute X 1,000indicates both the ve


Figure 16-7. Electric Mach Speed Indicator Figure 1



STANDBY MACH SPEEDINDICATOR

The standby Mach speed indicator (Figure 16-10) receives pneumatic information directlyfrom the standby pitot and static pressure sys-tems. The IAS scale is gradated from 60 to 400

knots. At 200 knots and upwards, the IASscale is the inner scale, and the Mach scale is

the outer sca

ing on bothcated within

STANDB

The standbypneumatic standby statof displayinfeet. A threein 100-foot ithe 10,000-fand-white sthis drum istriped scresure is set in of the instru

ID 802 AN

Temperature(TAS) informADC 2 and iID 802 annuinstrument (can be instal

side.


Figure 16-9. Vertical Speed Indicator



TAT and SAT are permanently displayed in 1°increments from –99 to +99° C. TAS is per-manently displayed in 1-knot increments from0 to 999 knobs. Display brightness is controlledby the BRT knob on the face of the instrument.

The ID advisory display messages are dividedinto five categories:

• AFCS operational status

• AP and YD engage status

• Mode inhibits and error

• System failures and cautions

• SAT/TAT/TAS

White air data messages on line one are dis-

played continuously. Line two displays ambermessages which either flash or come on stead-ily; these require the pilot to clear the messagewith the RESET button or, in some cases, bydisengagement of the autopilot.

CABIN DIFFERENTIALPRESSURE INDICATOR

The cabin differential pressure indicator(Figure 16-13) is a triple indicator which dis-plays cabin altitude, differential pressure,and rate of climb. It is pneumatic and con-nected to the standby static pressure system.Refer to Chapter 12, “Pressurization,” for ad-

ALTITUDE WA

The ADC which is tion to the autopilotvisual and audio waairplane is deviatinon the EFIS instrum

The altitude preset 1,013 mb corrected f1 supplies informatdirector, the referendisplayed on the piADC 2 i l d


Figure 16-13.

RESET

20°c SAT 25°c TAT 175KTAS

L AFCS FAIL

R AFCS MASTER

AP ENGAGED

L AFCS R AFCS

BRT

Figure 16-12. ID 802 Annunciator



A visual warning is given by a light in the up-

per right corner of each altimeter and on bothof the EADIs. An audio warning sounds for onesecond as soon as the visual warning appears.

Altitude warning is given when the airplanealtitude is within 1,000 feet of and more than250 feet from the prese t altitude. Warning isalso given when the airplane altitude is lowerthan the preset altitude by at least 250 feet and

at most 1,000 feet. Figure 16-14 shows graph-ically when the audio and light warnings occur.

PITOT-EQUIPM

MACH SP

GeneralMach speedspeed signa

The signal aever spee


+ 1,000 FT

A L T I T U D E

D E V I A T I O N

S E L E C T E D

A L T

S E L E C T E D

A L T

A P P R O A C H I N G

A P P R O A C H I N G

+ 250 FT

0 FT

WARNING LIGHT(S)

GOES OUT

WARNING LIGHT(S)

AUDIO WARNING

WARNING LIGHT(S)COMES ON

AUDIO WARNING(FOR 0.75 SEC)

WARNING LIGHT(SGOES OUT

WARNING LIGHT(S)COMES ON

- 250 FT



VMO /MMO is indicated by the striped barber

pole on the Mach speed indicators. The warn-ing signal is modulated with rising pitch forone second.

VMOVMO is linear between 350 and 370 knots fromsea level to 10,000 feet. From 10,000 to 25,000

feet the value is constant at 370 knots.

MMOThere are two airplane gross weight laws:

• For gross weights over 35,000 pounds,MMO is 0.87 from 25,000 to 33,000 feet.It is 0.84 from 37,000 to 51,000 feet. Var-

iation between these two points is lin-ear.

• For gross we igh t s l ess than 35 ,000pounds, MMO is 0.87 from 25,000 to38,000 feet. It is 0.84 from 42,000 to51,000 feet. Variation between thesetwo points is linear.

There are two possible installation configu-rations for M

MOswitching control:

• A switch is located to the left of the co-pilot’s RMI; the pilot manually controlsswitching.

• A manual switch is not installed. FMScomputers control switching automati-cally once the gross weight falls below35,000 pounds (provided that the ini-

tial airplane gross weight was input).Both ADCs receive the switching signal.

SLAT MONITORINGAND CONTROL

Auto slat extension is inhibited by a signal

AILERON Q AARTHUR UNITThe aileron Arthur formation from thepressure lines on aior from ADC 2 on aADC 1 on SNs 179

The elevator Arthur

both ADCs (speed hknots). The speed logic circuit controlon the warning pancomes on if the aileto the low-speed AIAS is equal to or h

LANDING GEAThe speed contact landing gear warninlanding gear controless than 160 knots,not in the downlockepower lever is in th

HORIZONTALA speed contact, plimits horizontal stion to –4° when IAto 210 knots.

CABIN PRESS

The automatic cabceives altitude andboth ADCs.

TURBO COOLUNIT CONTRO




f rom the rema in ing ADC i f t he

IAS/Mach function on the corre-sponding EFIS reversion controlleris selected. Refer to EFIS Rever-sionary Switching for additional in-formation.

ANGLE-OF-ATTACK

SYSTEMThe Teledyne angle-of-attack system senseslocal angle of attack along with airplane flapand slat position, converts these inputs into liftinformation, and displays it on a normalized,ten-gradation, zero (0) to one (1.0) dial (Figure16-16). This display has zero lift placed atzero on the dial, and the maximum usable lift

point, as defined by the AFM stall speed, is dis-played as 1.0. In effect, the indicator presentslift as a percentage. With flap and slat posi-tion information, the display is valid for all air-plane configurations. Therefore, the systemcontinuously displays the margin to the max-imum usable lift point (1.0), regardless of air-plane weight , bank angle, G loading, orairplane configuration.

The REF SET knob on the face of the AOA in-dicator manually positions the adjustable refer-ence index and, at the same time, positions areference pointer located approximately oppo-site the adjustable reference index on the faceof the indicator. The adjustable reference indexcan be moved to any position between .15 and.70 on the percentage scale, or alte rnatively,

the reference pointer can be positioned to aspecific point on the V/VS scale. The rela-tionship between the reference index providesa cross-reference between V/VS ratio and nor-malized AOA scale points (i.e., setting thereference pointer to V/VS 1.45, .52, etc.). For

t i l f ll fl 1 3 V h th V/V

COMMGENERA

CommunicaFalcon 900 dress, air-toVoice comm

• Interc• Cabin

• Radio

• VHF

• HF

• FM

Navigation c

• VOR/I

• DME

• ADF

• ATC

Different com

for use on th(audio paneFigure 16-17navigation e





ATTACK

OF

REFSET

1.0

.8

.6

.4

.20

1.5

1.4

1.3

1.2

REFERENCE POINTER

V/VS SCALE

ZERO LIFT

REF SET KNOB

AOA VANE

ANGLE

V / VS




ICSLH

5VHF

1

10PWR

1

25

CONTROL

LEFT CIRCUIT–BREAKER PANEL

MASKMIKEJACK

HEADSETJACK

HEADSET RAILMIKE JACK

LOUDSPEAKER

ANNUNCIATOR ANDHF TRAFFIC LIGHTS

;

;

;

;

;

;

ANNUCIATORS TRAFFIC

VHF 1 FWD HF

HF 1 MID XFR

VHF 2 AFT HF 1

HF 2 LAV HF 2

VHF 1TEST VIP

PILOTCONTROL WHEEL

PILOT PTTBUTTON

COPILOTCONTROL WHE

HANDMIKE

JACK

L

MICROPHONE

EMG SPK ST VOL NAV 1 NAV 2 FILT DME 1

MKR 1 MKR 2 ADF 1 ADF 2 FILT DME 2

VHF 1 VHF 2 VHF 3 HF 1 HF 2 FM MASK CPIT

VHF 1 VHF 2 VHF 3 HF 1 HF 2 FM PAGE CABIN

AUDIO

AUDIO

5

MASKMIKEJACK

HEADSEJACK

HEADSET MIKE J

PILOT MIXER UNIT



AUDIO PANELS

General

The pilot’s and copilot’s audio panels aremounted in the left and right side consoles, re-spectively (Figure 16-18). They control trans-mitting and receiving of all radios on theFalcon 900. They are independent panels pow-ered by separate buses (pilot’s, A2; copilot’s,

B2). Figure 16-19 graphically shows radioand intercom connections.

The audio panels installed are Baker mixerunits. They have two types of pushbuttons forcommunication control:

• Interlocking pushbuttons for the MI-CROPHONE channels

• Separate pushbuttons for the AUDIOreception channels

The MASK annot interlock wphone pushbut

Except for MASKonly one button canthe row of MICRObutton is depressed

• Hand mike abutton

• Headset rail mleased)

• Mask mike pressed)

The use of one of causes reception m

The headset ramike are contrtalk button on

Depressing PAbuttons releaselected pushbut

Microphone C

PAGE pushbutton—public address in thdepressed, it causeheadsets or cockpit lton depressed), excsystem modulation independent of the S





FILT pushbuttons—The filter is used to elim-

inate the 1,020-Hz frequency from NAV andADF signals. There are two FILT pushbut-tons: one, located between NAV 2 and DME1, filters the VOR channels when depressed;the second pushbutton, located between ADF2 and DME 2, filters ADF only.

When the FILT pushbuttons are depressed,the VOR and ADF signals are adjusted by the

VOL knob and can be attenuated. When theyare released, the VOR signals are adjusted bythe VOL knob and are attenuated by the mut-ing system; the ADF signals are not adjustedby the VOL knob but can be attenuated by themuting system. The muting system is opera-tional during VHF, HF, FM, or public addresstransmissions (in the latter case, with thePAGE pushbutton depressed).

MKR 1, MKR 2 pushbu ttons—When de-pressed, the marker signals are received with-out volume adjustment by the mixer unit. Thesignal level is not adjusted by the VOL knobbut can be attenuated by the muting system.

Muting System Operation

The muting system is operational during thefollowing transmissions:

• VHF

• HF

• FM

• Public address

The following reception channels are attenuated:

• VHF 1, 2, and 3

• HF 1 and 2

• FM

The receptio

ated during

• Audio

• Interc

CONTRO

General

The type of ccon 900 mayor preferencthe center ptwo COMs, tders. This cwhich are ty900.

Collins VControlThe VIR-3VOR/ILS nVOR, localicon.

An accurateceived signatection fromDigital frequmance, redu

Collins CControl

The CTL-22dition to thesix additionmemory, to bthis switch inumber is b



25-kHz channel is available whenever de-

sired.

Collins CTL-62 ADFReceiver ControlThe CTL-62 ADF control provides variablerate, two-knob tuning. In addition to the ac-tive and preset frequencies displayed, mem-ory storage is provided for four additional

frequencies.

Collins CTL-92 ATCTransponder Control

The CTL-92 ATC transponder control allowsvariable rate, two-knob tuning and a pilot-se-lectable preset code. The transponder control

also alerts the pilot when an emergency codehas been selected by flashing the selectedcode. This gives the pilot an opportunity to en-sure that the code is correct before transmit-ting. The transponder control can also displayencoded altitude.

RADIO MAGNETIC INDICATOR

The bearing/relative bearing of an ADF 1 or2 beacon is displayed on RMI 1 or 2 by meansof a single green pointer (ADF 1) or a doubleyellow pointer (ADF 2). Source (ADF/VOR)selectors 1 and 2 (Figure 16-20) are located onthe face of the instrument. The heading index(lubber line) is at the top center. DME 1 and2 digital readouts are displayed in two win-dows above the compass rose.

ELECTRICA

SOURCESFigure 16-21 show


Figure 16-20.



R ev i s i on2

1

6 - 2 1

F OR T RAI NI N G P

URP O S E S

ONL Y

Figure 16-21. DC P

CENTERCIRCUIT-BREAKER

PANEL

LEFT CIRCUIT-BREAKER

PANEL

1 5 0 A

1 5 0 A

GA L L E Y 1 B A R

A 4

GA L L E Y 2 B A R

A 6

P I L OT F R ONT

WI ND S HI E L DA 3

S T A NDB Y HY DRA UL I C

P UMP A 5

1 5 0 A

8 0

A

LEFT MAIN BUS

MAIN BUS-TIEROTARY SELECTOR

225A

FLIGNOR

BUS (4PA)

(14P)

TO EXT

POWER

CONTACTOR

START BUS

BUS A1

OFF

L AVMASTER

AVIONICMASTER

BUS A2

1 3 0 A

1 3 0 A L B

U S

T I E D

TIE

D

BAT20

0

25

30V-



CHAPTER 17

MISCELLANEOUS SYS

CONTENTS

INTRODUCTION .....................................................................

OXYGEN SYSTEM ..................................................................

General................................................................................

Crew System.......................................................................

Passenger System................................................................

Passenger Oxygen Controller .............................................

POTABLE WATER SYSTEM (SN 3)........................................

General................................................................................

Operation ............................................................................

POTABLE WATER SYSTEM (SNs 4 AND SUBSEQUENT) .

General................................................................................

Operation ............................................................................

WATER WASTE SYSTEM .......................................................

QUESTIONS..............................................................................




ILLUSTRATIONS

Figure Title

17-1 Oxygen System Components Locations .....................

17-2 Crew Oxygen System..................................................

17-3 Use of Smoke Goggles................................................

17-4 Use of Emergency Escape Breathing Device..............

17-5 Passenger Oxygen System ..........................................

17-6 PASSENGER OXYGEN Controller Controls and Fun

17-7 Passenger Oxygen Mask and Box (Typical) ...............

17-8 Potable Water System Schematic (SN 3) ....................

17-9 Potable Water System Components ............................

17-10 Potable Water System Controls and Indicators (SN 3)

17-11 Potable Water System Schematic (SNs 4 and Subsequ

17-12 Valve Control Safety Access Door .............................

17-13 Toilet Disposal And Drain System .............................




INTRODUCTION

The Falcon 900 oxygen system is supplied by a single highposition has a quick-donning mask with a built-in regulat

masks automatically drop out of overhead storage boxes can be manually released. Passengers can receive two diffthe setting on a controller on the copilot’s console. Twovided in the cabin.

The water system provides potable water to the galley, th

16 °12

8

4

0

NO 1 FUELTRANS

NO 1 FUELLOW

BATTHOT


MISCELLANEO




OXYGEN SYSTEM

GENERAL

The oxygen system uses gaseous oxygen storedin a high-pressure cylinder and provides crew-members and passengers with low-pressureoxygen as required. There are three separatesystems: crew, passengers, and first-aid.

The oxygen supply system consists of a high-pressure cylinder f i t ted with a pressure-reducing valve, a shutoff valve, a f i l lerconnection, two high-pressure gages, and anoxygen controller on the copilot’s console.Except for the controller and the high-pressuregage mounted on it, all of the components arelocated under the left cabin floor, aft of the pas-senger floor (Figure 17-1).

Oxygen Cylinder

The high-pressure oxygen cylinder has a totalcapacity of 2,200 litres (NTPD) under 1,850 psi(127.5 bars) at 21°C. Minimum usable pressureis 170 psi (12 bars). An optional oxygen cylinderwith a total capacity of 3,310 litres is also

available.

Pressure-Reducing Valve

The pressure-reducing valve functions to reducethe oxygen pressure to 70 psi (4.8 bars) for use inthe passenger and crew low-pressure systems. A90-psi safety valve prevents overpressurization inthe low-pressure system. The pressure-reducingvalve also incorporates a blowout disc sensingcylinder pressure. The disc is calibrated torupture at 2,750 to 3,050 psi (191 to 213 bars).

Shutoff Valve

When the shutoff valve on the cylinder is closed,

Cylinder High-P

A high-pressure gacylinder pressure. Irange graduated frorange extending fro

Copilot’s HighMounted on the PAStroller, the gage hasated from 0 to 2,2000 to 200 psi, an ambpsi, and another am2,200 psi.

The red range indicin the cylinder; ifimmediate descenamber range from 2(at 21°C) representcylinder pressure foout passengers on bsure of 700 psi permtwo hours for each cnot be undertaken wless than 700 psi at a21°C (70°F). At thisto isolate the pasConsidering the sizthis situation couldleak in the passengder was insufficient

The second amber rindicates that the othe pressure rating pressure must be moessary, to prevent bl

CREW SYSTE

Oxygen Mask






CYLINDER

PRESSURE

20

15

10-100

5

0

PSI

X

O

YGEN

RED RANGE A

HIGH-PRESSUR





THIRD CREWMEMBER’SMASK BOX

COPILOT’S MASK

TEST SOCKET

MASK BOX

OXYGENCONTROLLER

MASK BOX

TO PASSENGEOXYGEN SYST

SMOKE HOOD

PILOT’S MASK

COPILOT’SCONSOLE

PNEU(SHOW

TEST BUTTON ANDEMERGENCY OVERPRESSURE

MANUAL CONTROL




Figure 17-4. Use of Emergency Escape Breathing Device

P U L L

T O A C

REMOVE UNIT FROM STORAGE CONTAINER. TEAR OFF RED PULL STRIP, AND REMOVE UNITFROM PLASTIC PROTECTIVE WRAPPER.

PULL ACTUATION RING IN THE DIRECTIONINDICATED.

HOLD THE DEVICE BY THE OPEN END OF THEHOOD WITH THE LIFE SUPPORT PACK AWAYFROM THE USER.

BEND OVER AND GRASP HOOD OPENING WITHTHUMB, AND PULL HOOD OVER HEAD.

RAISE TO STANDING POSITION, AND ADJUSTHOOD AND LIFE SUPPORT PACK FOR MOSTCOMFORTABLE FIT. CHECK NECK SEAL FORSECURE FIT.

1 2

3 4

5 6




PASSENGER OXYGEN

Controller

The controller (Figure 17-5) provides manualor automatic operation of the passenger oxy-gen system. It is supplied with oxygen regu-lated at 70 psi. Below 16,000 feet, it deliversoxygen at 19 psi (1.3 bars); above 16,000feet, it provides 70 psi (4.8 bars) oxygen pres-sure. Between the altitudes of 11,500 and16,000 feet, initial pressure is 70 psi for 5 sec-onds in order to release the masks, then it is

automaticalpressure mocompared wvalve, and irelated cont

The controlllection of NAID, or CLOa high-pressu

indicator, anation is pres

CYLINDER

PRESSURE

20

X

O

YG

EN

15

10

-100

5

0

PSI

ONORMA

CLOSED

FIRST AID

PASS ON

PASSENGER OXYGEN

SELECTOR

NORMAL

FIRST AID

PASS ON

OVERRIDE

CLOSED

HALF OPEN

HALF OPEN

HALF OPEN

ORFULLY OPEN

FULLY OPEN

FULLY OPEN

PASS ON

PASS ON

INDICATOR FUNCT

CABIN DEPRESSURIZATION—EMERMASK BOXES UNDER 70 PSI (4.8 BASWITCH TO FIRST AID FOR SUPPLY

OXYGEN SUPPLY AT 19 PSI FOR CO(MASK BOXES DO NOT OPEN). CABFUNCTION IS THE SAME AS WITH S

(OPENING OF MASK BOXES).

NORMAL FLIGHT—THE OXYGEN COREADY TO OPERATE (SUPPLY SYS

CABIN ALTITUDE ABOVE 16,000 +2,0

CABIN DEPRESSURIZATION—AUTOAND OXYGEN SUPPLY PER TWO PRCABIN ALTITUDE BETWEEN 11,500PRESS = 19 (1.3 BAR)




Passenger Mask Boxes

Each passenger mask box (Figure 17-7) housesa passenger mask and is equipped with a dualpressure-operated latch, an internal door tomaintain the mask in the stowed position, anda magnet-held cover.

Rated operating pressure of the latch is 29 to58 psi (2 to 4 bars). It is not actuated when thesystem is operating in the low-pressure mode

(19 psi). If the latch fails to actuate automat-ically, the cover can be manually removed.Cover removal, automatic or manual, allowsthe mask to drop out.

The dual mask box in the lavatory houses twomasks. When oxygen pressure builds to 70

psi, the latch is acturelease the internapushes the masks ouof the passengers.

Passenger MaThe passenger maska nosepiece and a ma breath-in/breath-o

air valve. Oxygen iliter economizer baof the mask is reguwith the mask box on the supply presrate is 2.8 to 3 liters to 0.875 liters per m

PRESSURE-OPERATEDLATCH

OXYGEN VALVEOPENING PIN

CORD

MASK




VIDEO

CEILING

LIGHT

LIGHT

AFT

DRAIN

HEATER

INVERT

XXXXX

FLUSH

TABE

VALANCE

FWD

10 5 10

7.5 15 5

7.5 5

WATER

HEATER

20

CIRCUIT-BREAKERPANEL

WASHBASINCABINET

WATER PUMP

PUSHBUTTON

REFILLING PORT

VENT PORTGALLEY CABINET

LIGHTS–MISC

GALLEY

MASTER

CAPT WH

READING

INSTR

RH

XX AFT CABIN

XXXX

LH CABIN

READING

RH CABIN

READING

FWD CABIN

XX

XX

XX

XX

XXXXX

XX XXXX

XXXX

LANDING

RM

ANTI???

BELLY

XXX

LAV

MASTER

Figure 17-10. Potable Water System Controls and Indicators (SN 3)




POTABLE WATER

SYSTEM (SNS 4 ANDSUBSEQUENT)

GENERAL

The potable water system stores water in a14.5-gallon (55-liter) water tank (Figure 17-11), located beneath the aft lavatory vanity, fordistribution of water under pressure to the aftvanity, forward galley, and optional forwardlavatory if installed.

The system includes the water tank, regulatorvalves, pressure and quantity gages, waterheater, air accumulator, shutoff valves, filland vent valves, and drain mast.

The system the aft lavatulated enginair chargingleft baggagesure gage ontank pressurquire a full the system; hlator air prewater that cstart.

OPERAT

Delivery With the air aat approxima

supply on/of

LEGEND

BLEED AIR

APU ENGNO. 2

ENGNO. 1

RESTRICTOR

RELIEFVALVE

(120 PSI)

AIRCHARGINGVALVE

AIRACCU

WATER HEATER

TO CABIN WATERDISTRIBUTION SYSTEM

(GALLEY, ETC.)

VANITY SINKAND FAUCETS

FILTERREGULATOR

(120 PSI)




system control panel and then into the top of thetank. Water tank pressure is also indicated on apressure gage in the vanity panel.

Water from the tank is routed through a sup-ply on-off valve and is made available to theaft vanity, forward galley, and optional for-ward lavatory. After an engine is started, aconstant air pressure source is available fromthe bleed-air system. Hot water under pres-sure is available to the vanity sinks, providedthe hot water circuit breakers behind the aftlavatory toilet seat and on the center cockpitoverhead panel are engaged.

Tank FillingThe tank can be filled from inside the airplanethrough the filler neck. Before moving the

cap, pressing down on the cap relieves airpressure. The fill and vent valve ports areused for filling the tank from an outside watersource. A push-pull handle in the external ser-vicing panel in the aft left wing root fairingactuates the valve.

Draining

Prior to draining the storage tank, trip theheater circuit breakers, close the air shutoff valve, and open the forward and aft vanitysupply valves. If tank pressure drops to zero,vent the system by pulling the push-pull han-dle to open the fill and vent valve.

NOTE

Detailed instructions for operation,filling, and draining of the systemare available in the “Water SystemInformation’’ section of the Falcon200 Supplemental MaintenanceManual

WATER WA

Disposal of water lavatory compartmeand the water tank. Ffrom these areas ttected by anti-icing

The water tank is gvent pipe leading to

the drain shutoff vathe outside drain through the drain (Figure 17-12) preof the two valves in

The toilet unit incling 9.6 U.S. gallo(Figure 17-13).

A drain valve is cona drain port beneathbe actuated to manuwhen disposal of toflushing port is usand to refill it when

Complete instructio

ing and servicing are38-310 of the Main




CLOSED

OPEN

GALLEY WATER S

SHUTOFF VA

DOO

DOOR

ISOLATIONVALVE

Figure 17-12. Valve Control Safety Ac




LOOKINGFORWARD

CABIN FLOOR

DRAINPAN

DRAIN

VALVE

SEALEDDUCT

LEVER

FLUSHING PORT

DRAIN PORT

Figure 17-13. Toilet Disposal and Drain Syst




1. Oxygen cylinder pressure is read on:

A. Two gages: one in the cockpi t

and one nea r the oxygen cylinder

B. One gage in the cockpit

C. One gage near the oxygen cylinder

D. Gages at both pilot stations and near

the oxygen cylinders

2. With the PASSENGER OXY GEN

controller selector knob in NORMAL,

oxygen is available:

A. Only to the crew

B. Only to the passengers

C. To the crew and the first-aid stations

D. To all stations throughout the

airplane

3. The PASS ON indicator shows:

A. Operation at 19 psi only

B. Operation at 70 psi only

C. Operation at 19 or 70 psi


4. The passenger oxygen masks deploy:

A. When the selector knob is placed in

NORMAL

B. When the selector knob is placed in

OVERRIDE

C. When the selector knob is placed inFIRST AID

D. If cabin pressure alt i tude is above

11,000 feet

6. The em

A. Con

B. Useoxy

C. Conmashose

D. Is togog

7. The uppt o 2 ,2 0OXYG

A. Nor

B. The

hauC. Presexce

D. Non

8. The pot

A. Fromby g

B. O n lsour

C. Onlthe

D. Non

9. Potable

A. The

B. TheC. The

D. All

10. The sel

A U

QUESTIONS




WALKAROUNDThe following section is a pictorialround. It shows each i tem cal led ouexterior power-off preflight inspectifold-out page at the beginning of theround section should be unfolded befoing to read.

The general location photographsspecify every checklist item. Howevitem is portrayed on the large-scatographs that fol low.



1. NORMAL STATIC PORTS:

COVER/CONDITION—REMOVE/CHECK

4. ANGLE-O

COVER/CO

5. STANDBY

COVER/CO

2. PI LOT P ITOT PROBE:

COVER/CONDITION—REMOVE/CLEAR

FALCON 900 WALKAR




7. COCKPIT W INDOWS—CHECK/CLEAN

WINDSHIELD WIPER—CHECK/STOWED

NOSE CONE (2 LATCHES)—CLOSED/LOCKED

11. STANDBY STATIC

COVER/CONDITIO

10. TEMPERATURE PR

COVER/CONDITIO

8. NOSE LANDING GEAR: HYDRAULIC LEAKS,

SHOCK ABSORBER HEIGHT, ANTISKID

TACHOMETER CONNECTOR (AS REQUIRED),

TIRE CONDITION, AND TORSION LINK PIN IN-

STALLED—CHECK




13. COPILOT PITOT PROBE:


16. BELLY AN

(STANDAR

14. NORMAL STATIC PORTS:

COVER/CONDITION—REMOVE/CLEAR

17. ICE DETE




22. EMERGENCY EXIT

EMERGENCY EXIT19. PARKING BRAKE ACCUMULATOR

PRESSURE—CHECK 1,000 PSI

23. RIGHT ENGINE AIR

COVER/CONDITIO

20. FUEL VENT VALVE—CLEAR/NO LEAKS




25. EMERGENCY EXIT LIGHT IN LEADING

EDGE—CHECK28. WING CHE

WING FUE

26. LANDING GEAR: TIRE CONDITION, HYDRAULIC

LEAKS, BRAKE WEAR, AND SHOCK ABSORBER

HEIGHT—CHECK

29. LEADING-




31. STATIC DISCHARGERS (4)—CHECK

36. PRESSURE REFUETROL PANEL ACCE32. AILERON/FLAPS—CHECK

33. AIRBRAKES—CHECK




38. HEAT EXCHANGER DUCT AIR INLET—CHECK 41. HEAT EXC

39. ENGINE NACELLE AND PYLON (3 LATCHES) AND

DRAINS—CHECK

42. REAR CO




44. ENGINE OIL AND FUEL FILTERS

CLOGGING—CHECK

ENGINE OIL LEVEL—CHECK

47. MAIN ELECTRICAL

OPTION)—CHECK

45. NO. 1 HYDRAULIC ACCUMULATOR

PRESSURE—CHECK

DOORS—CLOSED/CHECK

48. S-DUCT ACCESS D

CENTER ENGINE F

S-DUCT ACCESS D




50. NO. 2 HYDRAULIC RESERVOIR LEVEL—CHECK 53. NO. 2 HYD

BUBBLES

51. NO. 2 HYDRAULIC FILTER PLUGS—CHECK 54. FIRE EXT




56. NO. 1 HYDRAULIC RESERVOIR LEVEL—CHECK 59. CENTER ENGINE C

57. NO. 1 HYDRAULIC FILTER PLUGS—CHECK

HYDRAULIC BAY ACCESS DOOR—CLOSE/CHECK

60. ITEMS ON BOARD

LADDER (OPTION)

REAR COMPARTM




62. RIGHT ENGINE TAILPIPE:


PYLON STATIC DISCHARGER—CHECK

64. VERTICAL

ANTENNA

65. NAVIGATI




68. LEFT ENGINE TAILPIPE:


PYLON STATIC DISCHARGER—CHECK

67. CENTER ENGINE COWL

(3 L ATCHES)—CLOSED/LOCKED

70. APU GENERATOR

OUTLET—CHECK

APU/REAR COMPA

CHECK

71. NO. 1 HYDRAULIC

DOOR—CLOSED/C




73. ENGINE NACELLE AND PYLON (3 LATCHES) AND

DRAINS—CHECK

BAGGAGE COMPARTMENT ENTRANCE

LIGHT—CHECK

79. STATIC DI

74. FUEL TANK PRESSURIZATION DRAINS—

CHECK/NO LEAKS

75. WATER SYSTEM DRAIN MAST—CHECK

76. WATER SYSTEM FILLING AND CONTROL PANEL

ACCESS DOOR—CLOSED/CHECK

80. WINGTIP

NAVIGATI




82. WING CHECK FOR FUEL LEAKS—CHECK

WING FUEL SUMP DRAIN—CHECK FOR LEAKS

85. LEFT ENGINE AIR

COVER/CONDITIO

HEAT EXCHANGER

83. GRAVITY-FUELING PLUG AND GRAVITY-FUELING

BONDING—CHECK

86. FUEL VENT VALVE






88. OXYGEN SYSTEM FILLING AND

VENT—CHECK/CLEAR

90. CABIN ACCESS DOOR—CHECK

89. OXYGEN VALVE AND PRESSURE—OPEN/CHECK

ANSWERS TO QUEST




ANSWERS TO QUEST

CHAPTER 2

1. C

2. B

3. D

4. D

5. C

6. B

7. D

8. C

9. A

10. D

11. C

12. D

CHAPTER 31. C

2. A

3. B

4. D

5. A

6. C

CHAPTER 4

1. B

2. A

3. C

4. A

5. C

6. A

CHAPTER 5

1. A

2. D

3 C

10. D

11. D

12. D

13. A

14. C

15. A

16. B

17. C

18. D

19. D

20. B

21. D

22. C

CHAPTER 6

1. C

2. D

3. A

4. D

5. C

6. A

7. B

8. C

9. A

10. B

CHAPTER 7

1. B

2. D

3. A4. C

5. A

6. C

7 B

3.

4.

5.

6.

7.

CHAPT

1.

2.

3.

4.

5.

6.

7.

8.

9.

10.

CHAPT

1.

2.

3.

4.

5.

6.

7.

8.

CHAPT

1. 2.

3.

4.

5


CHAPTER 16



1. A

2. D

3. C

4. B

CHAPTER 17

1. A

2. D

3. C4. D

5. B

6. A

7. C

8. A

9. D

10. C

APPENDIX A




APPENDIX A

This appendix contains the following conversion tables:

Table

A-1 Conversion Factors .....................................................

A-2 Fahrenheit and Celsius Temperature Conversion .......

A-3 Inches to Millimeters ..................................................

A-4 Weight (Mass): Ounces or Pounds to Kilograms........

A-5 Weight (Mass): Thousand Pounds to Kilograms ........




Table A-1. CONVERSION FACT

Multiply

centimeterskilogramskilometerskilometersliterslitersmetersmetersmillibarsfeetgallons

inchesin. Hg (32° F)nautical milesnautical milespoundsquarts (liquid)statute milesstatute miles

By

0.39372.20460.6210.5390.2641.0539.373.2810.029530.30483.7853

2.5433.86391.1511.8520.45360.9461.6090.868

To

incpostanagaquincfeein.melite

cemistakilkillitekilna




Table A-3 INCHES TO MILLIMETERS (0 0001 I



Table A-3. INCHES TO MILLIMETERS (0.0001 I

INCHES 0.0000 0.0001 0.0002 0.0003 0.0004 0.0005

MILLIMETERS

0.000 0.0025 0.0050 0.0076 0.0101 0.0127 0.001 0.0254 0.0279 0.0304 0.0330 0.0355 0.0381 0.002 0.0508 0.0533 0.0558 0.0584 0.0609 0.0635 0.003 0.0762 0.0787 0.0812 0.0838 0.0863 0.0889 0.004 0.1016 0.1041 0.1066 0.1092 0.1117 0.1143

0.005 0.1270 0.1295 0.1320 0.1346 0.1371 0.1397 0.006 0.1524 0.1549 0.1574 0.1600 0.1625 0.1651 0.007 0.1778 0.1803 0.1828 0.1854 0.1879 0.1905 0.008 0.2032 0.2057 0.2082 0.2108 0.2133 0.2159 0.009 0.2286 0.2311 0.2336 0.2362 0.2387 0.2413

INCHES 0.000 0.001 0.002 0.003 0.004 0.005

MILLIMETERS

0.00 0.025 0.050 0.076 0.101 0.127 0.01 0.254 0.279 0.304 0.330 0.355 0.381

0.02 0.508 0.533 0.558 0.584 0.609 0.635 0.03 0.762 0.787 0.812 0.838 0.863 0.889 0.04 1.016 1.041 1.066 1.092 1.117 1.143

0.05 1.270 1.295 1.320 1.346 1.371 1.397 0.06 1.524 1.549 1.574 1.600 1.625 1.651 0.07 1.778 1.803 1.828 1.854 1.879 1.905 0.08 2.032 2.057 2.082 2.108 2.133 2.159 0.09 2.286 2.311 2.336 2.362 2.387 2.413

INCHES 0.00 0.01 0.02 0.03 0.04 0.05

MILLIMETERS

0.0 0.254 0.508 0.762 1.016 1.270 0.1 2.540 2.794 3.048 3.302 3.556 3.810 0.2 5.080 5.334 5.588 5.842 6.096 6.350 0.3 7.620 7.874 8.128 8.382 8.636 8.890 0.4 10.160 10.414 10.668 10.922 11.176 11.430

0.5 12.700 12.954 13.208 13.462 13.716 13.970 0.6 15.240 15.494 15.748 16.002 16.256 16.510 0.7 17.780 18.034 18.288 18.542 18.796 19.050 0.8 20.320 20.574 20.828 21.082 21.336 21.590 0.9 22.860 23.114 23.368 23.622 23.876 24.130

INCHES 0.00 0.1 0.2 0.3 0.4 0.5

MILLIMETERS

0 2 54 5 08 7 62 10 16 12 70




0 1 2 3 4 5 6

kg kg kg kg kg kg kg

oz

0 - 0.028 0.057 0.085 0.113 0.142 0.170

10 0.283 0.312 0.340 0.369 0.397 0.425 0.454

lb

0 - 0.45 0.91 1.36 1.81 2.27 2.72

10 4.5 5.0 5.4 5.9 6.4 6.8 7.3

20 9.1 9.5 10.0 10.4 10.9 11.3 11.8

30 13.6 14.1 14.5 15.0 15.4 15.9 16.3

40 18.1 18.6 19.1 19.5 20.0 20.4 20.9

50 22.7 23.1 23.6 24.0 24.5 24.9 25.4 60 27.2 27.7 28.1 28.6 29.0 29.5 29.9

70 31.8 32.2 32.7 33.1 33.6 34.0 34.5

80 36.3 36.7 37.2 37.6 38.1 38.6 39.0

90 40.8 41.3 41.7 42.2 42.6 43.1 43.5

100 45 46 46 47 47 48 48

0 10 20 30 40 50 60

200 91 95 100 104 109 113 118

300 136 141 145 150 154 159 163 400 181 186 191 195 200 204 209

500 227 231 236 240 245 249 254

600 272 277 281 286 290 295 299

700 318 322 327 331 336 340 345

800 363 367 372 376 381 386 390

900 408 413 417 422 426 431 435

1000 454 458 463 467 472 476 481

Table A-4. WEIGHT (MASS): OUNCES OR POUNDS TO

(1 oz = 0.028 349 52 kg) ( 1 lb = 0.453 592 4 kg)




lb 0 100 200 300 400 500

(000)* kg kg kg kg kg kg

1 454 499 544 590 635 680

2 907 953 998 1043 1089 1134

3 1361 1406 1451 1497 1542 1588

4 1814 1860 1905 1950 1996 2041

5 2268 2313 2359 2404 2449 2495

6 2722 2767 2812 2858 2903 2948

7 3175 3221 3266 3311 3357 3402

8 3629 3674 3719 3765 3810 3856

9 4082 4128 4173 4218 4264 4309

10 4536 4581 4627 4672 4717 4763

11 4990 5035 5080 5126 5171 5216

12 5443 5488 5534 5579 5625 5670 13 5897 5942 5987 6033 6078 6123

14 6350 6396 6441 6486 6532 6577

15 6804 6849 6895 6940 6985 7031

16 7257 7303 7348 7394 7439 7484

17 7711 7756 7802 7847 7893 7938

18 8165 8210 8255 8301 8346 8391

19 8618 8664 8709 8754 8800 8845

20 9072 9117 9163 9208 9253 9299

Table A-5. WEIGHT (MASS): THOUSAND POUN

(l lb = 0.453 592 4 kg)

*Multiply lb value by 1,000




The Annunciator section presents a c

resentation of all the annunciator lighairplane.

Please unfold page ANN-5 to the right it open for ready reference as the annuare cited in the text.




120

10080

MASTER

OIL

60

40

20

0

N%1000

800

600

400

200

0

T5°C

0

2025

30V –

BAT

GEN

A –100

100200

300

350

0+

_

APU

STOPSTART

GENGEN 1 BAT 1 GEN 3

NORMAL

EXT POWER

APU DC SYSTEM

FUEL SYSTEMENGINES

CMPTR 1AUTO

MAN

OFF

IGN 1

CMPTR 2

AIR START

MOTOR-START STOP

CMPTR 3AUTO

MAN

GRD START

OFF

IGN 2 IGN 3

PRESS TO START

HOLD TO MOTOR

PILOT ST-BY COPILOT ENG 1 ENG 2 ENG 3 WINGPI

IRS 1 HRZN IRS 2 IRS 3

GEN 1E BAT 2

TEST

FLIGHTNORM

A

BAT 1 BAT 2

HP 1

BOOSTERNORM

ST-BY

OFF

ENG 2ENG 1 ENG 3

XTK 2*

LEVEL

°C

BOOSTER

L

AUTO

OPEN

CLOSED

LEVEL

XTK

X.BP

LEVELX.BP X.BP

BOOSTER



CMPTR 1

OIL 1

GEN 1

BAT 1

CMPTR 2

OIL 2

GEN 2

BUS TIED

CMPTR 3

OIL 3

GEN 3

BAT 2

L. PITOT

AOAPROBE

HOTBAT

L AOA

ST BYPITOT

L. WHLOVHT

AUTOSLATS

R. AOA

R. PITOT

R. WHLOVHT

FLAPASYM

BRIGHT TEST

FIRELIGHTS

DIM

ENG 2 FAIL PWRINC

RESET

L AFCS R AFCS

BRT20°c SAT 25°c TAT 175KTAS

L AFCS FAIL

R AFCS MASTER

AP ENGAGEDLH

AVIONICSRH

AVIONICS

DISCH2

1

0

FAULT

FUELSHUTOFF

TRANS

FIRE 1

D2

1

0

F




FUEL 1

LOFUEL 1

BLEEDOVHT

FUEL 2

#2 P BK CABIN

LOFUEL 2

ECUOVHT

FUEL 3

FUELING

LOFUEL 3

COND'GOVHT

AP

AILZERO

NOSECONE OVHT

XTK 2OPEN

XTK 2CLOSED

REARDOORS

MISTRIM

AILFEEL

BLEEDAPU

T/OCONFIG

BAG ISOL

MACHTRIM

BAGACCESS

PITCHFEEL

ITT

°C

0

200

4 0 0 6 0 0

800ITT

°C

0

200

4 0 0 6 0 0

800ITT

°C

0

200

4 0 0 6 0 0

800

IRS

XFR

SGIASM

OFF

HSI ADI

BRT

COMPOS

TGT

GAIN MODEPRESET

GCR RCT

STBYOFF

TEST WXGMAP

RANGE TILT

2510 0

50 100 +15

_15

200300

REVUNLOCK

OFF

MASTER

OFF

MASTER

L RPUMP 1

PUMP 3

PUMP 2

ST/BYPUMP

1/1

1/2

0

4

HYDRASKID

TEST

ON

OFF

OFFOFF

ON

AUTO

3

2

1

0

1/1

1/2

0

4

HYDR

Q TY P SI X 1000Q TY P SI X 1000

3

2

1

0

#1

#1

#2

BRAKEST-BYPUMP

AFT

CAB

SMK

LAV

SMK

HORN SIL

EBATTONERESET

DEPLOYED

TRANSIT

EBAT1

ON OFF

AUTO

MAN

EBAT2

ON OFF

AUTO

MAN

ATC1

ATC2

RESE

L AFCS

IR S

GAINPRES

STO

* *

* ON AIRCRAFT WITHOUT ELECTRONIC TRANSFER VALVE XTK2, THESE LIGHTS ARE NOT FEATURED.

ECUA/I



PILOT CHECK

9NORMAL PROCEDURES

Revision 4

FA

INSERT LATEST REVISED PAGE. DESTROY SUP

LIST OF EFFECTIVE PAGES

Original and changed pages are



Original ....................................0 Incorporates

Revision ...................................1

Revision ...................................2 Incorporates

Revision ..................................3

Revision ..................................4

TOTAL NUMBER OF PAGES IN THIS PUBLI

CONSISTING OF THE FOLLOWIN


Title ............................................4ii..................................................4N-i...............................................3N-ii..............................................4N-1 – N-24 .................................4

*Zero in this column indicates an original

These are suggested training proceduand in no way supersede current prooutlined in the FAA-approved Flight and any revisions thereto. In the conflict the Flight Manual takes preced

Page No.

L-1 – L-4....P-1AR – P-2P-1BR – P-2

NORMAL PROCEDU



PREFLIGHT — EXTERIOR INSPECTION .............

PREFLIGHT — INTERIOR INSPECTION ..............

BEFORE START (POWER OFF)............................

BEFORE START (POWER ON) .............................

AFTER APU START OR ENGINE 2 START OR IFA GROUND POWER UNIT IS USED .................

STARTING ENGINES .............................................

START.....................................................................

BEFORE TAXI.........................................................

TAXI.........................................................................

BEFORE TAKEOFF ................................................

AFTER TAKEOFF ...................................................

CRUISE ...................................................................

DESCENT................................................................

APPROACH.............................................................

BEFORE LANDING.................................................

AFTER LANDING....................................................

PARKING.................................................................

FALCON 900 A-B NORMAL PROCEDURES CHE

INTRODUCTION

The operating procedures of this section have been d



g

regard for the design features and operating characterisFALCON 900 airplane.

They have been approved by the DGAC for guidance inprocedures for safe operation.

Observance of these procedures is not mandatory asuch procedures is not intended to prohibit or discour

use of improved or equivalent alternate procedures experience with the airplane. When alternate procresponsibility for compliance with applicable airworthirests with the operator.

When an electronic system is used to display all procor adapted procedures, the operator remains responfor use and updating of this system.

The checks preceded by (*) are to be performed onc

PREFLIGHT — EXTERIOR INSPE

The exterior inspection is presented as a checklist and just before entering the aircraft for the flight.

If flight is delayed, exterior inspection will have to be

PREFLIGHT — EXTERIOR INS

Forward Fuselage



g

1. Normal Static Ports: (Two) Cover/Condition...

2. Pilot Pitot Probe: (One) Cover/Condition.......

3. Stand-By Pitot Probe: (One) Cover/Condition..

4. Left Angle-Of-Attack Sensor:

(One) Cover/Condition ...................................5. Left Stand-By Static Port: (One) Cover/Condition

6. Cockpit Windows............................................

7. Windshield Wipers .........................................

8. Nose Landing Gear:

• No Hydraulic Leaks, Tire Condition ........

• Shock Absorber Height ...........................

• Torsion Link Pin .......................................

• Antiskid Tachometer Connector...............

• Grounding Wire........................................• Chock.......................................................

9. Taxi Light: Condition ......................................

10. Nose Wheel Well ...........................................

• Maint. Access door ..................................

• Nose Cone Drain .....................................

• Wheel Well Doors ....................................

11 Nose Cone Closed and Latched: (Five)


PREFLIGHT — EXTERIOR INSPEC

15. Co-Pilot Pitot Probe: (One) Cover/Condition ...R



16. Normal Static Ports: (Two) Cover/Condition...R

17. Right Stand-By Static Port(One) Cover/Condition ...................................R

18. Front Toilet Service Door: (If Installed) ...........

19. Belly Anticollision Light ....................................20. Antennas ..........................................................

21. Water Drain Mast (If Installed) .........................

22. Ice Detection Light ...........................................

23. Landing Light....................................................

24. Parking Brake Accumulator Pressure:(At Least 1,000 PSI).........................................

25. Fuel Vent Valve: No leaks ...............................

26. Fuselage Fuel Sump Drain ..............................

27. Emergency Exit ................................................

28. Overwing Emergency Light..............................

Right Wing

29. Right Engine Air Inlet .......................COVER R

30. Center Engine Air Inlet.....................COVER R

31. Ground Emergency Light .................................

32. Leading Edge Condition...................................

33 G it Fill C

39. Right Landing Gear:

• No Hydraulic Leaks, Tire Condition ........

• Shock Absorber Height ...........................



• Brake Wear Indicators.............................

• Fairing Door...........................................C

Rear Fuselage – Rear Cone-Tail Surfaces

1. Nacelle, Ventilation-Drains............................

2. Right Engine Tail Pipe ..................................

3. Engine Pylon Static Discharger ....................

4. Pressure Fueling Access Door .....................

5. Pressure Fueling Control Door .....................

6. Antennas .......................................................

7. Heat Exchanger Air Inlet and Outlet .............

8. Lavatory Service Panel andDrain Access Doors ......................................

9. External Power Connector Access Door ......

10. Hydraulic Coupling No. 2 Access Door ........

Rear Compartment

• Rear Compartment Door .........................

• Batteries ..............................................CO

• Main Electrical Box(Circuit Breakers In, Switches Guarded).

• No. 2 Hyd Accumulator Pressures(Pressure Holding No. 2 – 1,700 PSI,Accumulator No. 2 – 1,500 PSI,Thrust Reverser Accumulator – 1,500 PSI)



E i Oil L l



• Engine Oil Level .........................................• Fire Extinguisher Pressure(s).....................

• No. 1 Hyd Accumulator Pressures(Pressure Holding No. 1 – 1,700 PSI,Accumulator No. 1 – 1,500 PSI) ................

• LH Ground Test Panel Access Doors ........

• S-Duct Access Door...................................

• No. 2 Hyd Bay

Stand-By Pump Motor .........................

Stand-By Pump Selector Handle Set T

“In Flight and Saftied”..........................Hyd 2 Filter Plugs (4) ..........................

Hyd 2 Bay Access Door ......................

Hyd 2 Reservoir Level.........................

Hyd 2 Air Trap..........................CHECK

• No. 1 Hyd Bay

Hyd 1 Air Trap..........................CHECK

Hyd 1 Reservoir Level.........................

Hyd 1 Filter Plugs (5) ..........................

Hyd 1 Bay Access Door ......................• Center Engine Computer ...........................

• Maintenance Test Panel ............................

13. Right Tailplane:

• Leading Edge and Elevator Condition ....

• Static Dischargers (Three) ......................



14. Vertical Stabilizer:

• Leading Edge and Rudder Condition......

• White Navigation Light ............................

• Static Dischargers (Two).........................

• Antennas .................................................15. Center Engine Tail Pipe................................

• Center Engine Static Discharger (1) .......

16. Thrust Reverser Stowed Position .................

• Center Engine Latches (3) ......................

17. Left Tailplane:• Leading Edge and Elevator Condition ....

• Static Dischargers (Three) ......................

18. Left Engine Tail Pipe: Cover/Condition..........

19. Left Engine Pylon Static Discharger

and Nacelle Latches (3).................................

20. APU Generator Ventilation: Inlet/Outlet ........

21. APU Ventilation .............................................

22. Hydraulic No. 1 Coupling Access Door .........

23. Baggage Compartment Door.........................

• Door Control Access Panel......................

• Latch Positions (4) ..................................

24 W t T k Filli P l



Left Wing



Left Wing1. Left Landing Gear:

• No Hydraulic Leaks, Tire Condition ...........

• Shock Absorber Height ..............................

• Brake Wear Indicators................................

2. Flaps/Airbrakes/Aileron ....................................

3. Static Dischargers (Four) .................................

4. Navigation/Strobe Lights – Wing Tip Fairing....

5. Wing Check for Fuel Leaks..............................

6. Wing Fuel Sump/Drain ................................CH7. Gravity Filler Cap .............................................

• Gravity Fueling Bonding.............................

8. Leading Edge Condition...................................

9. Left Engine Air Inlet .........................COVER R

• Pylon Heat Exchanger Air Inlet ..................

Left Forward Fuselage

1. Fuel Vent Valve: No Leaks ..............................

2. Ice Detection Light ...........................................

3. Left Landing Light ............................................

4. Cabin Access Door ..........................................

5 O S t Filli d V t

PREFLIGHT — INTERIOR INS

1 Oxygen: Pressure/Valve



1. Oxygen: Pressure/Valve ................................

2. First Aid Kit.....................................................

3. Axe ................................................................

4. Cabin/Cockpit Fire Extinguishers .................I

5. Emergency Exit Safety Device (If Installed) ..

6. Documents and Keys ....................................

7. Survival Equipment ........................................

8. Smoke Hood(s) .............................................

9. Smoke Goggles (Two) ..................................

10. Circuit Breakers .............................................

11. Park Brake Handle.........................................

12. Headsets and Mikes .....................................

13. Flashlights .....................................................

14. Charts/Plates.................................................


NORMAL PROCEDUR

NOTE

Items marked with an asterisk (*) are to be a



Items marked with an asterisk ( ) are to be aon the first flight of the day only with the sam

BEFORE START (POWER O

Cockpit Check

1. Park Brake................................................INTE

Overhead Panels

1. Circuit Breakers................................................

2. LH AV Master and RH AV Master ....................

3. FMS Master (LH and RH).................................

4. Bus-Tied Switch ...............................................

5. APU Generator Switch .....................................

6. BAT 1-BAT 2 Switches.....................................

7. GEN 1-GEN 2-GEN 3 Switches.......................

*8. IRS 1-IRS 2-IRS 3 Battery Voltages ................

*9. E. BAT Battery Voltage....................................

10. APU Master Switch ..........................................

11. DC Power Selector...........................................

12. CMPTR 1-2-3 Switches ...................................

13 Start Selector S itches (3)

19. APU Bleed Switch .........................................

20. Isolation Valve Knob......................................

21. Crew and Passenger Valve Switches (2) ......



22. BAG Switch....................................................

23. Windshield Switches (3) ................................

24. Anti-lce Switches (4)......................................

25. Pitot Heating Switches (3) .............................

26. Wiper Switches (2) ........................................

27. Exterior and Interior Lights (7).......................

28. Instrument and Panel Lighting Rheostats (6)

Left Console

1. Data Loader...................................................

2. Oxygen Mask ................................................

3. IRS 1 (and Optional IRS 3) ...........................

4. Audio Control Panel.......................................

• SPK, ST Audio VHF 1 and VHF 2...........

• Microphone VHF AND C’PIT ...................

• All Other Push Buttons ............................

• VOR-DME Potentiometer.........................

Pilot Instrument Panels

1. Clock .............................................................

2. EFIS Dim Control (2).....................................


BEFORE START (POWER OFF

4. Normal L/G Control ..........................................

5 Gear Pull Handle



5. Gear Pull Handle..............................................

6. Brake Selector Switch ......................................

7. ST-BY Pump Switch.........................................

8. Temperature Controllers (2) Selector and Knob

9. THRUST REVERSER Switch ..........................

Co-Pilot Instrument Panel

1. Automatic Cabin Pressure Controlle ................

2. DN/UP Knob.....................................................

3. Auto/Manual Selector Switch............................

4. NORM/EMERG Selector Switch.......................

5. DUMP Switch ...................................................

6. Clock ................................................................

7. ELT Switch........................................................

8. GPWS Flaps O’RIDE Switch ...........................

Right Console

1. Passenger Oxygen Valve Selector ..................

2. Oxygen Pressure .............................................

3. VHF 3 (If Installed) ...........................................

4. Audio Control Panel..........................................

• SPK ST Audio VHF 1 and VHF 2

6. IRS 2..............................................................

7. Oxygen Mask ................................................

8. Cond Lever (If Installed)................................

9 N L (If I ll d)



9. Nose Lever (If Installed) ................................

Pedestal

1. Power Levers (3) ...........................................

2. Radios (VHF 1)..............................................

3. Airbrake Handle .............................................

4. Normal Tailplane Circuit Breaker...................

5. Flap-Slat Handle ............................................

6. Emergency Slats Switch................................

7. MMO Switch (If Installed) ..............................

8. Landing Gear Emergency Extension Handles


PILOT NOTES



BEFORE START (POWER

1. Battery Switches (2) ......................................

• BAT 2BAT 1



• ........................

2. Battery Voltages............................................

*3. HRZN Battery Voltages..................................

4. Bus-Tie Switch...............................................

5. Park Brake/No. 2 Park Brake Light ...............6. Fire Detection ................................................

7. COND BATT Pushbutton Light (SB-125)......

8 Navigation Lights ...........................................

9. Lights and Wording “N

A. APU Start

1. Booster 2.......................................................

• Light .....................................

2. APU Master ...................................................

3. APU Start Switch (One Second) ...................

4. APU N1-T5 Indicators–APU Generator Volts

Discontinue start (STOP PUSHBUTTON) i

rise within 10 seconds. Wait 5 minutes prioa second start.

5. APU Bleed Air Switch ....................................

CAUTION

FUEL 2

T/OCONFIG

ENG 2FAIL

BAT 2BAT 1


AFTER APU START OR ENGINE 2 SA GROUND POWER UNIT IS

1. LH AV Master, RH AV Master..........................R T

I N G

S



,

• Lights (LH and RH)................

2. FMS Master (LH and RH) ................................

• Lights (LH and RH)................

3. Maintenance Test Panel ..................................

4 Voice Recorder.................................................

5. IRS (3)/FMS (2) VLF Omega............................

• POS Sensors..............................................

• Data Base Validity......................................

6. Clocks ..............................................................

7. Fasten Belt/No Smoking Lights........................

8. Emergency Lights.............................................

9. Exterior/Interior Lights......................................

10. Standby Horizon ..............................................

11. Master Warning Panel......................................

12. Landing Gear Panel .........................................

13. Fuel Quantity Indicators ...................................

14. Fuel/Gross Weight Counters ...........................

15. Takeoff Data/Bugs ...........................................

16. Cabin Pressure Controller ................................

FMSMASTER

AVMASTER

B E

F O R E

S T A R

E N G I N E S

23. No. 2 Stall Test Pushbutton ..........................

24. Standby Pump Switch ...................................

25. Cabin Pressure Controller .............................

• Cabin Aural Warning



Cabin Aural Warning ...............................

26. Battery Temperature Indicator.......................

27. ADC 1 Then ADC 2 Pushbuttons..................

• VMO/MMO Aural Warnings.....................

28. EFIS Annunciators/GPWS.............................

29. Altimeters and ASEL .....................................

30. Radios Altimeter ............................................

31. Radios/E. BAT 2............................................

32. VHF 1 ............................................................

33. and Lights...............

34. Seats and Rudder Pedals..............................

STARTING ENGINES

Starting Problems

Discontinue a start whenever any one of the follow

• The ITT does not rise within 10 seconds power lever to idle.

• Oil pressure does not rise within 10 seco• N1 remains close to zero when N2 is 20%

• The ITT rises rapidly and approaches the

REARDOORS

CABIN


STARTING ENGINES (Co

Perform a dry motoring whenever fuel is suspected

in the tailpipe:



1. Power Lever.....................................................

2. Start Selector Switch .....................................M

3. Start Button ...............................HOLD DEPRE

If the light remains on although N2 sp50% and all idle parameters are

1. Start Selector Switch .....................................M

START

Engine 2 Start

1. Anticollision Lights Switch................................

2. Booster Switch .................................................

• Light CheckedFUEL 2

IGN

If Ignition Light

Goes Out... Re

2. Start Selector Switch ....GRD START 2. Associated

3. Associated

Circuit Bre

5. Power Lever...................................................

• 2 Light ..................................

6. ITT, N1, Fuel Flow and Oil Pressure.............

When N2 Reaches 50%, 2 Light .........IGN

IGN



With N2 Stabilized:

• Light..................................

• Hydraulic Pressure #2..............................

• Light ........................................

• Light

If Normal Start ........................................

If EXT Power Start ..................................

7. Idle Speed Parameters .................................

8. Power Selector Switch ..................................

9. Ground Power Unit (As Applicable) ..............

• Light ....................................

Complete page N-14 and N-15 items 1 th

Engine 3 and 1 Start

1. Booster Switch (3 or 1) .................................

• Light .....................................

2. GEN 2 Ammeter ............................................

Use same starting procedure as used for Engin

1. Engine Anti-Ice...............................................

BEFORE TAXI

FUEL

GEN 2

GEN 2

OIL 2

PUMP 2


BEFORE TAXI (Cont)

5. Booster 2 Switch...............................................

• Light Checked.........................FUEL 2



g

*6. Isolation Valve Knob.........................................

• Light .........................................

7. Compass Headings (5) ....................................

8. Windshield: Pilot and Co-pilot ..........................

9. Windshield Side................................................

10. Warning Panel Lights (7 or 8) ..........................

11. Hydraulic System 1 and 2 Pressure and Quan

12. Standby Pump Switch .....................................

13. Anti-Skid System ..............................................

*14. Airbrakes ..........................................................

*15. No. 1 Stall System............................................

*16. Airbrakes ..........................................................

17. Flaps–Slats.......................................................

18. No. 1 and No. 2 Stall Systems .........................

19. Mach Trim ........................................................

20. Yaw Damper ....................................................

21. AP Transfer by (LH FCS and RH FCSPush Buttons on ID-802)..................................

22. COMM/NAV/RADAR/IRS ................................

ISOL

TAXI

1. Taxi Light.......................................................

2. Parking Brake Handle....................................3 No 1 and No 2 Brakes



3. No. 1 and No. 2 Brakes ................................

4. Brake Selector...............................................

5. Thrust Reverser.............................................

6. Engine Computers ........................................

7. Pilot Window .................................................

8. APU STOP Button .........................................

9. APU Master (N1 Zero)...................................

10. APU Bleed .....................................................

*11. Engine Anti-Ice/Wing (or Wing-BRK) Anti-ice(5 SECO

12. Engine Anti-lce ................................AS REQ13. Flight Controls ...............................................

14. Takeoff Briefing..............................................

BEFORE TAKEOFF

1. Radar (2)/Transponder..................................

2. Parking Brake Handle ...................................

3. Flaps–Airbrake–Trims–Speeds (FATS) ........

4. Anticollision Lights.........................................

5. Landing Lights ...............................................

6. Start Selector Switches (3) (If Necessary).....


AFTER TAKEOFF

1. Landing Gear ...................................................

2. Wing Anti-Ice Switch........................................



3. Flaps–Slats ..................................CLEAN/HYD

4. Start Selector Switches (3) ..............................

5. Climb Power ....................................................

6. Taxi Light..........................................................7. Fasten Belt/No Smoking Switches...................

8. Cabin Pressure and Temperature....................

9. Entrance Curtain ..............................................

After 10,000 Feet1. Landing Lights ..................................................

After 18,000 Feet or Transition Level

1. Altimeters.........................................................

2. Oxygen.............................................................

3. Station Check:

a. Circuit Breakers..........................................

b. Electrical Panel...........................................

c. Engine Instruments ....................................

d. Hydraulic Panel ..........................................

e. Fuel Panel and Quantities..........................

CRUISE

1. Fuel Management..........................................

2. Station Check ................................................



DESCENT

1. Cabin Pressure Controls ...............................

2. Fasten Belt ....................................................

3. Anti-lce Systems ............................................

4. Altimeters ....................................................S

5. Landing Computations/Bugs..........................

6. Approach Briefing ..........................................

Below 10 000 Feet

WT/1,000 LB 24 26 28 30 32 34 36

VREF 100 104 108 112 115 119 122

HOLDING 172 179 187 193 201 207 212

VREF Correction for Configuration (New Bug)Clean Wing +30 KT Outboard Slats

Slats Only +20 KT Outboard Slats

Slats + 7° Flaps +15 KT Outboard Slats

Slats + 20° Flaps + 5 KT Outboard Slats

All wind Gust + 1/2 Steady Wind (Max 20 KT)


APPROACH

1. Entrance Curtain...............................................

2 No Smoking Sign



2. No Smoking Sign .............................................

3. Altimeters/Radio Altimeters.........................SE

4. Fuel Crossfeeds (3) .........................................

5. Flaps–Slats ...................................+ 20 DEGRE

BEFORE LANDING

1. Landing Gear................................................DO

2. Anti-Skid ...........................................................

3. Hydraulics ........................................................

4. Flaps-Slats....................................+ 40 DEGRE

5. Start Selector Switches (If Necessary) ............

6. Autopilot............................................................

AFTER LANDING

1. Thrust Reverser ...........................................S

2 Anti-lce: Wing (or Wing-BRK)



2. Anti-lce: Wing (or Wing-BRK)........................

3. Pitot Heat Switches (3)..................................

4. Start Selector Switches (3)............................

5. Anticollision Lights..........................................6. Landing Lights ...............................................

7. Taxi Light.......................................................

8. Radar (2)/Transponder ..................................

9. Flap + Slats Handle .......................................

10. Airbrake Handle .............................................

11. Windshield Heat Switches (3) .......................

12. Trims (3).........................................................

13. Bus-Tied Switch.............................................

14. APU................................................................

NOTE

As a minimum, the No. 2 and No. 3 engin

kept running until parked at the ramp, pr

minute conditioning period at idle power is


PARKING

1. Parking Brake/No. 2 Brake Light.............INTE



2. Flight Data (Fuel-IRS) ......................................

3. AVIONICS/FMS/EFIS Masters .........................

4. Engine Anti-Ice Switches (3) ............................

5. Taxi Light ..........................................................

6. IRS (2 or 3) ......................................................

7. Radar (2)/Transponder .....................................

8. Standby Horizon ..............................................

9. Standby Pump..................................................

10. VHF 1/No. 2 Emergency Battery ......................

11. Engines (3): (After 2 Minutes At Idle Speed)...

12. Booster Pump Switches (3) With APU (2) .......

13. Anticollision/Navigation Lights ..........................

14. Fasten Belt/No Smoking/Emergency Lights.....

15. lnterior Lights ....................................................

16. APU/No. 2 Booster Pump................................

17. Batteries (2) ......................................................

18. Engine Computers (Last Flight) .......................19. Generators Switches (Last Flight) ...................

20. Aircraft Chocked ..............................................

LIMITATIONS

Weight

Ramp:

5AR 45 70



–5AR.............................................................45,70

–5BR .....................................................................

Taxi/Takeoff

–5AR.............................................................45,50

–5BR .....................................................................

Landing....................................................................

Zero Fuel Weight

–5AR.............................................................28,20

–5BR .....................................................................

NOTE

Zero fuel weight may change based on S

Load Factors

Clean .......................................................................

Flaps Extended .......................................................

Do not extend or maintain the slats or flap

20,000 feet.

CAUTION


LIMITATIONS (Cont)

Speeds

VMO:

Sea Level to 10 000 Feet



Sea Level to 10,000 Feet......................................

10,000 to 25,000 Feet ...........................................

MMO:

Above 35,000 Pounds:

25,000 to 33,000 Feet......................................

33,000 to 37,000 Feet .....................................

Above 37,000 Feet...........................................

Below 35,000 Pounds:

25,000 to 37,000 Feet......................................

37,000 to 42,000Feet.......................................

Above 42,000 Feet...........................................

VA ...............................................................................

VLO /MLO.......................................................................

VLE /MLE .......................................................................

VFE Slats + 7° Flaps...................................................

VFE Slats + 20° Flaps.................................................

VFE Slats + 40° Flaps.................................................

VWWE (Wipers) ............................................................

Pilot Window ..............................................................

Tire Groundspeed ......................................................

VMCA:

5AR

L I M I T A T I O N

S

Altitudes

Airbrakes ..............................................NO LOWER

Maximum Operating Altitude...................................

Takeoff Landing Altitude .......................................–

Autopilot:



Autopilot:

Minimum Height during FMS Approach .............

Minimum Height Radio Altimeter Operative .......

Minimum Height Radio Altimeter Inoperative .....

Minimum Decision Height...................................

Minimum Height Except during Approach ..........

Engine Operation

TFE 731-5AR Engines

N1 ..........................................100 TO 103% (FIVE

N2 ..........................................101 TO 103% (FIVE

ITT:

Start.....................................................................

Takeoff ................................................................

Maximum Continuous .........................................

Generator Load:

To 43,000 Feet ...................................................

Above 43,000 Feet .............................................

One Minute Transient .........................................

Oil Pressure:

Idle ......................................................................

T k ff/Cli b/C i


LIMITATIONS (Cont)

TFE 731-5BR Engines

N1 ..........................................100 TO 103% (FIVE S

N2 101 TO 103% (FIVE S



N2 ..........................................101 TO 103% (FIVE S

ITT:

Start.......................................................................

Takeoff ..................................................................

Maximum Continuous ...........................................

Generator Load:

To 43,000 Feet ......................................................

Above 43,000 Feet ................................................

One Minute Transient ............................................Oil Pressure:

Idle .........................................................................

Takeoff/Climb/Cruise ............................................

Transient....................................55 PSI (THREE M

Oil Temperature:

To 30,000 Feet ......................................................

Above 30,000 Feet ................................................

Transient.........................................149°C (TWO M

PERFORMANCE

TAKEOFF SPEEDS AND BFL—SLATS +20° FLAP



TAKEOFF SPEEDS AND BFL—SLATS +7° FLAPS

TAKEOFF/MAXIMUM CONTINUOUS POWER

WITHOUT ANTI-ICE—ALTITUDE 0–4,000 FEET .

TAKEOFF/MAXIMUM CONTINUOUS POWERWITHOUT ANTI-ICE—ALTITUDE 5,000–9,000 FE


WITH ANTI-ICE—ALTITUDE 0–4,000 FEET.........


WITH ANTI-ICE—ALTITUDE 5,000–9,000 FEET..

MAXIMUM CLIMB THRUST: ALL ENGINES

OPERATING CLIMB—0–25,000 FEET...................

MAXIMUM CLIMB THRUST: ALL ENGINES

OPERATING CLIMB—25,000–50,000 FEET.........

MAXIMUM WEIGHT IN LEVEL FLIGHT

LIMITED BY MAXIMUM CRUISE THRUST ...........

CRUISE MACH—0.75 ............................................

CRUISE – MACH 0.80 ............................................

MAXIMUM CRUISE THRUST ................................

LANDING DATA ......................................................

P -2 A R

FALCON 900 DRYSLATS + FLAPS 20° NO

TAKEOFF DATA NO

PRESSURE ALTITUDE AND T

TAKEOFF SPS +

G.W. D

INTERPOLATION ONLY

731-5AR



R ev i s i on

3

PRESSURE ALTITUDE AND TGX

1000

24

26

28

30

32

34

36

38

40

42

44

45.5

ATA

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

SEA LEVEL

0° 10° 20° 30° 40°

88 88 88 88 882600 2600 2750 3000 3300

88 88 88 88 882600 2600 2750 3000 3300

88 88 88 88 882600 2600 2750 3000 3300

88 88 88 89 922600 2600 2750 3000 3300

91 91 91 94 962700 2750 2900 3250 3700

95 95 95 98 1013000 3050 3200 3550 4150

100 100 100 102 1063300 3350 3500 4000 4650

104 104 104 107 1103650 3700 3800 4400 5150

108 108 108 111 1144000 4050 4200 4850 5700

112 112 112 115 1184300 4400 4600 5400 6300

116 116 116 1194750 4800 5050 6000

118 118 118 1215100 5150 5400 6400

0° 10°

88 88 2700 2800

88 88 2700 2800

88 88 2700 2800

88 88 2700 2800

91 91 2850 2900

95 95 3100 3200

100 100 3450 3550

104 104 3850 3900

108 108 4200 4250

112 112 4600 4650

116 116 5000 5050

118 118 5400 5500

R ev i s i on

3

FALCON 900 DRYSLATS + FLAPS 20° NO W

TAKEOFF DATA NO S

TAKEOFF SPES +INTERPOLATION ONLY



P - 3 A R

PRESSURE ALTITUDE AND TEG.W.X

1000

24

26

28

30

32

34

36

38

40

42

44

45.5

DATA

V1

BFLV1

BFL

V1BFL

V1BFL

V1

BFLV1

BFL

V1BFL

V1BFL

V1

BFLV1

BFL

V1BFL

V1BFL

4000’

0° 10° 20° 30° 40°

88 88 88 88 88

2800 3000 3300 3550 385088 88 88 88 882800 3000 3300 3550 3900

88 88 88 88 912800 3000 3300 3550 4000

88 88 90 93 962800 3000 3300 3750 4450

91 92 94 97 101

2950 3300 3700 4200 500095 96 99 102 105

3300 3550 4100 4650 5550

101 101 103 106 1103700 3950 4600 5200 6250

104 105 108 110 1144050 4400 5150 5800 7000

108 109 112 115

4500 4900 5650 6450112 113 116 1194900 5400 6300 7150

116 117 1205400 5900 6900

118 119 1235800 6300 7400

0° 10°

88 88

3050 3300 88 88

3050 3300

88 88 3050 3300

88 89 3050 3300

92 94

3350 3700 96 99

3700 4050

101 103 4100 4600 5

105 107 4550 5100 5

109 111

5000 5600 6113 115 5500 6200 7

117 119 6050 6850 7

119 1226450 7400

P -4 A R

FALCON 900 DRYSLATS + FLAPS 7° NO

TAKEOFF DATA NO

TAKEOFF SPS

INTERPOLATION ONLY



R ev i s i on

3

TAKEOFF DATA NO

PRESSURE ALTITUDE AND TG.W.X

1000

26

28

30

32

34

36

38

40

42

44

45.5

DATA

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

4000’

0° 10° 20° 30° 40°

88 88 88 89 902750 2800 3050 3300 3750

88 89 91 93 952750 2900 3250 3700 4250

93 94 95 97 993000 3250 3600 4100 4800

97 98 99 101 1033300 3600 4050 4600 5400

101 102 103 105 1073650 3950 4450 5100 6050

105 106 107 109 1124000 4400 5000 5700 6800

109 110 111 113 1164450 4900 5550 6400 7600

112 114 115 117 1204900 5400 6200 7200 8500

116 118 119 1225400 6000 6850 8000

120 122 123 1265950 6600 7600 8900

123 124 126 1296400 7050 8200 9600

0° 10°

88 88 2850 3050

89 91 3000 3250

93 95 3350 3600

98 99 3700 4050

102 103 4100 4500

106 107 4600 5000

110 111 5100 5600

114 115 5600 6200

117 119 6200 6900

121 123 6800 7600

124 125 7250 8200

R ev i s i on

3

FALCON 900 DRSLATS + FLAPS 7° NO

TAKEOFF SPS

INTERPOLATION ONLY



P - 5 A R

SLATS + FLAPS 7 NOTAKEOFF DATA NO

PRESSURE ALTITUDE AND TG.W.X

1000

26

28

30

32

34

36

38

40

42

44

45.5

DATA

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

8000’

0° 10° 20° 30° 40°

88 88 89 913100 3300 3700 4200

91 92 93 953350 3700 4200 4800

95 96 98 1003750 4100 4750 5500

99 100 102 1044200 4650 5350 6200

103 104 106 1084650 5200 6000 7000

107 108 110 1125150 5800 6700 7850

111 112 114 1165750 6500 7500 8750

115 116 118 1216400 7250 8400 9800

119 120 1227100 8000 9350

123 124 1267800 9000 10400

126 1288400 9700

0° 10°

88 893400 3750

92 933800 4250

96 97 4300 4800

100 1024800 5400

104 1055350 6050

108 1106000 6800

112 1146750 7600

116 1187500 8500

120 1228400 9500

124 1269400 10600

128 12910150 11400


TAKEOFF/MAXIMUM CONTINUOU

WITHOUT ANTI-ICE

NOTE: *SAT IS USED FOR T.O.; TAT IS U

T.O. MMC M



TEMP °C LAT CENT LAT CENT LAT CENT L

+40 TO 93.2 91.5 93.1 91.4 93.1 91.5 9

MC 90.3 88.1 90.1 87.9 89.8 87.6 8

+35 TO 94.5 93.0 94.5 93.0 94.4 93.0 9

MC 91.6 90.0 91.4 89.8 91.2 89.5 9

+30 TO 95.7 94.4 95.7 94.4 95.7 94.4 9

MC 92.8 91.5 92.6 91.3 92.4 91.0 9

+25 TO 96.8 95.8 96.8 95.8 96.8 95.7 9

MC 94.0 92.9 93.8 92.7 93.6 92.5 9

+20 TO 97.4 96.6 97.8 96.9 97.7 96.9 9

MC 95.1 94.3 95.0 94.1 94.8 93.9 9

+15 TO 96.4 95.8 97.9 97.4 98.5 97.9 9

MC 95.1 94.3 95.4 94.7 95.7 95.1 9

+10 TO 95.6 95.0 97.1 96.6 98.5 98.1 9

MC 94.2 93.4 95.4 94.8 96.6 96.2 9

+5 TO 94.7 94.2 96.2 95.8 97.6 97.3 9

MC 93.4 92.7 94.6 94.1 95.8 95.4 9

0 TO 94.0 93.4 95.4 94.9 96.7 96.4 9

MC 92.5 91.8 93.7 93.2 94.8. 94.5 9

–5 TO 93.2 92.5 94.5 94.0 95.9 95.6 9

MC 91.7 91.0 92.9 92.4 94.1 93.7 9

–10 TO 92.2 91.7 93.6 93.3 95.0 94.6 9

MC 90.9 90.2 92.1 91.5 93.2 92.8 9

ALT 0 1000 2000

TAKEOFF/MAXIMUM CONTINUO

WITHOUT ANTI-ICE

NOTE: *SAT IS USED FOR T.O.; TAT IS

T.O.MC



TEMP °C LAT CENT LAT CENT LAT CENT

+40 TO 93.0 91.3 92.9 91.2

MC 89.1 86.6 88.8 86.3 88.6 87.4

+35 TO 94.3 92.9 94.3 92.8 94.2 92.7

MC 90.6 88.8 90.4 88.5 90.3 88.3

+30 TO 95.6 94.3 95.6 94.2 95.5 94.2

MC 91.9 90.4 91.7 90.2 91.6 90.0

+25 TO 96.8 95.6 96.7 95.6 96.7 95.6

MC 93.1 91.9 92.9 91.7 92.8 91.6

+20 TO 97.7 96.8 97.7 96.6 97.7 96.7

MC 94.2 93.3 94.0 93.1 93.9 93.0

+15 TO 98.5 97.8 98.5 97.8 98.4 97.8

MC 95.3 94.6 95.1 94.4 95.0 94.3

+10 TO 99.1 98.7 99.1 98.7 99.1 98.6

MC 96.2 95.8 96.0 95.6 95.9 95.4

+5 TO 99.9 99.6 99.8 99.6 99.8 99.5

MC 96.9 96.7 96.7 96.5 96.6 96.4

0 TO 100.0 100.0 100.0 100.0 100.0 100.0

MC 97.3 97.4 97.3 97.4 97.2 97.3

–5 TO 99.7 99.9 100.0 100.0 100.0 100.0

MC 97.3 97.4 97.9 98.2 97.7 98.1

–10 TO 98.8 98.9 99.9 100.0 100.0 100.0

MC 96.8 97.0 97.9 98.2 97.9 98.4

ALT 5000 6000 7000



WITH ANTI-ICE




T.O. M

MC M


+10 TO 95.0 94.3 96.6 95.8 97.4 96.7 9

MC 92.8 88.7 93.0 88.5 93.2 88.3 9

+5 TO 94.2 93.4 95.6 95.0 97.0 96.6 9

MC 92.1 90.1 93.1 89.9 94.2 89.7 90

TO 93.4 92.6 94.8 94.2 96.1 96.7 9

MC 91.3 89.8 92.5 90.4 93.8 91.1 9

–5 TO 92.5 91.7 94.0 93.4 95.3 94.9 9

MC 90.5 89.6 91.7 90.8 93.0 92.0 9

–10 TO 91.6 90.8 93.1 92.4 94.5 94.0 9

MC 89.6 88.2 90.8 89.7 92.1 91.2 9–15

TO 90.8 90.1 92.3 91.7 93.6 93.2 9

MC 88.9 87.5 90.1 88.9 91.3 90.4 9

–20 TO 89.9 89.3 91.4 90.8 92.7 92.3 9

MC 87.9 86.6 89.1 88.0 90.4 89.5 9

–25 TO 89.1 88.4 90.5 90.0 91.8 91.4 9

MC 87.1 85.7 88.3 87.1 89.5 88.6 9–30

TO 88.2 87.4 89.5 88.9 90.9 90.5 9

MC 86.2 84.9 87.4 86.3 88.6 87.7 8

ALT 0 1000 2000


WITH ANTI-ICE




T.O

MC


+10 TO 97.3 96.6 97.3 96.5 97.2 96.5

MC 92.8 87.9 92.6 87.7 92.5 97.6

+5 TO 97.9 97.5 97.9 97.4 97.9 97.4

MC 93.8 89.3 93.6 89.1 93.5 89.0

0 TO 98.6 98.2 98.5 98.2 98.5 98.2

MC 94.6 90.6 94.4 90.4 94.3 90.3

–5 TO 98.9 98.6 99.1 99.0 99.1 99.0

MC 95.1 91.9 95.0 91.7 94.9 91.6

–10 TO 98.0 97.9 99.0 99.1 99.6 99.7

MC 95.1 93.0 95.6 92.8 95.4 92.6

–15 TO 97.0 97.0 98.1 98.2 99.2 99.4

MC 94.8 93.5 95.8 93.6 95.8 93.5

–20 TO 96.1 96.1 97.1 97.3 98.2 98.5

MC 93.9 92.9 94.9 94.3 95.6 94.2

–25 TO 95.3 95.2 96.2 96.3 97.2 97.5

MC 93.0 92.7 94.0 93.8 94.7 94.3

–30 TO 94.3 94.2 95.3 95.4 96.2 96.5

MC 92.0 91.8 93.0 92.9 94.0 93.9

ALT 5000 6000 7000


–30°/–20° –20°/–10° –10°/

ANTI-ICE – MINIMUM N1 – 3 ENGINES O

TAT°C



>20000 80% 76% 73%

<20000 76% 73% 65%>10000

<10000 68% 65% 61%

NOTE

WITH ONE ENGINE INOPERATIVE I

N1 BY 9% IF N1>65% OR 6% IF N1<65%

When the brake heating system is used, t

required N1 speed with two or all engine

must be increased by 1%.

CAUTION

ALT

PILOT NOTES




MAXIMUM CLIMB THRUS

ALL ENGINES OPERATING C

CLIMB 260 kt/0.72

Pressure Altitude (ft.)

0 5 000 10 000 15 000



50 86.9 84.2 88.0 85.4

45 88.7 86.3 89.7 87.5

40 90.4 88.4 91.0 89.4 89.7 87.4

35 91.7 90.1 92.1 90.9 90.7 89.3 91.0 89.7

30 93.0 91.7 93.2 92.3 92.1 90.8 92.0 91.1

25 94.1 93.1 94.2 93.5 93.1 92.1 92.9 92.4

20 95.3 94.4 95.1 94.8 93.9 93.4 94.0 93.6

15 95.5 94.8 95.9 95.9 94.9 94.5 94.9 94.9

10 94.6 93.9 96.8 96.9 95.7 95.7 95.7 96.0

5 93.8 92.9 97.4 97.8 96.5 96.6 96.4 96.8

0 93.0 92.3 96.9 97.4 97.0 97.5 97.0 97.6

–5 92.1 91.4 96.0 96.3 97.6 98.3 97.6 98.4

–10 91.3 90.6 95.1 95.5 97.1 97.9 98.2 99.2

–15 90.4 89.8 94.2 94.6 96.2 97.0 98.8 99.7

–20 89.6 88.9 93.4 93.8 95.2 96.0 99.4 100.0

–25 88.7 88.0 92.5 92.8 94.3 95.0 100.0 100.0

–30 87.8 87.1 91.6 91.9 93.3 94.1 100.0 100.0

–35 87.0 86.3 90.6 91.0 92.4 93.1 99.2 100.0

–40–45

–50

0 5,000 10,000 15,000

TA

T

(°C)




MAXIMUM WEIGHT IN LEVEL

LIMITED BY MAXIMUM CRUISE

TEMP

MILR 0.75 0.8 0.84

TEMP

MILR



Pressure altitude: 35,000 FT

–15°C N.L. N.L. N.L. N.L.

–10°C N.L. N.L. N.L. N.L.

–5°C N.L. N.L. N.L. 43,100

0°C N.L. N.L. N.L. 41,100

5°C N.L. N.L. N.L. 37,700

10°C N.L. N.L. N.L. 33,500

15°C N.L. N.L. 42,150 26,550

20°C 43,850 42,650 36,050


–15°C N.L. N.L. 44,400 38,050

–10°C N.L. N.L. 43,400 37,050

–5°C N.L. N.L. 42,300 35,700

0°C 44,700 44,200 40,950 33,690

5°C 42,950 42,500 39,450 31,100

10°C 41,300 40,700 37,800 28,000

15°C 38,950 38,250 34,600 26,450

20°C 36,000 35,250 29,900


–15°C 38,600 38,850 36,050 30,750

–10°C 38,050 37,950 35,250 29,750

5°C 37 150 37 050 34 350 28 550

Pressure altitude

–15°C N.L.

–10°C N.L.

–5°C N.L.

0°C N.L.

5°C N.L.

10°C N.L.

15°C 42,70

20°C 40,40

Pressure altitude

–15°C 43,65

–10°C 42,80

–5°C 41,65

0°C 39,95

5°C 38,35

10°C 36,65

15°C 34,20

20°C 31,40

Pressure altitude

–15°C 34,60

–10°C 34,20

5°C 33 25

PILOT NOTES



P -1 6 A R

CRUISE—

WT

X1000

44

ALT

FF

N1

TAS

–10

721

91.3

420

ISA

744

93.4

430

+10

767

95.5

440

9

350



R ev i s i on

3

42

40

38

36

34

32

30

28

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

FF

N1

TASFF

N1

TAS

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

420

698

90.4

420

67989.7

420

663

89.1

420

648

88.5

420635

88.0

420

623

87.5

420

61287.0

420

602

86.6

420

430

721

92.5

430

70191.8

430

683

91.2

430

668

90.6

430655

90.0

430

642

89.5

430

63089.1

430

619

88.6

430

440

742

94.6

440

72293.9

440

703

93.2

440

688

92.6

440673

92.1

440

660

91.5

440

64891.0

440

637

90.6

440

9

9

9

8

8

8

8

8

R ev i s i on

3

WT

X1000

44

ALT

FF

N1

TAS

–10 ISA +10 –

410

CRUISE—



P -1 7 A R

42

40

38

36

34

32

30

28

TAS

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

FF

N1TAS

FF

N1

TAS

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

698

99.3

418

66097.0

418

626

95.1

418

595

93.5418

566

92.0

418

543

90.8

418

52489.8

418

508

88.9

418

68399.3

428

647

97.4

428

614

95.7428

585

94.2

428

560

92.9

428

54091.9

428

524

91.0

428

634

97.8438

603

96.3

438

577

95.0

438

55694.0

438

539

93.1

438

6

9

4

5

94

5

9

4

5

9

4

59

4

4

9

4

P -1 8 A R

CRUISE—

WTX

1000

44

ALT

FF

N1

TAS

–10

739

93.4

446

ISA

820

95.6

457

+10

846

97.7

467

350



R ev i s i on

3

42

40

38

36

34

32

30

28

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

FF

N1

TASFF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

772

92.7

446

75392.0

446

736

91.3

446

720

90.8

446705

90.3

446

692

89.8

446

680

89.3

446

669

88.9

446

798

94.4

457

77894.2

457

760

93.5

457

743

92.9

457728

92.4

457

713

91.9

457

701

91.4

457

690

91.0

457

823

97.0

467

80296.3

467

783

95.6

467

766

95.0

467750

94.4

467

735

93.9

467

721

93.4

467

710

93.0

467

R ev i s i on

3

WT

X1000

44

ALT

FF

N1

TAS

–10 ISA +10 –

410

CRUISE—



P -1 9 A R

42

40

38

36

34

32

30

28

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

740100.0

444

697

98.4

444

657

96.2

444

622

94.4

444

596

93.2

444

57792.3

444

560

91.4

444

680

98.5

454

643

96.7

454

616

95.4

454

59694.5

454

579

93.6

454

664

98.9

465

636

97.6

465

61596.6

465

597

95.7

465

6

1

4

6

9

4

5

9

4

59

4

5

9

4

P -2 0 A R

MAXIMUM CR

WTX

1000

44

ALT

FF

N1

TAS

FF

–10981

99.3

467

984

ISA946

99.4

472

950

+10873

98.5

471

877

9

350



R ev i s i on

3

42

40

38

36

34

32

30

28

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

FF

N1

TASFF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

984

99.2

470

98699.2

472

988

99.2

474

991

99.2

476993

99.2

478

994

99.1

480

994

99.1

480

984

98.8

480

950

99.4

476

95299.4

478

955

99.3

480

957

99.3

482959

99.3

484

961

99.3

486

963

99.3

488

965

99.3

489

877

98.5

476

88098.5

479

883

98.4

482

885

98.4

485886

98.4

487

888

98.4

489

889

98.4

490

891

98.4

492

9

9

9

9

9

9

9

9

R ev i s i on

3

WT

X1000

44

ALT

FF

N1

TAS

–10 ISA +10 –

410

MAXIMUM CR



P -2 1 A R

42

40

38

36

34

32

30

28

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

701

99.5

418

71599.6

436

723

99.6

448

729

99.6

455

735

99.5

461

739

99.5

465

72798.8

467

702

97.5

467

67399.4

422

689

99.2

444

698

99.3

457

705

99.3

465

709

99.3

470

71299.2

474

714

99.2

477

643

98.2

441

655

98.5

461

662

98.6

471

66598.6

475

668

97.5

480

6

9

4

6

9

4

6

9

4

6

9

4

69

4

6

9

4

P -2 2 A R

LANDIN

FALC

LANDING DATA SLATS + 40 FLAPS – STG.W. LANDING DISTANCE/LAN

X VREF

1000 SEA LEVEL 2000’ 4000’



R ev i s i on

3

1000 SEA LEVEL 2000 4000

24 100 2150/3500 2250/3750 2350/3900 2

26 104 2250/3750 2350/3900 2450/4050 2

28 108 2350/3900 2450/4050 2600/4300 2

30 112 2500/4100 2600/4300 2700/4500 2

32 115 2600/4300 2750/4600 2850/4750 3

34 119 2800/4700 2900/4800 3050/5100 3

36 122 2950/4900 3100/5200 3200/5350 3

38 126 3150/5250 3300/5500 3400/5700 3

40 129 3350/5600 3500/5850 3650/6100 3

42 132 3550/5950 3700/6200 3850/6450 4

PERFORMANCE

TAKEOFF SPEEDS AND BFL—SLATS +20° FLAP

TAKEOFF SPEEDS AND BFL—SLATS +7° FLAPS





WITHOUT ANTI-ICE—ALTITUDE 0–4,000 FEET .

TAKEOFF/MAXIMUM CONTINUOUS POWERWITHOUT ANTI-ICE—ALTITUDE 5,000–9,000 FE


WITH ANTI-ICE—ALTITUDE 0–4,000 FEET.........


WITH ANTI-ICE—ALTITUDE 5,000–9,000 FEET..

MAXIMUM CLIMB THRUST: ALL ENGINES OPER

CLIMB 260 KNOTS/MACH 0.72—0–25,000 FEET.


CLIMB 260 KNOTS/MACH 0.72—25,000–50,000 F


CLIMB 300 KNOTS/MACH 0.80—0–25,000 FEET.


CLIMB 300 KNOTS/MACH 0.80—25,000–50,000 F

MAXIMUM WEIGHT IN LEVEL FLIGHTLIMITED BY MAXIMUM CRUISE THRUST ...........

CRUISE MACH—0 75

P -2 B R

FALCON 900 DR

SLATS + FLAPS 20° NOTAKEOFF DATA NO


TAKEOFF SPS +SECOND SEGMENT LIMITED

G.W.X

1000

DATA

SEA LEVEL

0° 10° 20° 30° 40° 0° 10°

731-5BR



R ev i s i on

3

26

28

30

32

34

36

38

40

42

44

46

46.5

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

90 90 90 90 902600 2700 2800 2900 3300

90 90 90 90 902600 2700 2800 2900 3300

90 90 90 90 902600 2700 2800 2900 3300

90 90 90 92 942600 2700 2800 2900 3400

94 94 94 96 992800 2900 3000 3200 3800

98 98 98 100 1033100 3200 3300 3500 4250102 102 102 104 106

3500 3450 3650 3850 4700

106 106 106 107 1093800 3800 3950 4250 5200

109 109 109 110 1144150 4150 4350 4650 5800

113 113 113 114 117

4450 4500 4750 5100 6350

117 117 117 118 1214800 5000 5150 5700 6950

118 118 118 119 1225000 5200 5300 5900 7100

90 90 2800 2900

90 90 2800 2900

90 90 2800 2900

90 90 2800 2900

94 94 3000 3100

98 98 3300 3400102 1023650 3750

106 1064000 4150

109 1094350 4550

113 113

4750 4950

117 117 5100 5300

118 118 5300 5450

R ev i s i on

3

FALCON 900 DRYSLATS + FLAPS 20° NOTAKEOFF DATA NO


TAKEOFF SPS +SECOND SEGMENT LIMIT

G.W.X

1000

DAT 4000



P - 3 B R

26

28

30

32

34

36

38

40

42

44

46

46.5

A

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

0° 10° 20° 30° 40°

90 90 90 90 902900 3000 3200 3500 3700

90 90 90 90 902900 3000 3200 3500 3750

90 90 90 91 942900 3000 3200 3500 3950

90 90 93 95 982900 3000 3400 3800 4450

94 95 97 99 1023250 3400 3750 4250 4950

98 99 101 102 1063600 3750 4250 4800 5550

102 103 105 107 1103950 4150 4700 5300 6200

106 107 109 110 1144350 4550 5250 5950 6950

109 110 112 1144750 5000 5800 6500

113 114 115 1185200 5450 6300 7250

117 117 119 1225600 6050 6800 7900

118 118 120 1235750 6200 7100 8100

0° 10°

90 90 3200 3350

90 90 3200 3350

90 90 3200 3350

92 93 3250 3400

96 97 3600 3800

100 101 4000 4300

104 105 4400 4750

107 108 4900 5250

111 112 5350 5800

114 116 5900 6350

118 119 6500 7200

119 120 6700 7400

P -4 B R

FALCON 900 DR

SLATS + FLAPS 7° NOTAKEOFF DATA NO


TAKEOFF SPS SECOND SEGMENT LIMITED

G.W.X

1000

DATA

4000

0° 10° 20° 30° 40° 0° 10°



R ev i s i on

3

26

28

30

32

34

36

38

40

42

44

46

46.5

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

90 90 90 90 902800 2900 3100 3300 3650

90 90 91 92 932800 2900 3100 3350 3900

93 93 94 96 982900 3100 3300 3700 4400

97 97 98 99 1023200 3400 3800 4200 4900

101 101 102 104 1063500 3700 4100 4650 5500

104 104 106 108 1103800 4200 4600 5200 6250

108 109 110 112 1144300 4600 5100 5800 6850

112 113 114 116 1184700 5100 5700 6450 7750

116 117 118 119 1225200 5500 6250 7200 8650

119 120 122 123 1275700 6100 6800 7900 9550

124 124 125 1276100 6600 7600 8700

124 125 126 1286300 6700 7700 8850

90 903050 3250

90 913050 3250

94 943200 3500

98 98 3550 3850

101 1023900 4300

106 1064350 4800

109 1104850 5350

113 1135300 5950

117 1185850 6600

121 1216400 7300

124 1257000 7950

125 1267150 8150

R ev i s i on

3

FALCON 900 DRSLATS + FLAPS 7° NOTAKEOFF DATA NO


TAKEOFF SPS SECOND SEGMENT LIMITED

G.W.X

1000

DATA

8000

0° 10° 20° 30° 40° 0° 10°



P - 5 B R

26

28

30

32

34

36

38

40

42

44

46

46.5

A

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

V1BFL

0 10 20 30 40

90 90 90 903150 3300 3550 3900

91 91 93 933200 3500 3800 4400

94 95 97 983650 3900 4300 4950

98 99 101 1024000 4400 4850 5600

102 103 104 1064450 4850 5400 6250

106 107 109 1104950 5500 6100 7100

110 111 113 1145500 6150 6800 7900

114 115 117 1186200 6800 7600 8800

118 119 121 1226800 7600 8450 9850

121 122 1257500 8350 9350

125 126 1298250 9200 10300

126 127 1308400 9400 10600

0 10

90 90 3550 3650

92 92 3700 3900

95 96 4200 4350

99 1004700 4900

103 104 5150 5500

108 108 5850 6200

112 112 6600 6900

115 116 7350 7700

119 120 8200 8650

123 124 9000 9500

127 12810000 10600

128 12910200 10850



WITHOUT ANTI-ICE



ALT 0 1000 2000

T.O. MMC M



+40 TO 94.4 92.8 94.3 92.8 94.3 92.8 9

MC 93.1 91.7 93.1 91.7 83.1 91.6 9

+35 TO 95.7 94.3 95.7 94.2 95.6 94.2 9

MC 94.3 93.1 94.3 93.1 94.3 93.1 9

+30 TO 96.9 95.7 96.8 95.7 96.8 95.6 9

MC 95.5 94.6 95.5 94.6 95.5 94.5 9

+25 TO 97.9 97.0 97.9 97.0 97.9 97.0 9

MC 95.6 95.1 96.1 95.6 96.5 96.0 9

+20 TO 97.6 97.1 98.7 98.1 98.7 96.1 9

MC 94.7 94.2 96.0 95.7 97.3 97.2 9

+15 TO 96.8 96.3 98.2 97.9 99.3 98.6 9

MC 94.0 93.5 95.3 95.0 96.5 96.5 9

+10 TO 95.9 95.4 97.3 97.0 98.6 98.5 1

MC 93.1 92.6 94.4 94.0 95.6 95.6 9

+5 TO 95.1 94.6 96.5 96.2 97.8 97.7 9

MC 92.3 91.8 93.5 93.3 94.8 94.8 9

0 TO 94.3 93.8 95.6 95.3 99.6 96.8 9

MC 91.4 91.0 92.7 92.5 93.9 93.9 9

–5 TO 93.4 92.9 94.7 94.4 96.0 95.9 9

MC 90.6 90.2 91.2 91.7 93.1 93.1 9

–10 TO 92.5 92.0 93.8 93.5 95.1 95.0 9

MC 89.7 89.3 91.0 90.8 92.2 92.2 9

TO 91 7 91 2 93 0 92 7 94 2 94 1 9


ALT 5000 6000 7000


WITHOUT ANTI-ICE


T.O.MC



+40 TO 94.3 92.6 94.2 92.6

MC 92.9 91.4 92.7 91.2

+35 TO 95.6 94.2 95.5 94.1 95.5 94.1

MC 94.2 92.9 94.0 92.7 93.7 92.4

+30 TO 96.8 95.6 96.7 95.6 96.7 95.5

MC 95.3 94.4 95.1 94.2 94.8 93.9

+25 TO 97.8 96.9 97.8 96.8 97.8 96.8

MC 96.4 95.8 96.2 95.6 95.9 95.3

+20 TO 98.7 98.0 98.6 98.0 98.6 98.0

MC 97.3 97.1 97.1 96.9 96.9 96.6

+15 TO 99.4 99.1 99.4 99.0 99.4 99.0

MC 98.2 98.2 98.0 98.0 97.8 97.8

+10 TO 100.0 100.0 100.0 100.0 100.0 100.0

MC 98.5 98.9 98.7 99.0 98.5 98.8

+5 TO 100.0 100.0 100.0 100.0 100.0 100.0

MC 98.4 98.9 99.3 99.9 99.1 99.7

0 TO 100.0 100.0 100.0 100.0 100.0 100.0

MC 97.9 98.5 99.2 100.0 99.4 100.0

–5 TO 100.0 100.0 100.0 100.0 100.0 100.0

MC 97.0 97.7 98.3 99.3 99.2 99.7

–10 TO 99.5 100.0 100.0 100.0 100.0 100.0

MC 96.1 96.8 97.4 98.3 98.7 99.2

TO 98 6 99 3 100 0 100 0 100 0 100 0


T.O. M

MC M


WITH ANTI-ICE




MC M


+10 TO 95.1 94.5 96.5 96.1 97.9 97.6 9

MC 91.4 90.3 92.8 90.9 94.1 91.4 9

+5 TO 94.3 93.7 95.7 95.3 97.1 96.8 9

MC 90.6 89.5 92.0 90.6 93.3 91.6 9

0 TO 93.5 92.8 94.9 94.4 96.2 95.9 9

MC 89.8 88.7 91.2 90.2 92.5 91.7 9

–5 TO 92.6 92.0 94.0 93.5 95.3 95.0 9

MC 89.0 87.9 90.3 89.6 91.6 91.9 9

–10 TO 91.7 91.1 93.1 92.6 94.4 94.1 9

MC 88.1 87.0 89.4 88.9 90.7 90.0 9

–15 TO 90.9 90.3 92.2 91.8 93.5 93.2 9

MC 87.3 86.2 88.6 88.1 89.9 89.2 9

–20 TO 90.0 89.4 91.3 90.9 92.6 92.3 9

MC 86.4 85.3 87.7 87.2 89.0 88.3 8

–25 TO 89.1 88.5 90.4 90.0 91.7 91.4 9

MC 85.4 84.5 86.9 86.2 88.1 87.4 8

–30 TO 88.2 87.6 89.5 89.1 90.8 90.5 9

MC 84.7 83.6 86.0 85.1 87.2 86.5 8

ALT 0 1000 2000

T.O

MC


WITH ANTI-ICE




MC


+10 TO 98.1 97.7 98.1 97.6 98.0 97.5

MC 95.5 93.2 95.4 91.2 95.2 91.0

+5 TO 98.5 98.6 98.8 98.6 98.7 98.5

MC 95.8 93.6 97.9 92.5 96.9 92.3

0 TO 99.5 99.5 99.5 99.5 99.4 99.5

MC 96.1 94.0 96.9 93.8 96.7 93.6

–5 TO 99.1 99.3 99.8 99.8 99.7 99.8

MC 95.4 94.3 96.5 94.8 96.8 94.6

–10 TO 98.6 99.1 100.0 100.0 100.0 100.0

MC 94.7 94.6 96.0 95.8 96.8 95.6

–15 TO 97.8 98.2 99.2 99.5 100.0 100.0

MC 93.8 93.8 95.1 95.2 96.2 95.7

–20 TO 96.8 97.2 98.4 99.0 100.0 100.0

MC 92.9 93.0 94.2 94.5 95.6 95.7

–25 TO 95.9 96.3 97.4 98.0 99.1 99.5

MC 92.0 92.1 93.3 93.6 94.7 95.0

–30 TO 94.9 95.3 96.4 97.0 98.2 99.0

MC 91.1 91.1 92.3 92.6 94.7 94.2

ALT 5000 6000 7000


ANTI-ICE – MINIMUM N1 – 3 ENGINES O

–30°/–20° –20°/–10° –10°/

>20000 80% 76% 73%

TAT°C

ALT



NOTE

WITH ONE ENGINE INOPERATIVE I

N1 BY 9% IF N1>65% OR 6% IF N1<65%

>20000 80% 76% 73%

<20000 76% 73% 65%>10000

<10000 68% 65% 61%

PILOT NOTES




MAXIMUM CLIMB THRUS


CLIMB 260 kt/0.72

50 92.3 90.8 92.4 91.1


0 5,000 10,000 15,000



45 93.4 92.2 93.5 92.4 92.1 91.0

40 94.4 93.6 94.5 93.8 93.2 92.435 95.4 94.9 95.4 95.1 94.1 93.8 92.1 91.6

30 95.0 94.8 96.4 96.4 95.1 95.1 93.1 92.9

25 94.2 94.0 97.3 97.6 96.1 96.4 94.1 94.3

20 93.4 93.2 98.2 98.7 97.0 97.6 95.2 95.7

15 92.6 92.4 98.3 99.2 97.8 98.5 96.1 96.8

10 91.8 91.6 97.4 98.3 98.4 99.4 96.9 97.8

5 91.0 90.8 96.6 97.4 99.0 100.0 97.5 98.6

0 90.2 89.9 95.7 96.6 99.7 100.0 98.2 99.4

–5 89.3 98.1 94.8 95.7 100.0 100.0 98.9 100.0

–10 88.5 88.2 93.9 94.8 100.0 100.0 99.6 100.0

–15 87.6 87.4 93.0 93.9 99.5 100.0 100.0 100.0–20 86.8 86.6 92.1 93.0 98.5 100.0 100.0 100.0

–25 85.9 85.7 91.2 92.0 97.6 99.2 100.0 100.0

–30 85.0 84.8 90.3 91.1 96.6 98.2 100.0 100.0

–35 84.1 83.9 89.4 90.2 95.6 97.2 100.0 100.0

–40 83.2 83.0 88.4 89.2 94.6 96.1 100.0 100.0

–45 82.3 82.1 87.5 88.2 93.6 95.1

–50

T

A

T

(°C)

MAXIMUM CLIMB THRU

ALL ENGINES OPERATINGCLIMB 260 kt/0.72

50


25,000 30,000 35,000 40,000



45

4035

30

25

20 94.4 95.3

15 95.6 96.6 95.6 96.7

10 96.6 97.7 96.6 97.8 96.2 97.4

5 97.4 98.6 97.5 98.7 97.1 98.4

0 98.1 99.4 98.3 99.7 97.9 99.4

–5 98.9 100.0 99.2 100.0 98.8 100.0 98.1 99.

–10 99.6 100.0 100.0 100.0 99.7 100.0 99.1 100

–15 100.0 100.0 100.0 100.0 100.0 100.0 99.9 100–20 100.0 100.0 100.0 100.0 100.0 100.0 100.0 100

–25 100.0 100.0 100.0 100.0 100.0 100.0 100.0 100

–30 100.0 100.0 100.0 100.0 100.0 100.0 100.0 100

–35 100.0 100.0 100.0 100.0 100.0 100.0 100.0 100

–40 100.0 100.0 100.0 100.0 100.0 100

–45 99.6 99.6 99.6 99.6 99.6 99.

–50 98.5 98.5 98.5 98.

T

A

T

(°C)


MAXIMUM CLIMB THRUS


CLIMB 300 kt/0.8

50 92.7 91.6 92.8 91.9 91.5 90.7

45 93 7 92 9 93 8 93 2 92 5 92 0


0 5,000 10,000 15,000



45 93.7 92.9 93.8 93.2 92.5 92.0

40 94.7 94.2 94.7 94.5 93.4 93.3 91.4 91.235 94.9 94.8 95.7 95.7 94.4 94.5 92.4 92.5

30 94.1 94.0 96.6 97.0 95.4 95.8 93.4 93.8

25 93.3 93.2 97.5 98.1 96.3 97.0 94.5 95.1

20 92.5 92.5 97.8 98.7 97.2 98.1 95.4 96.3

15 91.8 91.7 96.9 97.9 97.9 99.0 96.3 97.4

10 90.9 90.9 96.1 97.0 98.6 99.8 97.1 98.2

5 90.1 90.1 95.2 96.2 99.2 100.0 97.8 99.0

0 89.3 89.2 94.4 95.3 99.9 100.0 98.4 99.7

–5 88.5 88.4 93.5 94.4 99.5 100.0 99.1 100.0

–10 87.6 87.5 92.6 93.6 98.5 100.0 99.8 100.0

–15 86.8 86.7 91.8 92.7 97.6 99.1 100.0 100.0–20 86.0 85.9 90.9 91.8 96.6 98.1 100.0 100.0

–25 85.1 85.0 90.0 90.8 95.7 97.2 100.0 100.0

–30 84.2 84.1 89.0 89.9 94.7 96.2 100.0 100.0

–35 83.4 83.3 88.1 89.0 93.8 95.2 100.0 100.0

–40 82.4 82.4 87.2 88.0 92.8 94.2

–45 81.6 81.5

–50

T

A

T

(°C)

MAXIMUM CLIMB THRU

ALL ENGINES OPERATING

CLIMB 300 kt/0.8

50

45


25,000 30,000 35,000 40,000



45

4035

30 92.9 93.7

25 94.1 95.1 94.3 95.4

20 95.2 96.3 95.4 96.6

15 96.2 97.4 96.5 97.7 95.9 97.1

10 97.1 98.2 97.4 98.7 96.9 98.1

5 97.8 99.0 98.3 99.6 97.7 99.1 97.1 98.4

0 98.5 99.8 99.0 100.0 98.6 100.0 97.9 99.4

–5 99.1 100.0 99.8 100.0 99.4 100.0 98.8 100.

–10 99.8 100.0 100.0 100.0 100.0 100.0 99.7 100.

–15 100.0 100.0 100.0 100.0 100.0 100.0 100.0 100.–20 100.0 100.0 100.0 100.0 100.0 100.0 100.0 100.

–25 100.0 100.0 100.0 100.0 100.0 100.0 100.0 100.

–30 100.0 100.0 100.0 100.0 100.0 100.0 100.0 100.

–35 100.0 100.0 100.0 100.0 100.0 100.

–40 100.0 100.0 100.0 100.0 100.0 100.

–45 99.6 99.6 99.6 99.6

–50 98.5 98.5 98.5 98.5

T

A

T

(°C)



LIMITED BY MAXIMUM CRUISE

TEMP

MI

LR 0.75 0.8 0.84

TEMP

MI

LR



Pressure altitude: 35,000 FT–15°C N.L. N.L. N.L. N.L.

–10°C N.L. N.L. N.L. 47,730

–5°C N.L. N.L. N.L. 46,250

0°C N.L. N.L. N.L. 45,270

5°C N.L. N.L. N.L. 43,240

10°C N.L. N.L. N.L. 40,13015°C N.L. N.L. 46,430 35,110

20°C N.L. 46,960 42,560 28,650


–15°C 47,860 48,070 45,650 39,910

–10°C 47,840 48,040 45,220 39,170

–5°C 47,200 47,340 44,280 38,090

0°C 46,540 46,630 43,300 36,970

5°C 45,215 44,870 41,570 34,810

10°C 43,610 42,990 39,740 31,500

15°C 40,960 40,290 37,270 26,80020°C 37,990 37,340 33,030

Pressure altitude–15°C N.L.

–10°C N.L.

–5°C N.L.

0°C N.L.

5°C N.L.

10°C N.L.15°C 46,41

20°C 43,54

Pressure altitude

–15°C 43,28

–10°C 43,22

–5°C 42,66

0°C 42,00

5°C 40,43

10°C 38,51

15°C 35,9420°C 33,11


LIMITED BY MAXIMUM CRUIS

TEMP

MI

LR 0.75 0.8 0.84

TEMP

MI

L




–15°C 39,240 39,260 37,190 32,440

–10°C 39,240 39,262 36,880 31,860

–5°C 38,580 38,590 35,950 30,780

0°C 37,760 37,730 34,990 29,650

5°C 36,150 36,100 33,400 27,090

10°C 34,170 34,120 31,53015°C 31,750 31,740 28,620

20°C 29,100 29,040 24,320


–15°C 32,100 32,130 30,360 26,410

–10°C 32,080 32,110 30,090 25,880

–5°C 31,150 31,360 29,120 24,760

0°C 30,030 30,250 28,020

5°C 28,420 28,640 26,480

10°C 28,460 26,670 24,350

15°C 24,250 24,51020°C

Pressure altitu

–15°C 35,4

–10°C 35,4

–5°C 34,7

0°C 33,8

5°C 32,3

10°C 30,315°C 28,1

20°C 25,3

Pressure altitu

–15°C 28,9

–10°C 28,8

–5°C 27,9

0°C 26,6

5°C 24,9

10°C 23,0

15°C20°C

P -1 8 B R

CRUISE—

WTX

1000

45

43

ALT

FF

N1

TAS

FF

N1

TAS

–10

710

89.9

420

685

88.9

420

ISA

732

92.1

430

707

91.0

430

+10

755

94.1

440

729

93.1

440

9

9

350



R ev i s i on

3

41

39

37

35

33

31

29

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TASFF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

420

666

88.1420

649

87.5

420

634

86.8

420621

86.2

420

609

85.7

420

598

85.2420

588

84.7

420

430

687

90.2430

669

89.5

430

654

88.9

430640

88.3

430

627

87.7

430

616

87.2430

605

86.7

430

440

708

92.3440

689

91.6

440

673

90.9

440659

90.3

440

646

89.7

440

633

89.1440

622

88.6

440

8

8

8

8

8

8

8

R ev i s i on

3

WTX

1000

45

43

ALT

FF

N1

TAS

FF

N1

TAS

–10

720

99.6

426

ISA +10 –

410

CRUISE—



P -1 9 B R

41

39

37

35

33

31

29

TAS

FFN1

TAS

FF

N1

TAS

FF

N1

TASFF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

426

66696.9

420

625

94.7

418

593

92.9

418563

91.3

418

538

89.9

418

518

88.6

418

501

87.7

418

68999.2

430

646

97.0

428

613

95.1

428582

93.5

428

555

92.0

428

534

90.8

428

517

89.8

428

633

97.3

438600

95.6

438

573

94.1

438

550

92.8

438

533

91.9

438

6

9

45

9

4

5

9

4

5

9

4

4

8

4

P -2 0 B R

CRUISE—

WTX

1000

45

43

ALT

FF

N1

TAS

FF

N1

TAS

–10

778

91.8

446

756

90.9

446

ISA

804

93.9

457

781

93.1

457

+10

830

96.1

467

806

95.2

467

–

7

9

4

7

9

4

350



R ev i s i on

3

41

39

37

35

33

31

29

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TASFF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

446

737

90.2

446

720

89.5

446

704

88.9

446690

88.3

446

676

87.8

446

664

87.3446

653

86.9

446

457

761

92.3

457

743

91.6

457

727

91.0

457711

90.4

457

697

89.9

457

684

89.4457

673

88.9

457

467

785

94.4

467

766

93.7

467

749

93.1

467733

92.5

467

718

91.9

467

705

91.4467

693

90.9

467

4

7

9

4

6

9

4

6

8

46

8

4

6

8

4

6

84

6

8

4

R ev i s i on

3

WT

X1000

45

43

ALT

FF

N1

TAS

FF

N1

–10 ISA +10 –

410

CRUISE—



P -2 1 B R

41

39

37

35

33

31

29

TAS

FFN1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FFN1

TAS

FF

N1

TAS

703

97.7

444

659

95.4

444

620

93.4

444

590

91.8

444

57090.8

444

553

89.8

444

682

97.7

454

642

95.6

454

610

94.0

454

58992.9

454

571

92.0

454

663

97.8

465

630

96.2

465

60795.1

465

589

94.1

465

6

9

4

5

9

4

59

4

5

9

4

P -2 2 B R

MAXIMUM C

WTX

1000

45

43

ALT

FF

N1

TAS

FF

N1

TAS

–10

1023

99.3

470

1022

99.1

473

ISA

996

99.7

477

1001

99.6

480

+10

926

98.9

479

931

98.6

484

350



R ev i s i on

3

41

39

37

35

33

31

29

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

990

97.9473

961

96.9

473

984

95.9

473

1039

99.1

480

1017

98.3

480

990

97.4480

967

96.6

480

1004

99.6483

996

99.2

484

967

98.2

484

941

97.3

484

1015

99.4

491

1016

99.4491

1002

98.9

492

934

98.7487

936

98.6

489

838

98.6

491

940

98.5

493

942

98.4

495

943

98.4497

945

98.3

499

R ev i s i on

3

WTX

1000

45

43

ALT

FF

N1

TAS

FF

N1

TAS

–10

725

99.8

427

ISA +10 –

410

MAXIMUM CR



P -2 3 B R

41

39

37

35

33

31

29

FF

N1

TAS

FF

N1

TAS

FF

N1

TASFF

N1

TAS

FF

N1

TAS

FF

N1

TAS

FF

N1

TAS

743

99.9

442

754

99.8

452

762

99.6

459768

99.5

464

758

98.8

467

725

97.1467

698

95.8

467

712

100.0

438

726

99.9

454

733

99.8

463739

99.6

469

744

99.4

475

747

99.3478

723

98.1

478

661

98.3

450671

98.1

467

675

97.9

474

678

97.7479

680

97.6

483

6

1

46

9

4

6

9

4

6

94

6

9

4

P -2 4 B R

S + 40 LFALCO

SLATS + FLAPS

G.W. PRX VREF

1000 SEA LEVEL 2000’

24 100 2150/3600 2200/3700



R ev i s i on

3

26 104 2250/3700 2375/3900

28 108 2400/4000 2450/4100

30 112 2475/4100 2600/4300

32 116 2600/4400 2700/4500

34 119 2775/4600 2900/4800

36 122 2950/4900 3050/5100

38 126 3510/5200 3300/5500

40 129 3400/5700 3500/5900

42 132 3600/6000 3700/6150

44 135 3600/6100 3700/6300



INSERT LATEST REVISED PAGE. DESTROY SU

LIST OF EFFECTIVE PAGES

Dates of original and changed pag

Revision .........3 .........August 1999Revision .........4..............May 2000

Revision .........5 ..............July 2001Revision .........5.01........June 2002Revision .........6 .........August 2006

TOTAL NUMBER OF PAGES IN THIS PUBLCONSISTING OF THE FOLLOW




Title ..............................................6ii ...................................................6E-i.................................................5E-ii................................................6E-1 – E-21....................................6A-i.................................................5A-ii – A-iii......................................6A-iv...............................................5A-v................................................6A-vi...............................................5A-vii ..............................................6A-viii .............................................5

A-1 – A-2......................................5A-3 – A-4......................................6A-5 ...............................................3A-6 – A-7......................................5A-8 – A-11....................................6A-12 – A-13..................................5

PageNo.

A-14 – A-69S-1 ............S-2 – S-3....S-4 ............S-5 ............S-6 ............S-7 ............S-8 ............S-9 ............QR-1 .........QR-2 .........QR-3 .........

QR-4 .........QR-5 .........QR-6 – QR-QR-8 – QR-QR-10 – QRQR-12 .......

EMERGENCY PROCEDU

Compliance with the order prescribed for application is recommended. Where more than one phase is specif

PHASE I Specifies immediate action to be accomplineed for reference to the checklist.



PHASE II Items shall be completed only after PHAScompleted by checklist.

PHASE III Items shall be completed as soon as time

AURAL WARNINGS shall be identified before silencinaural warnings will enable better coordination during the

the emergency procedures.

FALCON 900A OR B PILOT CHECKLIST

WARNING PANEL

CMPTR 3CMPTR 2CMPTR 1 ——

OIL 3OIL 2OIL 1 L. PITOT



GEN 3GEN 2GEN 1 AOA PROBE

BAT 2BUS TIEDBAT 1HOTBAT

FUEL 3FUEL 2FUEL 1XTK 2OPEN C

LOFUEL 1

LOFUEL 2

LOFUEL 3

AILZERO

FUELINGREV

UNLOCK—— AP M

7 7 7

4

COND'GOVHT

ECUOVHT

BLEEDOVHT

NOSECONE OVHT

EMERGENCY PROCEDU

ENGINE FIRE

INTERNAL ENGINE FIREON THE GROUND



ON THE GROUND

APU FIRE

FIRE IN BAGGAGE COMP

MAIN WHEEL WELL OVE

AIR CONDITIONING SMO

ELECTRICAL SMOKE OR

SMOKE REMOVAL

INADVERTENT THRUST RDEPLOYMENT IN FLIGH

TWO ENGINES INOPERATAPPROACH AND LANDIN

TWO ENGINES INOPERATGO-AROUND

ALL ENGINES INOPERAT

LOSS OF BOTH HYDRAUL

FIRE

SMOKE

ENGINES

HYDRAULIC

ENGINE FIRE

NOTE

The following procedures must be followenot the FAULT light is on.


1

Phase I

1 Power Lever of Engine Concerned



Phase II

6. Booster Switch..............................................

Light ..........................................

7. GEN Switch ..................................................

Light ..........................................

8. Engine Anti-ice .............................................

and possibly Light..PUMPOIL

GEN

FUEL

1. Power Lever of Engine Concerned...............

2. FUEL SHUTOFF Switch of Engine Concern

Light.........................................

3. Airspeed.......................................................

4. Fire Extinguisher DISCH Switch ..................

If fire warning persists:

5. Fire Extinguisher DISCH Switch ..................

TRANS

INTERNAL ENGINE FIRE ON THE

During starting attempt or shutdown:

Phase I

1. Power Lever .....................................................

2. FUEL SHUTOFF Switch ...................................

TRANS



Light ..............................................

3. Start Selector ..............................................MO

4. Start Button .......................................................

5. Bleed Air Switches (Crew and Passenger) .......

6. Evacuation ..............................................INITIA – – – END – – –

APU FIRE

NOTE

The following procedure must be followed not the FAULT light is on.

Phase I1. APU Master Pushbutton Light..........................

TRANS

FIRE

FIRE INBAGGAGE COMPARTMENT


Phase I

1. BAG Switch..................................................

In flight at or below 41,000 ft and with thement pressurized:

B



Phase II3. Land as soon as possible.

2

p

2. The co-pilot dons the smoke hood and fightsheld fire extinguisher.

The lavatory door must be properly clo

opening the baggage compartment doosmoke from entering the passenger cabin

During takeoff or landing, or in flight above baggage compartment not pressurized:

1. BAG Switch..................................................

Light.........................................

2. BAG COMP Extinguisher Switch .................

Do not open the baggage compartment door until t

BAG ISOL

CAUTIONOR

AIR CONDITIONING SMOK

Smoke at Air Conditioning Outlets

WARNING

Phase I1. Crew Oxygen Masks and Smoke Goggles.......

2. Microphone Selector ........................................



Phase II6. Crew Gaspers...................................................7. Isolation Valve Knob.........................................

Light ..............................................

8. Crew Bleed-Air Switch......................................9. COND Control Lever ........................................

If the smoke persists:

10. Crew Bleed-Air Switch......................................11. Passenger Bleed-Air Switch .............................

If the smoke disappears: If the sm

12. Continue the flight 12. Crew Tempwith the faulty Controller ..bleed air system isolated.

If the smoke disappears: If the sm

ISOL

3. “NO SMOKING” Sign.......................................4. Passenger Oxygen Controller .........................

5. Passenger Masks ............................................

ELECTRICAL SMOKE OR

Smoke and unusual odors.

WARNING


Phase I

1. Crew Oxygen Masks and Smoke Goggles...



Phase II

5. Crew Gaspers...............................................

If the origin of the fire If theor smoke is evident: smo

and 6. Suspected Equipment .....ISOLATED

Permit a total electrical Do noshutdown: electr

shutd

6. Gen 1, Bat 1, Bat 2,Gen 2, Gen 3Switches OFF

3

2. Microphone Selector ....................................

3. “NO SMOKING” Sign...................................

Only if there are no flames in the cabin.

4. Passenger Oxygen Controller andPassenger Masks ........................................

6. Bat 2 anSwitches

If fire or smoke is If fire o

visibly verified removed: visibly v

13. Continue the flight to the

nearest suitable airport.13. Land as so

If smo

14. Crew Air G

15. Descend t

or a safe a



The followmust notflames are

cabin or co

At 14,000 feet o

16. Pressuriza

UP—DN C

At a speed belo

17. LH Direct

Vision Win

PHASE III

18. Descend t

the safe a

or smoke i

removed.

CA

SMOKE REMOVAL


Phase I

1. Crew Oxygen Masks and Smoke Goggles...

2. Microphone Selector ....................................

3. “NO SMOKING” Sign...................................



Phase II

5. Crew Air Gaspers .........................................

6. Crew and Passenger Temperature Controlle

7. Descend to 14,000 feet or below or to the m

the route flown.

The following procedure must not b

flames are present in the cabin or cock


8 Pressurization Dump Switch DUMP (A1 BU

CAUTION

4

Only if there are no flames in the cabin:

4. Passenger Oxygen Controller and

Passenger Masks ........................................

INADVERTENT THRUST REVEDEPLOYMENT IN FLIGHT

and possibly lights on.

TRANSIT

DEPLOYEDREV

UNLOCK



Pitch down moment, abnormal noise and buffeting ma

Phase I

1. No. 2 Engine.....................................................

2. Thrust Reverser NORM/STOW Switch ............

3. Airspeed ...........................................................

Phase II

If the thrust reverser stows

4. Continue the flight with the NORM/STOW Switch

If the thrust reverser remains deployed:

4. Thrust Reverser NORM/STOW Switch ........MA


NOTE

The drag resulting from an idling No. 2 ethe thrust reverser deployed adversely a

OR

TWO ENGINES INOPERATAPPROACH AND LAND

Preparation

1. Reduce aircraft weight to the minimum pract

2. Determine the weight limitationfor enroute climb gradient ........................SE

3. Fuel Supply to Remaining Engine.................

4 BUS TIED S i h




4. BUS-TIED Switch .........................................Light ..........................................

5. Limit the On-Line Generator Load................3

6. Crew and Passenger Bleed-Air Switches.....

7. Brake Selector: (If Inoperative Engines are No. 1

8. ST-BY Hydraulic Pump Switch.....................9. Avoid icing conditions.

NOTE

With engine 1 and 3 inoperative—ExtenEmergency Procedure—Extend GEmergency Procedure.

Approach:

10. Flaps + Slats Handle ....................................

11. Emergency Slats Switch...............................

Make the decision to land or go-around at feet above ground level.

When committed for landing:

NOTE

BUS TIED

5

Free Fall Extension

15. Airspeed................................................NOT LE

Extend the main gear first, one at a time:

16. LH MAIN MANUAL GEAR RELEASE Handle ..

• Gently apply up to full rudder to the left while

KIAS max.) until the left green gear down ligh

Maintain wings level with appropriate aileron

NOTE



Illumination of the green gear down light

more than 30 seconds with full rudder deflec

• Gently come back to neutral rudder.

NOTE

Rapidly alternating large rudder applicationnation with large side-slip angles may result al failure at any speed.

17. RH MAIN MANUAL GEAR RELEASE Handle .

• Gently apply up to full rudder to the right whil

KIAS max.) until the right green gear down liged. Maintain wings level with appropriate aile

NOTE




TWO ENGINES INOPERATIVE—G

The decision to land or go-around must babove 1,000 feet above ground level. Theassociated with this Go-around procedumately 500 feet.

On the Go Around:

CAUTION




On the Go-Around:1. Engine Thrust (Full Power) ...........................

2. Landing Gear (If Down) ................................

The landing gear cannot be retractedand No. 3 engines are inoperative.

3. Accelerate the aircraft while in descent on n

At VREF + 25 knots:


5. Set and Maintain Enroute Climb Speed, (1.4

– – – END – – –

CAUTION

6

ALL ENGINES INOPERATI

Phase 11. Communications ...............................................

2. Establish the aircraft within the airstart envelop(See Abnormal Tab 3, page A-5.)

3. Reduce the electrical load to the lowest possibbattery current drains as possible, see note be

4 Relight the engines using the airstart procedure



4. Relight the engines using the airstart procedure(See Abnormal Tab 3 thru 7, page A-8 thru A-1

Phase 2

If no engine can be relighted:

5. Prepare for a Forced Landing or a Ditching. (Seor 9, page E-19 or E-20.)

6. ST-BY Hydraulic Pump (for EMERG Flaps + Sla

7. Extend the slats using the emergency system f necessary.

8. If a forced landing is anticipated, extend the

(190 KIAS) if possible, and if necessary perforEmergency Extension (Manual Unlocking). (Sepage A-24.)

NOTE

• Booster Pumps .........................................

• Windshield Heat .......................................

• Pitot ..........................................................


ENROUTE CLIMB GRADIE2 ENGINES INOPERATIVE AND



ENROUTE CLIMB GRADIEN2 ENGINES INOPERATIVE AND C



LOSS OF BOTH HYDRAULIC S

and possibly and

Both hydraulic indicators show a pressure drop.

PIT

FE

AIL

FEELPUMP 3

PUMP 2PUMP 1


Phase I



Phase II

3. Avoid high pitch attitudes and zones of air tu

Landing Preparation:


5. Landing Gear ............................................FR

See Abnormal Tab 13, page A-24 (Emergen

6. Approach Speed (Clean Configuration) ........

7. Rate of Descent ...................................APPR

The landing distance will be twice the tance. (Normal configuration flaps 40°a

CAUTION

Phase I

1. Auto-Pilot and Yaw Damper.........................

2. Airspeed...............................................260 K

7

LOSS OF ALL THREE GENERA

and

Corresponding GEN switches possibly tripped.

1. Bus Voltages.....................................................

2. Generator and Battery Ammeters .....................

3. Battery Load ..................................SHED AS M

CAUTION

GEN 3GEN 2GEN 1



To limit the battery load, accomplish the fothe sequence shown;

4. Booster Pump Switches (3) ..............................

5. Windshield Heating (3) .....................................

In icing conditions it is imperative to keep the following

• Pilot windshield heating system.• Engine and wing anti-icing system.

6. Pilot and Co-pilot Pitot Heating ....................

7. RH AV Master Pushbutton ..........................

8. Lavatory and Galley Master Circuit Breakers

8a. Unnecessary Lights .....................................

Attempt to reset the last generator that failed:9. BUS TIED Switch ........................................

10. BAT Switch (On Side Concerned) ................

11. Power Lever (Of Engine Concerned)............• Engine Idle Setting ...................................

12. GEN Switch (Of Generator Concerned) .......

If voltmeter:

RAPID DEPRESSURIZAT

Phase 11. Crew Oxygen Masks.....................................

2. Microphone Selector .....................................

3. and Light Pushbutto

4. Oxygen Controller and Passenger Masks..O

FASTEN BELTS




yg g

5. Emergency Descent......................................

– – – END – – –

EMERGENCY DESCEN

This procedure assumes structuralthe aircraft. If structural integrity is qlimit the airspeed to the lowest pracand avoid high maneuvering loads.

1. Autopilot........................................................

2. Power Levers................................................

3 Ai b k H dl

CAUTION

FORCED LANDING

Preparation:1. Communications Transmission .......................T

2. Transponder .....................................................

3. Passenger ........................................................

4. Fasten Belts/No Smoking Signs.......................



g g

5. Light .............................................

6. Cockpit Jump Seat (If Possible) ............UNOCC

Approach:

7. Wing Anti-Ice ....................................................

8. Crew and Passenger Bleed Air Switches.........

9. HP 1, PRV 2 and PRV 3 Bleed Air Switches....

10. Pressurization Switch........................................

11. Landing Gear ....................................................

12. Flaps + Slats.....................................................4

13. Approach Speed...............................................

Just Before Touchdown:14 Vertical Speed APPRO

AFT CABIN ISOL

DITCHING

Preparation:

1. Communications Transmission .....................

2. Transponder.................................................

3. Passenger ....................................................

4. Life Jackets...................................................

5. Fasten Belts/No Smoking Signs...................




6. Light ..........................................

7. Cockpit Jump Seat (If Possible) ............UNO

8. Audio Warn A (A1 Bus)and B (B1 Bus) Circuit Breakers .................

Approach (Parallel to the Major Swell):

9. Wing Anti-Ice ................................................

10. Crew and Passenger Bleed Air Switches.....

11. BLEED AIR BAG Switch...............................

Light .........................................

12. HP 1, PRV 2, and PRV 3 Bleed Air Switches

13. Pressurization Switch...................................

14. Landing Gear................................................

15. Flaps + Slats.................................................

16 Approach Speed

BAG ISOL

AFT CABIN ISOL

PILOT NOTES



ABNORMAL PROCEDU

Procedures in this section address foreseeable situatioin which the systems redundancy or selection of an amaintain an acceptable level of air worthiness.

I. IMMEDIATE ACTION—There are no memorythese procedures.

When a failure occurs, the crew should asrelationship to the immediate effect on air



relationship to the immediate effect on airand the continuance of the planned flight path.

Procedures that could have an immediate effecthat are fundamental to basic airmanship are:

• Air Data System Failures—Refer to stantransfer control.

• Engine Vibration (Actual)—Reduce Thrust.

• Pitch Trim Runaway—Override or disconne

II. SPECIFIC PRIORITIES—Where there is no

required, it is advisable to complete a Normal Chor due, prior to calling for the Abnormal Checkl

Because of the interrelationship of systems, have an effect on other systems.

The system that can produce multiple annunciat(2) Electric, (3) Hydraulic and should be deanumbered.


WARNING PANEL

77 7 21

777

CMPTR 3CMPTR 2CMPTR 1 ——

OIL 3OIL 2OIL 1 L. PITOT

GEN 3GEN 2GEN 1 AOAPROBE



1

17 17 17

161616

151515 16

17 20

BAT 2BUS TIEDBAT 1HOTBAT

FUEL 3FUEL 2FUEL 1XTK 2OPEN

LOFUEL 1

LOFUEL 2

LOFUEL 3

AILZERO

FUELINGREV

UNLOCK—— AP

18

COND'GOVHT

ECUOVHT

BLEEDOVHT

NOSECONE OVHT

29

18 18

8

18 18 18

ABNORMAL PROCEDU

ENGINE FAILURE BEFORE V1

ENGINE FAILURE AFTER V1

ENGINE FAILURE IN FLIGHT—S

ONE ENGINE INOPERATIVE—

APPROACH AND LANDING

ONE ENGINE INOPERATIVE—

GO-AROUND

LANDING DATA

ENGINE AIRSTART



ENGINE—AIRSTART

ENGINE FLAME OUT

AND HIGH SPEED AIRSTART

ENGINES-ABNORMAL AIRSTAR

AIRSTART COMPUTER IN NORMA

AIRSTART-COMPUTER IN MANU

FUEL CONTROL COMPUTER INOP

ENGINE OIL

NO. 2 ENGINE INLET DOOR OPE


LOSS OF NO. 1 HYDRAULIC SYS

FAILURE OF A NO. 1 SYSTEM P

LOSS OF NO. 2 HYDRAULIC SYS

UNWANTED OPERATION OF STLANDING WITH INOPERATIVE S

ENGINES

HYDRAULICS




THIS PAGE INTENTIONALLY LE

LOW BOOSTER PUMP PRESSU

FUEL TRANSFER SYSTEM MALFAIRPLANES EQUIPPED WITH X

TANK LEVEL ABNORMALLY LOW

ON AIRPLANES WITHOUT XTKTANK LEVEL ABNORMALLY LOW

WITH XTK 2 SYSTEM

FUEL ASYMMETRY

FUELING LIGHT ON IN FLIGHT

ONE GENERATOR INOPERATIV

GENERATOR RESET PROCEDU

TWO GENERATORS INOPERATELECTRICAL

FUEL



TWO GENERATORS INOPERAT

BATTERY OVERHEAT

BATTERY FAILURE

EITHER AIR DATA COMPUTER IN

BOTH AIR DATA COMPUTERS INO

JAMMED OR ABNORMAL PILOTAND POSSIBLY STANDBY IASINDICATION AT HIGH ALTITUD

PITOT-STATIC SYSTEM-PROBE

ANTI-ICING MALFUNCTION

WING ANTI-ICE INOPERATIVE

AIRCRAFT WITHOUT BRAKEHEATING SYSTEM

WING ANTI-ICE UNWANTED OP

AIRCRAFT WITHOUT BRAKEHEATING SYSTEM

WING ANTI-ICE INOPERATIVE AIR

WITH BRAKE HEATING SYSTEWING ANTI-ICE UNWANTED OP

SYSTEMS

PITOT-STATICSYSTEM

ICE PROTECTION




THIS PAGE INTENTIONALLY LEF

DOORS UNLOCKED INDICATION

NO AUTOMATIC DEPLOYMENT PASSENGER MASKS

APU BLEED LIGHT

CRACK OR BUBBLES

HEAT SYSTEM INOPERATIVE

OUT-OF-TRIM CONDITION

MACH TRIM INOPERATIVE

ANGLE-OF-ATTACK STALL PROHEATING FAILURE

AUTOMATICFLIGHT CONTROL

SYSTEM (AFCS)

ANGLE-OF-ATTACK

WINDSHIELD

PRESSURIZATION



ANGLE-OF-ATTACK INDICATORHEAT SYSTEM INOPERATIVE

FAILURE OF ENGINEFIRE DETECTION SYSTEM

FAILURE OF APU FIRE DETECTIO

EITHER EADI CRT FAILURE

EITHER EHSI CRT FAILURE

SIMULTANEOUS FAILURE OF EAND EHSI CRTs ON THE SAM

SUCCESSIVE FAILURE OF EADAND EHSI CRTs ON THE SAME

LOSS OF ASCB CONTROL

EFIS—INVALID ATTITUDE DATA

EFIS—ANY HEADING DATA INV

EFIS—IRS ATTITUDE MISCOM

ANNUNCIATION WITH OR WITHEADING COMPARATOR

EFISMALFUNCTION

FIREPROTECTION

ATTACK




THIS PAGE INTENTIONALLY LEF

ENGINES

Aircraft will tend to yaw if failure is associaengine 1 or 3.

Light if failure of center engine.

ENGINE FAILURE BEFORE

Reject the takeoff:

ENG 2 FAIL

WARNING



1. Brakes .........................................................MA

2. Power Levers....................................................

3. Airbrake Handle................................................

4. Thrust Reverser ...............................................

ENGINE FAILURE AFTER

Continue the takeoff:

1. At VR ..............................................................R

2. Airspeed ...........................................................

3. Positive Rate of Climb ......................................

4. Crew and Passenger Bleed Air Switches.........

5. Wing Anti-Ice.....................................................

If the engine failure occurs at a speed above

CAUTION

ENGINE FAILURE IN FLIGHT—S

And possibly

Yawing tendency (Failure of No. 1 or 3 Engine)

Determine which engine has failed.

ENGINE SHUTDOWN

1. Power Lever Retardedand Maintained 1 Min to ...............................

2 Power Lever

PUMPGENOIL


2



2. Power Lever..................................................

3. Booster Switch..............................................

4. Gen Switch ...................................................

5. Engine Anti-Ice Switch..................................

6. Fuel Shutoff Switch (If Engine Cannot Be Re

In icing conditions, operate No. 2 eneven with the No. 2 engine shut dowlation valve must be open.

If Engine Shutdown Procedure was initiatedure during takeoff (after V1) five minutes and if obstacle clearance permits:

7. Crew and Passenger Bleed Air Switches .....

8. Maximum Continuous Thrust........................

If required see One Engine Inoperative Drift DoReference Tab 6 page QR-12)

CAUTION

ONE ENGINE INOPERATIVAPPROACH AND LANDIN

1. Hydraulic Implications .......................................

2. Electrical Implications........................................

3. Fuel Implications ..............................................

4. Landing and Climb Requirements.....................

5. EGPWS Flaps O’Ride Switch...........................

5a. Approach Configuration....................................

5b 20 Fl Sl t Ai d



5b. 20 Flaps + Slats Airspeed ................................Increase the LD by 200 ft.

5c. 7 Flaps + Slats Airspeed ..................................Increase the LD by 600 ft.

6. Crew Briefing ....................................................

7. Approach Checklists:

a. Passenger Door Curtain.............................

b. No Smoking Sign .......................................

c. Altimeters....................................................

d. Radio Altimeter DH.....................................

e. X-Bp (ALL–3) ..............................................

f. Landing Lights............................................

g. Flaps + Slats .........................................APP

h. Landing Gear Control.................................

8. Landing Checklists:

a. Landing Gear..............................................

b. Hydraulic Pressure.....................................

A i Skid

ONE ENGINE INOPERATIGO-AROUND

1. Maximum Engine Thrust...............................

2. Landing Climb Attitude..................................

3. Airbrake Handle............................................

If the approach was flown with 20 degree fla


When a positive rate of climb is established:


3




5. Landing Gear Control ...................................

6. Airspeed to 400 Feet AGL ............................

Complete items 7 thru 9 at no lower than 400 feet A

If the approach was flown with 7 degree flap



5. Landing Gear Control ...................................6. Airspeed to 400 Feet AGL ............................

At no lower than 400 feet AGL:

7. Level Flight Acceleration...............................

8. At V2

+ 25 Knots Flaps + Slats .....................

9. Enroute Climb Speed....................................

OR

ENGINE—AIRSTART

General

Do not attempt to relight an engine after an if the engine integrity is questionable, or if is not observed.

Wait ten seconds between two consecutitt t D t k th th

CAUTION

WARNING



attempts. Do not make more than three sairstart attempts.

In-Flight Relight Envelope

30

Altitude (x 1,000 FT)

20

10

MAXIMUM AIRSTART ALTITUDE IN MANUAL MOD

ENGINE FLAME OUT AND HIGH SP

N2 rpm 15% or higher:

NOTE

This immediate airstart procedure may behigh altitude, even at altitudes above tstart envelope.

1. Power Lever ...............................................IM

2. Start Selector Switch.....................................


4




Light .........................................

3. ITT Rise Within Ten Seconds .......................

4. Power Lever .................................................

After a successful relight:


Light ...........................................

6. Engine Instruments ......................................

If an airstart is unsuccessful, complete the ENGINE

Procedure (See Abnormal Tab 2, page A-2).

– – – END – – –

IGN

IGN

ENGINES—ABNORMAL AIRS

Abort an airstart whenever any one of the foll

occurs:

• The ITT does not rise within ten seconds after mo

to idle.

• The oil pressure does not rise within ten secon

• The ITT rises rapidly and approaches the 952°or 978°C (TFE731-5BR) limit.



• N1 remains close to zero when N2 = 20%

• N2 speed is not rising rapidly and smoothly afte

• If, during an airstart with the fuel computer in mN1 exceeds 80% with the power lever at idle.

1. Power Lever......................................................

2. Start Selector Switch ....................................MO

– – – END – – –

AIRSTART—COMPUTER IN NORMA

1. Preparation Phase

1. Establish Airplane within Airstart Envelopepage A-5.

2. Power Lever.................................................

3. Fuel Shutoff Switch.......................................

4. GEN Switch ..................................................

5. CMPTR Switch .............................................

6 BOOSTER Switch


5



6. BOOSTER Switch ........................................

7. Engine and Wing Anti-Ice Switches..............

8. BUS TIED Switch .........................................

Light .........................................2. Select Either Procedure A or B Below, Ba

Speed and N1 Other than Zero

Abort the airstart if any abnormal condition(s) Tab 4, page A-7.

If an airstart is unsuccessful, complete the Procedure (See Abnormal Tab 2, page A-2).

A. Windmilling Airstart—N2 more than 15

of N1 rotation

1. Start Selector Switch....................................

Light .........................................2. Power Lever..................................................

IGN

BUS TIED

B. Starter-Assisted Airstart-N2 less than 15

1. Start Selector Switch.........................................

Light .............................................

2. Start Switch.............................PUSH (TWO SE

At 10% N2 and indication of N1 rotation:

3. Power Lever......................................................

4. ITT Rise within Ten Seconds ............................

5. N1, Fuel Flow, Oil Pressure..............................

When N2 is above 50%

6 S S l S i h

IGN




7. Engine Instruments ..........................................

8. BUS TIED Switch .............................................

Light ..............................................

If the IGN light remains on after the engine is on svalues are within limits:


Light ..............................................


If the BUS TIED light stays on after the bus tied swFlight Norm:

9. Generator Volts and Amps ...............................END

IGN

BUS TIED

PUMPOILGENIGN

OR

AIRSTART—COMPUTER IN MAN

1. Preparation Phase

1. Establish Airplane Within Manual Mode AAbnormal Tab 3, page A-5.

2. Power Lever..................................................

3. Fuel Shutoff Switch.......................................

4. GEN Switch ..................................................

5. CMPTR Switch .............................................

6 BOOSTER Switch


6



6. BOOSTER Switch ........................................

7. Engine and Wing Anti-Ice Switches..............

8. BUS TIED Switch .........................................

Light .........................................

2. Select Either Procedure A or B Below, BaSpeed and N1 Other than Zero

Abort the airstart if any abnormal condition(s) ocTab 4, page A-7.

If an airstart is unsuccessful complete the EDOWN Procedure (See Abnormal Tab 2, pag

A. Windmilling Airstart—N2 more than 15than 10%


LightIGN

BUS TIED

6

If the BUS TIED Light stays on after the bus tied swFlight Norm:

7. Generator Volts and Amps ...............................

B. Starter-Assisted Airstart—N2 less than 1


Light .............................................

2. Start Switch.............................PUSH (TWO SE

When N2 speed is 15% and N1 rotation is observed

3 Power Lever

IGN

OR



3. Power Lever......................................................

4. ITT Rise within Ten Seconds ............................

When N2 is above 50%

5. Start Selector Switch (As Required) .............MO

6. Engine Instruments ..........................................


8. BUS TIED Switch .............................................

Light ..............................................

If the IGN light remains on after the engine is on svalues are within limits:


BUS TIED

PUMPOILGENIGN

OR

FUEL CONTROL COMPUTER INOPERA

1. Power Lever of the Affected Engine .............

2. Engine CMPTR Switch.................................

If the light stays on:

3. Engine CMPTR Switch .................................

Do not let ITT of effected engine exceed indicated I

Avoid rapid displacements of the power lever.

CAUTION

CMPTR




Maximum thrust may not be attainablemay be higher than normal. Fuel flow N1 speeds may be approximately 5% acceleration time is longer in manual m

– – – END – – –

ENGINE OIL

Illumination warns of low oil pressure condition or olubrication system.

If the indicated oil pressure is greater than 25 ps

1. Engine Thrust (If Possible)............................

2. Monitor oil pressure and temperature.

If the oil pressure indicated is less than 25 psi:

7

OR

NO. 2 ENGINE INLET DOOR OPEN

1. Power Lever......................................................If the engine surges or abnormal conditions are oEngine 2.

2. Power Lever......................................................

3. Fuel Shutoff Switch...........................................

4. BOOSTER Pump Switch ..................................

5. GEN Switch ......................................................



6. Engine Anti-Ice Switch......................................

In icing conditions, operate No. 2 engineven with the No. 2 engine shut-down. tion valve must be open.

7. BUS TIED Switch .............................................

Light .............................................

8. ST-BY Hydraulic Pump Switch.........................O

– – – END – – –


BUS TIED

CAUTION

HYDRAULIC SYSTE

LOSS OF NO. 1 HYDRAULIC

AND POSSIBLY

1. Hydraulic Pressure and Quantity ..................

2. Airspeed ....................................................26

SYSTEM STATUS

PITCHFEEL


PUMP 1

PUMP 3



LOSS OF R

Servo – actuator Barrel No. 1

Normal Slats control system Use EMERG and land withand VREF + 5

Normal and emergency landing gear Free Fall extecontrol systems.

# 1 Braking system Select # 2 Brawith antiskid s

Pitch Arthur Unit

NOTE

Increase Landing Distance by 6

ON APPROACH

3. New Bug Speed............................................(S Ab l T b 3 A 4)

8

LOSS OF NO. 2 HYDRAULIC SY

AND POSSIBLY

1. No. 2 Hydraulic Pressure and Quantity............

If the hydraulic quantity is normal and the standby p

2. Standby Pump Switch ......................................

3. Hydraulic Pressure (1500/2150 Psi) and Quantity

(The No. 2 hydraulic system is usable but witthe operating time of the components.)

If the hydraulic quantity is zero or the standby pum

AILFEEL

PUMP 2



y q y y p

2. Airspeed .....................................260 KNOTS/.7

3. ST-BY Pump Switch .........................................

SYSTEM STATUS

LOSS OF REM

Servo-actuator barrel No. 2

Roll Arthur Unit

Flaps System See NOTE Belo

Emergency Slats Use normal Slat

No. 2 Braking System Parking brake isaccumulator.

Nose wheel steering Use No. 1 brakinential brake pres

Airbrake System Increase the lan



FLIGHT CONTROLS

LANDING WITH INOPERATIVE ST

1. AUTO PILOT ....................................................

NOTE

As the flaps and slats are selected, the P

unit returns to the low-speed position, whica significant decrease of elevator feel force.

If the stabilizer is jammed in the +2 to –4 degr



If the stabilizer is jammed in the +2 to 4 degr


3. Airspeed............................................................

4. GPWS Flaps O’Ride Switch.............................


The landing gear not extended aural wa

not sound.

If the stabilizer is jammed in the –4 to –10 deg


3. Airspeed ...........................................................

– – – END – – –

CAUTIONOR


ARTHUR UNIT INOPERAT

OR

If the PITCH FEEL light is on reduce airspeed down to 26

1. FASTEN BELTS Light Pushbutton........................

The pitch and/or roll control forces may

lower than normal, depending on whethunit has failed in the “HIGH” or “LOW” spe

• LIGHT FORCES: Avoid large displacemmo ements of the fligh

CAUTION

PITCHFEEL

AILFEEL



movements of the fligh

OR

• HIGH FORCES: Use normal or emergency

For Approach and Landing

Light on: At VREF.

Light on: At VREF + 10 knots and increas

by 800 feet.

– – – END – – –

FLAP ASYMMETRY OR JAMM

May be on

With flaps extended up to 7°:

FLAPASYM

PITCHFEEL

AILFEEL

10



THIS PAGE INTENTIONALLY LEFT

SLAT MONITORING SYSTEM

• If the light comes on after takeoff or at a speed

1. Speed Range ...............................................

• If the light comes on at an speed of 280 knots

1. Reduce Airspeed .........................................

– – – END – – –

SLAT SYSTEM MALFUNC


OR



IN CRUISE:

and a possible rolling tendency.

1. Airspeed........................................................

2. Autopilot .......................................................

DURING APPROACH:

Follow the procedures for the appropriate “CASEslat indication occurs after the slats-flaps have bee

CASE 1: Slat/Flap handle at S+7°, flashing light out, 7°flaps extended.

INI

11

FLAPS

UP0

7

CLEAN

7°FLAPS–SLATS20°FLAPS–SLATS

2. EMERG Slat Switch .........................................

3. Follow the appropriate PROCEDURE below.

Do not change the position of the EMERGEswitch after selection.

Possibility 1 — ALL SLATS EXTENDED.

Steady GREEN + 7°Flaps.

4. Flap/Slat Handle ...............................................

CAUTION



5. Airspeed ...........................................................

Possibility 2 — Outboards visually checkedextended. Red light + 7°Flaps.

4. Flaps/Slats........................................................

5. Airspeed (Zero Wind)........................................

6. Increase the landing distance by 200 ft.

Possibility 3 — Outboards visually checked not extended.

Red light, no flaps extended.

4. Flaps/Slats Handle............................................


UNWANTED OUTBOARD SLAT E

AND

The red slats transit light comes on then goes outcomes on flashing and the aural stall warning soun

1. RH Auto Slat Circuit Breaker (B1 Bus) .........

If the slats do not retract:

2. LH Auto Slat Circuit Breaker (A1 Bus)..........3. RH Auto Slat Circuit Breaker ........................

4. Continue the flight at an indicated airspeed of ith th i it b k f th d f ti t

IGNAUTOSLATS



12

with the circuit breaker of the defective system

Depending on the final position of the circuit breake

If the RH AUTO SLAT C/B remains pulled:* Manual action of the flap-slat handle remains ope

* Emergency slats switch is inoperative.

If the LH AUTO SLAT C/B remains pulled:

* Manual action of the flap-slat handle:

• Causes only extension of outboard slats. Iretracted position and the slat green light is

• Has no action on flaps when flap-slat hdegrees position, but will operate flaps whenand 40 degrees positions.

– – – END – – –

AIRBRAKE(S) DO NOT RETRACT

FOR APPROACH AND LANDING:

1. Flaps + Slats Handle .......................................4

With the airbrake(s) extended to position 1:

2. Airspeed...........................................................


With the airbrake(s) extended to position 2:

2. Airspeed ...............................................................


OR



g y

– – – END – – –

AIRBRAKES DO NOT EXTEND IN

Increase the landing distance by 10%.

– – – END – – –


LANDING GEAR AND B

ABNORMAL LANDING GEAR E

If, after the LANDING GEAR handle has been plaction, and one or more of the following indicatioEMERGENCY LANDING GEAR EXTENSION.

• One or more green gear down indication IS N• The landing gear handle light is flashing.

• The landing gear not extended (“GEAR”) voice w



The landing gear handle must be mainta

Emergency Landing Gear Extension

1. EMERGENCY GEAR PULL Handle .............

If all three green gear down lights illumin

gear handle light (red) is not illuminated, the

and locked. Do not actuate any landing gear

If at least one (1) green gear light does n

landing gear handle light is flashing, apply th

SION procedure.

Free Fall Extension

1 Airspeed NOT

CAUTION

3. RH MAIN MANUAL GEAR RELEASE Handle .

• Gently apply up to full rudder to the right while

KIAS max.) until the right green gear down lig

Maintain wings level with appropriate aileron in

NOTE




4. NOSE GEAR MANUAL RELEASE Handle ......

Accelerate until illumination of the green gear li



Accelerate until illumination of the green gear li(190 KIAS max.)

NOTE

Free fall extension of all 3 landing gearapproximately 2 minutes to complete.

Do not actuate any landing gear control onclanding gears are locked down. The landingbe maintained down.

– – – END – – –

LANDING GEAR—ABNORMAL RE

CAUTION

NO. 1 BRAKE SYSTEM OR ANTI-SKI

Results of abnormal anti-skid test bef

1. Brake Selector Switch...................................

2. Braking................................................APPL

3. Increase the landing distance by 50%.

NOTE

For take-off operation of the airplane

WARNING




For take off operation of the airplane tional basis with anti-skid inoperative in annex 5 of the AFM. Operation on

subject to prior authorization defined

– – – END – – –

NO. 1 AND NO. 2 BRAKE SYSTEMS

The aircraft can be brought to a stop by simultanreverser and the parking brake system.

1. Runway Requirements.................................

2. Thrust Reverser............................................

3 Parking Brake INT

NOSEWHEEL STEERING INOPE

1. Release the steering control wheel to neutral.

2. Use differential braking to steer the aircraft.

– – – END – – –

NOSEWHEEL SHIMMY

Hold the nosewheel steering control depressed.

– – – END – – –



END


FUEL SYSTEM

LOW BOOSTER PUMP PRE

CASE 1

1. No. 2 Booster Switch....................................

If the FUEL 2 light stays on:

2. Associated Fuel Quantity..............................

If a significant fuel loss is evident:

3. No. 2 Engine Power Lever............................

4 N 2 E i F l Sh t ff S it h

FUEL 2



4. No. 2 Engine Fuel Shutoff Switch .................

Light ..........................................

5. GEN Switch ..................................................

6. Engine Anti-ice Switch..................................

In icing conditions, operate No. 2 engine anti-ice eveshutdown. The isolation valve must be open.

7. BUS TIED Switch (If Bus Loading is Normal)

8. ST-BY Hydraulic Pump Switch.....................9. Group 2 Fuel Tank Fuel................................

10. No. 2 Booster Switch ....................................

11. X-BP 1-2 and X-BP 3-2 ................................

1-2 and 2-3 Lights ..............

When No. 2 Fuel Tank is Consumed:12 No 2 Booster Switch

X-BPX-BP

TRANS

OR

If the fuel light stays on:

3. X-BP 1-3 ...........................................................

Light.....................................................

4. Associated Fuel Quantity..................................

If a fuel loss is evident:5. Associated Engine Power Lever.......................6. Associated Fuel Shutoff Switch ........................

Light .............................................

7. GEN Switch ......................................................8. Engine Anti-ice Switch......................................

See ONE ENGINE INOPERATIVE APPROACH AND LAbnormal Tab 2, page A-3.

If the fuel light goes out:The flight may be continued maintaining balanced fuel lev

TRANS

X-BP

OR



3. XTK Switch Set to Low Level Side....................

Light ....................................................

CASE 3 — Side engines supplied with X-BP 1

AND

1. X-BP 1-3 ...........................................................

1-3 Light ................................................

2. X-BP 1-2 and X-BP 3-2 ....................................If the and Lights Go Out:

3. No. 1 and No. 3 Booster Switches....................

4. Fuel Quantity Indicators....................................

To consume fuel in side tanks when no leak is sus

5. Flight Altitude .............................................31,0

6. X-BP 1-2, X-BP 1-3, and X-BP 3-2 ...................

FUEL 3FUEL 1

X-BP

FUEL 3FUEL 1

XTK

OR


FUEL TRANSFER SYSTEM MALFAIRPLANES EQUIPPED WITH XT

CASE 1

If No. 2 fuel total quantity indicates approxim(Green Range)

1. No. 2 Rear Tank Quantity .............................

If the No. 2 rear tank level is 3,300 pounds and

that the rear tank is full:2. XTK 2 Switch ................................................

Light .........................................

If the XTK 2 OPEN light does not go out and there is

XTK 2OPEN

XTK 2OPEN

OR



If the XTK 2 OPEN light does not go out and there is the rear CG limit:

3. Manual XTK 2 Valve ....................................C

Light .........................................

If the No. 2 fuel quantity indicates approximat(Amber Range)

1. No. 2 Rear Tank Quantity .............................

If the No. 2 rear tank level is above 1,400 pounds:

2. XTK 2 Switch ................................................

Light .........................................

NOTE

If the light does not go ou

pump 1 or3

has failed, or if side tank tion has been used, the attitude shall

XTK 2OPEN

XTK 2OPEN

XTK 2OPEN

TANK LEVEL ABNORMALLOW ON AIRPLANES WITHOUT XTK

OR OR

1. Associated Fuel Quantity ..................................

2. Associated X-BP...............................................

Light.....................................................

3. Booster of Affected Tank.................................. – – – END – – –

TANK LEVEL ABNORMAL

X-BP

LO

FUEL 3

LO

FUEL 2

LO

FUEL 1



TANK LEVEL ABNORMALLOW ON A/C WITH XTK 2 SYS

OR OR

This light indicates that a fuel level below 200 pounds isto a cruise flight time of 15 minutes at an altitude of 2,50

1. Associated Fuel Quantity Gage .......................

2. Associated X-BP...............................................

Light.....................................................

3. Booster of Affected Tank..................................

IFLO

FUEL 2

X-BP

LOFUEL 3

LOFUEL 2

LOFUEL 1


FUEL ASYMMETRY

1. Asymmetric Fuel Quantity Indications..........

2. Possible Abnormal Aileron Trim....................

CASE 1 —Side Tank Asymmetry:

3. X-BP 1-3.. .....................................................

Light.................................................

4. XTK Switch ............................................SET

Light.................................................

Only If Both Booster 1 and 3 Switches are On

XTK

X-BP



5. Booster Switch On Low Level Side ..............

CASE 2 —Tank 2 To Side Tank Asymmetry:3. X-BP 1-2 or X-BP 3-2 ...................................

Light.................................................

If No. 2 Tank Fuel Level is Higher:

4. Booster No.1 or No. 3 Switch .......................

If No. 2 Tank Fuel Level is Lower:

5. Booster No. 2 Switch....................................

– – – END – – –

FUELING LIGHT ON IN FLIGHT

X-BP




ONE GENERATOR INOPERAT

CASE 2 Switch is

1. Bus Voltages................................................2. Batteries and Generators Load....................

If LH Main Bus Voltage Is:

3. GEN 1 S

(2 reset aSee GenP d

GEN 1GEN 1



Procedur

4. If resettin

GEN 1 S

Above the green range

3. GEN 3 Switch ..............OFF

If indication is now

Light .......................ON

O

GEN 1GEN 3

CASE 3 Switch is n

1. Bus Voltages....................................................

2. Batteries and Generators Load.........................

If LH Main Bus Voltage Is:

No

3. GEN 3 Switch(2 Reset attemSee GeneratProcedure; T

4 If resetting c

GEN 3

GEN 3



4. If resetting cGEN 3 Switch

Above the green range

3. GEN 1 Switch ..............OFF

If indication is now

Light .......................ON Li

Light .....................OFF Li

4 LH M i B V l 4 GEN 3 S i

GEN 1GEN 3

GEN 3GEN 1


ONE GENERATOR INOPERAT

CASE 4 Switch is

1. Bus Voltages.................................................

2. Batteries and Generators Load.....................

3. GEN 2 Switch...............................................(2 Res

See Generator R


2. GEN 2 Switch ...............................................LH and RH Main Bus Voltage ...........NOT AB

GEN 2

GEN 2



g

3. Battery Ammeter ...........................................

If the battery load is normal:

4. BUS-TIED Switch .........................................

Light...........................................

5. Limit the load on the operating generators.

– – – END – – –

GENERATOR RESET PROC

1. Bus-Tie..........................................................

2 B tt S it h A i t d B

BUS TIED



THIS PAGE INTENTIONALLY LEFT B


18TWO GENERATORS INOPE

CASE 1 AND

If one or no generator switch has tripped:

1. Bus Voltages.................................................


Shed the load on the bus, if necessaryload on the operating generator andNever tie the buses without previously cthe voltages and amperages on each bthe prescribed limits.

CAUTION

GENGEN



the prescribed limits.

Do not attempt to reset the generatothe tripped switch.

On the generators for which the switcTRIPPED:

If

3. GEN 2 Swi(2 reset atteSee Gener

Procedure;

GEN 2

CAUTION

CASE 2 AND

GEN 2 and GEN 1 or GEN 3 switches have tr

1. Bus Voltages....................................................

2. Batteries and Generators Load.........................

Shed the load on the bus, if necessary, tload on the operating generator and/oNever tie the buses without previously chethe voltages and amperages on each busthe prescribed limits.

CAUTION

GEN1 OR 3

GEN 2

GEN 2



3. BUS-TIED Switch...............................CHECKE4. BAT No. 2 Switch..............................................

5. Power Lever of No. 2 Engine............................

• Engine Idle Setting.........................................

6. No. 2 Generator .........................................TRY


7. Power Lever of No. 2 EngineAfter Generator Reset Attempt .........................

8. LH and RH Main Bus Volts/Amps ....................

If volts and amps are normal:


TWO GENERATORS INOPERAT

CASE 3 AND

GEN 1 and GEN 3 switches have tripped:

1. Bus Voltages.................................................


Shed the load on the bus, if necessaryload on the operating generator andNever tie the buses without previously cthe voltages and amperages on each b

CAUTION

GEN 3GEN 1

G



g p gthe prescribed limits.

3. BUS-TIED Switch...............................CHECK

4. BAT No. 1 Switch..........................................

5. Power Lever of No. 1 Engine........................

• Engine Idle Setting.....................................6. No. 1 Generator .........................................TR




BATTERY OVERHEAT

Additionally, the battery temperature indicator red

annunciation and light on (mod 2629).

1. Associated BAT Switch ....................................

If the battery temperature keeps rising:


NOTEIf required, the faulty battery may be switche

for landing, provided the light h

out. Monitor the battery temperature indicat

HOTBAT



– – – END – – –

BATTERY FAILURE

OR Associated BA

1. Associated BAT Switch.....................................ATT

– – – END – – –

BAT 2BAT 1


PITOT-STATIC SYST

EITHER AIR DATA COMPUTER IN19


"On-Side" Airspeed Scale flag D

EFIS on EADI re

ASEL flag C

on EADI so

"On-Side" Vertical Modes Reversion to C

ASEL

IAS



On Side Vertical Modes Reversion to CAP-FD basic mode so

CPLD MSG: CPLD CDATA INVALIDon ID 802.

"On-Side" Altimeter Flags. Uflight rate-of-climb otinstruments indicators. sid

ID 802 SAT-TAT-TAS Dashes on DCPLD. corresponding "x"on-side." line.

Autoslats Extension Possibility Se

inhibition atlight.

M

high speed. Ta

AUTOSLATS

BOTH AIR DATA COMPUTERS INO


EFIS Airspeed Scale flag Use

Mach ASEL on EADI airsp

Wind on EHSI flag

on EADI

AP-M TRIM Autopilot, Mach flashing

YD-FD trim and yaw on EADIdamperCommand barsgo out of view.

on

warning panel

AP

AP

ASEL

IAS



warning panel

and

MSG: DADCDATA INVALIDin ID 802.

Flight Altimeters, rate Flags. UseInstruments. o f c l imb Instr

indicators.ID 802 SAT-TAT-TAS Dashes on Avoi

corresponding Condline. (See

Autoslats Extension light. See

inhibition at Malf

high speed. Tab

AUTOSLATS

MACHTRIM


20

JAMMED OR ABNORMAL PILOT, CPOSSIBLY STANDBY IAS/MACH

AT HIGH ALTITUDE

Pilot and copilot IAS/MI indications

abnormal and possibly:

VMO /MMO audio warning sounds; IAS EADI compa

nates; AP disengagement and/or DADC DATA INVAL

APMACHTRIM

AUTOSLATS

AILFEEL

WARNING



ID 802; and a disagreement with standby IAS/Mach in

If it is certain that the VMO/MMO warnin

not modify flying parameters.

1. Autopilot and Yaw Damper ...........................

If the VMO /MMO audio warning sounds:

2. AUDIO WARN AND AUDIO WARN B Circui

3. Use the standby altimeter to stabilize and fly

4. Engine Thrust Setting.........................MAXIM

CAUTION

C. Descent:

Follow either 1 below for operations without anti-ioperations with engine and wing anti-icing on.

1. Without anti-icing

6. N1 RPM.............................................................

7. Vertical Speed..............................................–2,

8. Aircraft Attitude ...............................................0

2. With engine and wing anti-icing on:

3. N1 RPM ............................................................(See N1 Anti-icing Table, Abnormal Tab 21, Pa

4. Airbrakes ..........................................................

5. Vertical Speed..............................................–1,



6. Aircraft Attitude ...............................................0

NOTE

Check airplane altitude frequently on thaltimeter.

If prior to the problems, flight was performedtemperature lower than the authorized min

(see table below), descend as soon as poair-data indications become normal again.

Re-engage CB's AUDIO WARN A and AUDB at intervals and leave them engaged ifwarning has stopped sounding.

If the IAS/MI indications are doubtful, thestatic temperature may be incorrect.


ICE PROTECTION

PITOT-STATIC PROBE ANTI-ICING

OR OR

Compare the instrument readings with the readsystems.

WING ANTI-ICE INOPERATIVE

WITHOUT BRAKE HEATING

ST BYPITOT

R PITOTL PITOT

21



CASE 1

The Wing Anti-Ice Switch is ON:

1. No. 1 Engine N1 Speed................................

GREEN

If Indication

2. No. 1 Engine N1 Speed.... 2. No. 1 Engine(In Icing) MAINTAIN Speed

Is now

WING ANTI-ICE INOPERATIVE AIWITHOUT BRAKE HEATING SYST

CASE 2


1. No. 1 Engine N1 Speed....................................AMBER

If Indication

2. No. 1 Engine N1 2. No. 1 EngineSpeed (In icing) ..........MAINTAIN

MINIMUM N1 3. No. 3 EnginSEE TABLE BELOW Speed ........

If Indicat

Is now



If Indicat

4. No. 3 Engine(In Icing) .....

Do not reducthe minimumditions (See T

– – – END – – – CASE 3


1. If in icing conditions do not reduce N1 speed belo2. Avoid or leave icing conditions as soon as poss

– – – END – – –

CASE 4

Is no


WING ANTI-ICE UNWANTED OP

AIRCRAFTWITHOUT BRAKE HEAT

The Wing Anti-Ice Switch is OFF:

1. Wing Anti-ice Switch.....................................

If the green wing light does not come on, it mthe indicating system:


If the green light comes on:


3 Isolation Valve Knob

22

OR



3. Isolation Valve Knob.....................................

Light .........................................

4. HP 1 and PRV 3 Switches............................

5. Reduce No. 1 and No. 3 engine thrust settings

– – – END – – –

WING ANTI-ICE INOPERAAIRCRAFT WITH BRAKE HEATI

CASE 1

A S it h i Wi iti

ISOL

WING-BRK

WING

OFF

If Indication

5. In icing conditions, do 5. Wing-Brk Switcnot reduce N1 speed * In icing conditionto less than 1% above speed below ththe specified value. (See Abnormal (See page A-47) N1 Anti-icing Table

* Avoid or leave soon as possible

.................................... * Wing anti-ice is n

B. Switch in Wing-Brk position:

1. Wing-Brk Switch.................................................

Is now rem

WING-BRK

WING

OFF



If indication

Brake heating 2. Wing-Brk Switcsystem inoperative

3. No. 1 Engine NSpeed.............

If Indication

Is now rem

Is now


WING ANTI-ICE INOPERAAIRCRAFT WITH BRAKE HEATING

CASE 2

Switch in Wing or Wing-Brk position:

1. No. 1 Engine N1 Speed ....REDUCE UNTIL AM

If indication

2. Maintain No. 1 Engine N1 . 2. No. 1 EngineS

Is now re

WING-BRK

WINGOR

OFF

W

WO



Speed

3. No. 3 Engine

If Indi

.........................................................

4. Maintain No. 3 Engine N1 If selectedSpeed

WingDo not reduce N1 speed Do n

below the minimum to le

Is now rem

CASE 4

Green and Amber lights out

A. Switch in Wing position:

1. Wing-Brk Switch.....................................................

If the amber flashes and then goes out:

2. Wing-Brk Switch ....................................................

If theAmber light comes on steady, and then goeconfirmation of wing anti-ice valve operation.

If the above indications are not observed, if innot reduce N1 speed below the specified valTab 21, page A-47, N1 Anti-icing Table.)

WING-BRK

WING

OFF




B. Switch in Wing-Brk position:

1. Wing-Brk Switch ....................................................

If indication

The brake heating system 2. Wing-Brk Switcshould be considered as

inoperative.

Is now rem

WING-BRK

WING

OFF


WING ANTI-ICE UNWANTED OAIRCRAFT WITH BRAKE HEATI

The Wing-Brk Switch is Off:

1. Wing-Brk Switch............................................

If Indication is

Is now re

2 Wing Brk Switch OFF 2 Wing Brk Sw

23

WING-BRK

WINGOFF



2. Wing-Brk Switch........OFF 2. Wing-Brk Sw

........................................... (Malfunction o

3. Isolation ValveKnob ............ISOLATION

Light.......ON

4. HP 1 and PRV 3Switches ..................OFF

5. Reduce Engine 1 and 3thrust setting as soon aspossible.

– – – END – – –

ISOL

ENGINE ANTI-ICE INOPERAT

CASE 1 With associated

Eng 1, Eng 2, or Eng 3 amber light on steady.

1. Increase thrust on the corresponding engine ungoes out and the green light illuminates.

2. Retain this thrust setting while in icing condition

If the amber Light does not go out:

3. If in icing conditions, do not reduce N1 speed bvalue (See Abnormal Tab 21, page A-47, N1 A


– – – END – – –




Eng 1 or Eng 2 or Eng 3 amber and green light on.

1. If in icing conditions, do not reduce N1 speed bvalue (see Abnormal Tab 21, page A-47, N1 An

2. Avoid or leave icing conditions as soon as poss – – – END – – –

CASE 3

The No. 2 engine anti-ice switch is on, but the green

1 No 2 Engine Anti-Ice Switch


ENGINE ANTI-ICE UNWANTED O


Eng 1 or Eng 3 amber light flashing.

1. Associated Engine Anti-Ice Switch...............

If the Indication

And the TAT ............ 2. Assis above +10°C Sw

2. Associated engine thrust reduced.

Is now r



– – – END – – –

CASE 2 With Eng 2 switc

Eng 2 amber light flashing.

1. No. 2 Engine Anti-Ice Switch........................

If the Indication

2 No 2 Engine 2 No 2 Engine

24

Is now

ICE PROTECTION–LATE ACTIV

1. Start Selector Switches (3) ...............................

2. No. 1 and No. 2 Engine Anti-Ice Switches .......

After waiting 30 seconds:

3. No. 3 Engine Anti-lce Switch............................

After waiting 30 seconds:

4. Wing or Wing-Brk Anti-Ice Switch .....................

5. Start Selector Switches (3)(When No Longer Required).............................

CAUTION



– – – END – – –


AIR CONDITIONIN

BLEED-AIR SYSTEM OVER

The following is a step-by-step identification proces

1. PRV 3 Switch................................................

If the light

Starts Blinking Starts BlinkingAnd Then Goes Out And Keeps Blinking

BLEEDOVHT

BLEEDOVHT



25

p g

2. Continue flight inthis configuration.

In Icing Conditions In Nand

2. Engine associated with to reblinking light ............REDUCE TO engi

IDLE THRUST

ECU OVERHEAT

1. Passenger Temperature Controller .................

2. Passenger Air-Conditioning Valve Switch (If R

3. COND Control Lever ........................................

If the ECU OVHT light stays on:

4. Passenger Air-Conditioning Valve Switch .......

5 Crew Temperature Controller...........................

6. Crew Air-Conditioning Valve Switch (If Require

ECUOVHT



If the ECU OVHT light is still on:

7. Airspeed .......................................................LE

NOTE

If not in icing conditions, the HP 1, PRV 2, switches may be turned off.

If the overheat warning persists:

8. Passenger Air-Conditioning Valve Switch .......


– – – END – – –


CABIN AIR CONDITIONING OVEAIRCRAFT EQUIPPED W

ANTI-ICING EMERGENCY CONTR

High air temperature occurs along w

1. Temperature Controllers...............................

Since anti-icing on the turbo cooler t

longer ensured pushbutton

used at high altitude (above 35,000 f

ECU A/I

CAUTION

COND’GOVHT



water concentration is practically cl

IF BELOW 35,000 FT SEE NOTE BELOW

If the temperature is not decreasing:

2. ECU A/I Pushbutton .....................................

Light ..................

After Selection

If the temperature

ECU A/I

ECU A/I26




PRESSURIZATION

IMPROPER CABIN VERTICAL

1. Pressurization Up-Dn Knob ...........ALIGNED

2. PRV 2 and 3 Switches ..................................

3. Bleed Air Crew and Passenger Switches .....

2

1

1

2

0

-1

UP

DN



If normal operation is not restored:

4. Pressurization Up-Dn Knob .................1 TO

5. Auto/Man Pressure Selector Switch .............

6. Pressurization Up-Dn Knob .......................A

– – – END – – –

TOO HIGH CABIN PRESS27

6

5

4

32

10-1

7

TOO HIGH CABIN ALTITUDOR

SLOW DEPRESSURIZATIO

The aural warning occurs along with th

Cabin altitude higher than 10,000 feet

1. Crew Oxygen Masks.........................................

2. Microphone Selector.........................................

3. Bleed Air Crew and Passenger Air-Conditionin

3a. PRV 2 and 3 Switches ......................................

4. BAG Switch ......................................................

Light ..............................................BAGISOL

CABIN



5. Nose Control Lever (If Installed) .......................6. Pressurization Up-Dn Knob ..............................

7. Auto/Man Pressure Selector Switch .................

8. Pressurization Up-Dn Knob.............................D

If cabin pressure cannot be restored:

9. Isolation Valve Knob.........................................

Light ..............................................

If cabin pressure is restored If cabin pressu

10. Cycle bleed air Passenger ... 10..NORM/EM

and/or Crew switches Switch

ISOL


DOORS UNLOCKED INDIC

CASE 1

No aural warning occurs.

1. Fasten Belts Sign .........................................

2. Cabin Entrance Door ....................................

If the door is, or seems to be, unlocked:3. Cabin Differential Pressure.........................R


– – – END – – –

CABIN



CASE 2

Access to the baggage compartment is not permitte

1. Bag Access Light .........................................

– – – END – – –

CASE 3

Visually check for proper closing and latching of the

BAGACCESS

REARDOORS

28

APU BLEED LIGHT

This light illuminates if the APU bleed-air valve is nowhenever the bleed switch is off, or when one of moved beyond 54° of power lever movement andautomatically closed.

– – – END – – –

WINDSHIELD

BLEEDAPU



CRACK OR BUBBLES

1. Airspeed...........................................................2

2. Cabin Differential Pressure...............................

3. Associated Windshield Heat Switch ...............

– – – END – – –

HEAT SYSTEM INOPERATI

XFR



ANGLE-OF-ATTACK

ANGLE-OF-ATTACK STALL PROBE HEA

OR

Avoid icing conditions.

– – – END – – –

ANGLE-OF-ATTACK INDICATOR HEAT SYSTEM INOPERATI

AOAPROBE

R. AOAL. AOA



The angle-of-attack indicator system must not be used

– – – END – – –

FIRE PROTECTION

FAILURE OF ENGINE FIRE DETECTI

PROBE


EFIS MALFUNCTIO

NOTE

In the event of a failure of either CRT, the

be flown by the pilot who has both CRTs o

In the event of the failure of the EHSI CRso equipped with the multifunction displacontroller may be selected to the Hif required.

EITHER EADI CRT FAILU

The CRT display on the EADI will go blank, or the cdifficult to interpret.


1. EADI On/Off Dimmer Knob...........................

2 MFD C t ll (If I t ll d) (At Di ti )



2. MFD Controller (If Installed) (At Discretion)..ASSOCIA

– – – END – – –

EITHER EHSI CRT FAILU

The CRT display on the EHSI will go blank, or the cdifficult to interpret.


1. EHSI On/Off Dimmer Knob...........................

2. MFD Controller (If Installed) (At Discretion)..ASSOCIA

END

SUCCESSIVE FAILURE OEADI AND EHSI CRTS ON THE SA

Both displays go successively blank.

1. MFD Controller (At Discretion) .........................

– – – END – – –

LOSS OF ASCB CONTRO

flag appears on both EADIs.

1. MFD Controller (At discretion)..........................

Items that are lost:.

Cross-side data.

Glide slope localizer and radio altimeter a

X DATA



Glide-slope, localizer, and radio altimeter, aand heading comparison monitoring.

– – – END – – –

EFIS—INVALID ATTITUDE D

flag appears on EADI and loss of attitude refe


1. IRS Pushbutton ................................................

2 OR Annunciators on EADI and EHSI

X

I

X

I

IRS


EFIS-IRS ATTITUDE MISCOMPARE AWITH OR WITHOUT HEADING CO

Miscompare annunciators appear on both

IRS DATA INVALID message appears on

1. Standby Horizon...........................................

2. Faulty IRS.....................................................

On the EFIS reversion controller panel on the faulty3. IRS Pushbutton ............................................

4. OR Annunciators............................

White annunciation with 3 IRS in

Amber “color” 2 IRS installation

X

IRS

X

IRS

IRS



Amber color 2 IRS installation.

– – – END – – –

EFIS—LOC OR GS MISCOMPARE A

and/or comparison annunciation

for any altitude below 1,200 feet.

1. Faulty ILS .....................................................

2 Proper ILS Radio for Each Side

GSLOC

EFIS-IRS HEADING COMPARISON ANWITHOUT AN ATTITUDE COMPARISON A

comparison annunciation appears on both

1. Standby Compass and RMI .............................

2. Faulty IRS.........................................................

If the faulty IRS is in the NAV mode, but the position inon the reversion controller for the faulty side IRS:

3. IRS Pushbutton ................................................

4. OR Annunciation (Amber or White).....

White annunciation with 3 IRS installation.

Amber “color” 2 IRS installations.

If h f l id IRS i i h NAV d b h i i

X

IRS

X

IRS

IRS



If the faulty side IRS is in the NAV mode but the positiincorrect, on the mode select unit (MSU) of the faulty

3. Selector.............................................................

4. POS SENSORS Page......................................

5. FMS off Faulty Side IRS....................ENTER C – – – END – – –

EFIS-IAS/MACH DATA INVA

flag appears on EADI and loss IAS and MacIAS



S -2

NORM

EXT POWER

OIL

APU STOP

START

IRS

GEN

BATTERY 1

TOAMMETER

FIREEXTINGUISHERS

CONDITIONS:

GENERATORS 1, 2, AND 3 OPERATING.BOTH BATTERY SWITCHES ON.

BUS TIE OPEN, POWER SELECTORSWITCH IN NORMAL.

GEN



R ev i s i on 6

EXTINGUISHERS

COCKPITDOME LIGHT

TOAMMETER

BATTERY 2



S -4

G1

FILLERPORT

CENTERWING TANKS

NEGATIVEPRESSURERELIEFVALVE

PROBE

R1 CROSSFEEDVALVE


MANIFOLD


GROUP 1CROSSFEED

MANIFOLD

GROUP 2

SHV



R ev i s i on 5 NO.

GROUP 2FUEL

MANIFOLD

R ev i s i on4

FLASHING IF

CORRESPONDINGSWITCH OFF

ON IF:• APU BLEED VALVE OPEN IN FLI• APU BLEED VALVE OPEN ON TH

GROUND AND POWER LEVER A A POSITION ABOVE 54° FCU

WARNING PANEL

OVERHEAD PANEL

BLEEDOVHT

BLEEDAPU

ISOLHEAT

APU

NORM

PRV3PRV2

LPLP

ISOLATION

PASSENGER CREW

BLEED AIR

BAG

AUTOONOFF

HP1

ISOL



S - 5

NOTE: HP1, PRV2, AND PRV3 WILL AUTOMATICALLY

CLOSE FOR 18 SECONDS WHEN ANTI-ICING OPERATION AND HIGH POWER SETTINGS ARE SELECTED.

BLEED-AIR DISTRIBUT


AIRPLANE SYSTEMS

Fuel tank pressurization 1-2

Pressurization jet pump

Waste water system

No. 1 engine air intake and heatexchanger inlet anti-icing



LP

MAIN(3 ENGINES)

AUXILIARY(NO.1 AND

NO. 2 ENGINES)



Low power setting 1-2-3

High power setting 1-2-3

Winganti-icing

Low power setting 1-2-3


No. 2engine

S-ductanti-icing

Low power setting

Medium power setting 1-2-3Air

Conditioning


R ev i s i on 6

LP WATERSEPARATOR


HP WATERSEPARATION

CREW AIR-

CONDITIONINGCIRCUIT

CREWCOLD AIR

CABIN

COLD AIROR

SUCTION

COLD AIR

PRESSURE-REGULATING

VALVE

COLD AIROR

SUCTION

CABIN


TURBINE



S -7

LEGEND

BLEED AIR TURBCOMP

TEMPERED AIRTURBTURB

CONDITIONINGCIRCUIT

STATICAIR

INLETNO. 1

ENGINEPYLON RAM

AIR INLET WAINJEC

EXTINGUISHING SYSTCONTROLS AND OPERA


NORMAL

ENGINE 1 FIREFUELSHUTOFFSWITCH

FIREPANEL

210

210

FIRE 1

FAULT

DISCH

210

210

FIRE 2

FAULT

DISCH

ENGINE 2 FIRE

FIREPANEL

FIRST SHOT SECOND S

ENGINE

1

1 2 3 4

ENGINE

2

1 2 3 4



ENGINE3

NORMAL

1 2 3 4

FIRE 2

210

210

FIRE 3

FAULT

DISCH

ENGINE 3 FIRE

FIREPANEL

NORMAL

FIRST SHOT


FIRST SHOT (LEFT)

FIRST SHOT (RIGHT)

R ev i s i on 5 . 0 1

ADC 1

STANDBY STATIC

STANDBY PITOT

PILOT PITOT

COPILOT STATIC

PILOT STATIC

MIN

HOR STABCRUISE STOP

ID 802 A(TA

AILERON ARTHUR QMONITORING

A/C SNs UP TO 178

AILERON ARTHUR QCONTROL

SNs 179 AND HIGHER

A

SLATMONITORING

VERIN

SLATCONTACTS

TRIP(CABI

RAT

STASPE



S - 9

VMO /MMOWARNING

LANDING GEARCONTACTS

ALTITUDEWARNING

ATC 1 CONTROL PANEL

QUICK REFERENCE

FMS INITIALIZATION CHECKLIST

OTHER FMS PROGRAMMING

FUEL CONVERSION CHART

ELECTRICAL—KEY BUS ITEMS

DUAL FUNCTIONING CIRCUIT BREAKERS

MECHANICS PANEL

LANDING DISTANCE ADDITIONS



LANDING DISTANCE ADDITIONS

ONE ENGINE INOPERATIVE DRIFTDOWN CHART

FMS INITIALIZATION CHEC

1. Select avionics switches on.

2. NAV IDENT page comes into view.

3. Check date – UCT – active NDB.

4. Select the MAINTENANCE page (L-4).

5. Check configuration.

6. Select IRS switch to NAV.

7. Select NAV IDENT (R-4).

8. Select POS INIT (R-4).

9. Enter present position and load.

• By ramp position name, or


1



• By coordinates, if known, or

• By airport ident (four letters)

10. Select flight plan; (R-4) active flight plan com

11. Enter stored flight plan by name; then go to

12. Enter a new flight plan name for adding to th

13. Build the new flight plan.

14 S l t FPL SEL (R 4)

OTHER FMS PROGRAMMI

If a configuration change is desired:

1. Select the MAINTENANCE page.

2. Select CHANGE CONFIG (R-1).

3. Select the desired configuration.

If unable to change configuration:

4. Select the MAINTENANCE page (R-4).

5. Select CONFIG PROBLEMS (R-3).

6. Select DATA LOAD (L-4).

7. Select CUSTOM DB (L-2).

8. Select TO COPILOT (L-2) or FR COPILOT (L-3



9. Select YES (R-4) when loading is complete.

10. Select MAINTENANCE (R-4).

If data loading is desired:

1. Insert disc into data loader.

2. Select the ON button.

3. Select the desired CDU (LEFT-RIGHT-AUX)

4 S l t DATA LOAD (L 4) l t d CDU


2

° F

G a l l o n s

L i t e r s

W e i g h t

2 0 5 0

7 7 5 9

1 3 8 3 8

2 1 0 0

7 9 4 8

1 4 1 7 5

2 1 5 0

8 1 3 8

1 4 5 1 3

2 2 0 0

8 3 2 7

1 4 8 5 0

2 2 5 0

8 5 1 6

1 5 1 8 8

2 3 0 0

8 7 0 5

1 5 5 2 5

2 3 5 0

8 8 9 5

1 5 8 6 3

2 4 0 0

9 0 8 4

1 6 2 0 0

2 4 5 0

9 2 7 3

1 6 5 3 7

2 5 0 0

9 4 6 2

1 6 8 7 5

2 5 5 0

9 6 5 2

1 7 2 1 2

2 6 0 0

9 8 4 1

1 7 5 5 0

2 6 5 0

1 0 0 3 0

1 7 8 8 7



C O N V E R S I O

N C H A R T

E L - 6 . 7 5 P O U N

D S

A T 6 0 °

L i t e r s

W e i g h t

3 9 7 4

7 0 8 8

4 1 6 3

7 4 2 5

4 3 5 3

7 7 6 3

4 5 4 2

8 1 0 0

4 7 3 1

8 4 3 8

4 9 2 0

8 7 7 5

5 1 1 0

9 1 1 3

5 3 0 0

9 4 5 0

5 4 8 8

9 7 8 8

5 6 7 7

1 0 1 2 5

5 8 6 7

1 0 4 6 3

6 0 5 6

1 0 8 0 0

6 2 4 5

1 1 1 3 8

ELECTRICAL—KEY BUS ITE

BUS A1 BU

1. GEAR CNTL 1 EMERGEN

2. AIR BRAKE CNTL 2. NORMAL S

3. STAND-BY PUMP CONTL 3. CABIN CO4. ELEVATOR ARTHUR 4. #2 HYD SY

5. SLAT INDIC AND NORMAL 5. #2 ENG FIRCONTROL DETECTIO

6. TRIM INDIC FAULT LIG

7. EMERG STAB TRIM 6. EFIS 2

8. #2 BRAKES 7. #2 ENG AN9. LEFT PITOT HEAT

10. CREW COND’G VALVE

11. CABIN PRESSURIZATIONCONTROLLER

12. WING ANTI-ICE

13. GROUP 2 FUEL XFR

14. #1 HYD SYS INDIC

15 #1 ENG FIRE DETECTION



15. #1 ENG FIRE DETECTIONAND FAULT LIGHT

16. EFIS 1

17. #1 ENG ANTI-ICE

BUS A2 BU

1. ANTI-SKID 1. EMERG A2. AIL TRIM 2. FLAP CNT3. FLAP/AIR BK INDIC 3. NOSE WH4. STAND-BY PITOT HEAT 4. THRUST R5. M.F.D 5. FMS 26. FMS 1 6. FUELING

7 #3 ENG FIRE DETECTION 7 CO PILOT

MECHANICS PANEL

Light is on when Bag Iand the valve is closed

Light is on when HEAed and the heat valve


3

BAG 1 COMPTBLEED AIR

ISOL HEAT

STABILIZERRELAY TEST ANTI-SKID

COOLINGFAN

SG 3

MFD

TES

NOR

RESET A/P TEST

ON

OFF

FUEL FLOWTEST

GRAVITYFUELING

ENG INDICTEST

BAG COMP

BLEED AIR

ISOL

BAG COMP

BLEED AIR



ed and the heat valve

Test Socket — If voltaghas been removed froHOR-STAB motor.

1. When red, there is a

b h

BLEED AIR

HEAT

STABILIZERRELAY TEST

ANTI-SKID

When the switch is selectright vent valves should orefueling/defueling valve.proper vent valve positionfirmed by extinguishing oFUELING” light on the aft

When pushed, slews all N

tors to the blue dot on the

When pushed, resets all

When red, respective E.Foverheated.

GRAVITY

FUELING

ENGINE

INDICATION

TEST

E.F.I.S.

MAGNETIC

INDICATORS

RESET



LANDING DISTANCE ADDI

The following are additives to be made to the landingcomputed for a 40°flaps + slats landing under norm

NOTELD – Landing Distance

Emergency Checklist

Loss of Both Hydraulic Systems

Clean configuration, VREF + 30 knots. Twice the noLanding field length is not addressed.

Approach and Landing—Two Engines Inoperati

If using:

• 7° flaps + slats, fly VREF + 20 knots; add


After making the additive for flaps and slats above




After making the additive for flaps and slats above, system is used, add 50% to the calculated LD.

Abnormal Checklist

Approach and Landing-One Engine Inoperative

If using:



Loss of No. 1 Hydraulic System

fl b d l l fl V

4

Landing with Inoperative Stabilizer

Use 20°flaps + slats, and fly VREF + 20 knots; add

Landing with Inoperative Elevator

Use 40° flaps + slats, and fly VREF + 10 knots; add 1

Pitch Feel Light On (Authur Unit Failed in Heavy F

Fly VREF + 10 knots; add 800 feet to LD.

Flap Asymmetry

If using:

• Up to 7°flaps + slats, fly VREF 20 knots; add

• 7°to 20°flaps + slats, fly VREF + 15 knots; add

• 20 to 40°flaps + slats, fly VREF + 5 knots; add

Landing with Flaps or Slats Malfunctions



Landing with Flaps or Slats Malfunctions

If using:

• 40°flaps + outboard slats only, fly VREF + 5 knots

• Flaps + slats clean, fly VREF + 30 knots; a

Landing with Airbrake Malfunctions

If:

• Airbrakes are extended to position 1, fly VR


1 ENGINE OUT – DRIFT DOWN: TIME (min) – DISTANCE (NM) – FTEMPERATURE DAY ISA

WEIGHT (x 1,0

ALTITUDE (ft)

50,000

45,000

41,000 and

below

DRIFT DOWN

SPEED

(kt)

46 44 42 40 38 36 34

173

196 183 187 175 166 161

190 185 180 175 170 166 161

ALTITUDEAT ENGINE

FAILURE (ft)

WEIGHT (x 1,0

46 44 42 40 38 36 34

4

22

1,25

31,42

48 4226 21

1,317 1,22

30 070 31 41

TIME (min)

DISTANCE (NM)

FUEL USED (lb)

FINAL ALT. (ft)

TIME (min)

DISTANCE (NM)

FUEL USED (lb)

FINAL ALT. (ft)

TIME (min)DISTANCE (NM)

FUEL USED (lb)

FINAL ALT (ft)

49,000

47,000

45,000

ONE ENGINE INOPERATIVE DRIFT



5

30,070 31,41

52 50 49 47 4

246 237 229 218 20

1,634 1,518 1,416 1,296 1,19

26,170 27,440 28,740 30,060 31,39

52 50 49 47 45 4

247 238 229 218 207 191,721 1,603 1,498 1,377 1,258 1,16

24,920 26,150 27,410 28,720 30,040 31,38

52 50 49 47 45 43 4

247 237 229 219 208 196 18

1,812 1,685 1,574 1,456 1,340 1,215 1,10

23,700 24,900 26,140 27,390 28,700 30,020 31,35

50 48 47 45 43 40 3

237 227 218 207 196 183 17

1,769 1,644 1,535 1,408 1,290 1,158 1,04

FINAL ALT. (ft)

TIME (min)

DISTANCE (NM)

FUEL USED (lb)

FINAL ALT. (ft)

TIME (min)

DISTANCE (NM)FUEL USED (lb)

FINAL ALT. (ft)

TIME (min)

DISTANCE (NM)

FUEL USED (lb)

FINAL ALT. (ft)

TIME (min)

DISTANCE (NM)

FUEL USED (lb)

43,000

41,000

39,000

37,000

PILOT NOTES



Annunciators

Falcon 900 A/B

#2 P BK



Falcon 900 A/B

1. Steady: No. 2 Brake Pressure >2. Flashing: (< 1200 psi Park Brak

#2 P BK



Annunciators

Falcon 900 A/B

MOVING

MAIN GEAR RED LI



Annunciators

Falcon 900 A/B

MOVING

NOSE GEAR RED L



Falcon 900 A/B

MOVING

1. Nose gear position

VS

2. Gear control handleOR

If nose gear on

THEN

Nose gear door(s) not fu

▼



Annunciators

Falcon 900 A/B

OIL



Falcon 900 A/B

1. Pressure < 25 psi.2. Chip Detector.

OIL



Annunciators

Falcon 900 A/B

T/OCONFIG



Annunciators

Falcon 900 A/B

L/R WHLOVHT



Overheat detected in left or rig

Falcon 900 A/B

L/R WHLOVHT



Annunciators

Falcon 900 A/B

MISTRIM



Falcon 900 A/B

MISTRIM

See ID-802 Displa



Annunciators

Falcon 900 A/B

ECUOVHT



1. Environmental control unit tem> 230° C.

2. Turbofan bypass valve open (g

Falcon 900 A/B

ECUOVHT



Annunciators

Falcon 900 A/B

ST/BY

PUMP(HYDRAULIC PAN



Annunciators

Falcon 900 A/B

PUMP

(HYDRAULIC PAN



Hydraulic Pump Pressure> 2150 psi (Light Out)

< 1500 psi (Light On)

Falcon 900 A/B

PUMP



Annunciators

Falcon 900 A/B

ENG 2 FAIL



Falcon 900 A/B

GROUND

ONLY

GROUND

OR

FLIGHT

1. On ground.2. Power lever > 84°3. No. 2 N

1< 85%.

1. S duct door not la

ENG 2 FAIL



Annunciators

Falcon 900 A/B

FUELING



Falcon 900 A/B

PRESSURE

REFUELING

GRAVITY

AND

DEFUELING

1. Refueling control p2. Refueling connect3. Vent value lever no

4. Vent valves (2) not

5. Gravity fueling swi6. Defueling switch o7. REF/DEF valve op

8. B2 bus not powere

FUELING



Annunciators

Falcon 900 A/B

CMPTR



1. Insufficient voltage to compute2. Input failure.

3. Switch not in Auto.

4. Flashing (on ground only).

Falcon 900 A/B

CMPTR



Annunciators

Falcon 900 A/B

AILZERO



Falcon 900 A/B

AILZERO

Emergency aileron actuator not in



Annunciators

Falcon 900 A/B

AILFEEL



DISAGREEMENT

Aileron Arthur Q Pos

VS

Airspeed (ADC 1 or AD

Falcon 900 A/B

AILFEEL



Annunciators

Falcon 900 A/B

PITCHFEEL



DISAGREEMENTElevator Arthur Posit

VS

Horizontal Stabilize

Falcon 900 A/B

PITCHFEEL



Annunciators

Falcon 900 A/B

AUTOSLATS



1. Stall vanes.2. Proximity switches.

3. ADC 265 knot speed switche

Falcon 900 A/B

AUTOSLATS



Annunciators

Falcon 900 A/B

FLAP

ASYM



1. 3° – 5° split (0° – 20°)2. 5° – 8° split (20° – 40°).

Falcon 900 A/B

FLAPASYM



Annunciators

Falcon 900 A/B

GEAR CONFIGURATIO▼



Falcon 900 A/B

▼Respective gear is down



Annunciators

Falcon 900 A/B

FIRE

(FIRE PANEL)



Main engine fire/overheat

Falcon 900 A/B

FIRE



(FIRE PANEL)

Annunciators

Falcon 900 A/B

FIRE

BAG COMP



Fire or smoke detected in baggag

Falcon 900 A/B

FIREBAG COMP



Falcon 900 A/B

DISAGREEMENTEngine Fuel Shutoff V

VS

Fuel Shutoff Switch Po

TRANS



Annunciators

Falcon 900 A/B

AIRBRAKES



Annunciators

Falcon 900 A/B

FIRE APU

(FIRE PANEL)



APU fire or overheat de

Falcon 900 A/B

FIRE APU



Annunciators

Falcon 900 A/B

(FIRE PANEL)

FAULT



Fire detector loop fa

Falcon 900 A/B

FAULT



Annunciators

Falcon 900 A/B

BAT



The battery is not connected tmain bus.

Falcon 900 A/B

BAT



Annunciators

Falcon 900 A/B

▼

▼



Falcon 900 A/B

▼

▼ DISAGREE

Slat/Flap

VSSlats Po



Annunciators

Falcon 900 A/B

(SLAT/FLAP CONFIGURAT



1. Steady: All slats are extended

2. Flashing: Only the outboard s

Falcon 900 A/B



Annunciators

Falcon 900 A/B

GEN



1. RCR/Main Bus.2. Start Relay.

Falcon 900 A/B

GEN



Annunciators

Falcon 900 A/B

BUS TIED



1. Bus Tie Relay.

2. BAT 2 Contactor.

Falcon 900 A/B

BUS TIED



Annunciators

Falcon 900 A/B

FUEL



Falcon 900 A/B

Boost pump pressure < 4

FUEL



Annunciators

Falcon 900 A/B

LO

FUEL



Falcon 900 A/B

< 200 lbs in respective fuel

LOFUEL



Annunciators

Falcon 900 A/B

COND’G

OVHT



Passenger or crew baseboard tem

Falcon 900 A/B

COND’GOVHT



Annunciators

Falcon 900 A/B

NOSE

CONE OVHT



Nose cone temperature

Falcon 900 A/B

NOSECONE OVHT



Annunciators

Falcon 900 A/B

BLEED

APU



Falcon 900 A/B

DISAGRE

APU Bleed V

VS

APU Bleed O

Power Levers >

Bleed Valve

BLEED APU



Annunciators

Falcon 900 A/B

BAG

ISOL



The baggage compartment isolati

(1 or 2 indications – see maintena

Falcon 900 A/B

BAGISOL



Annunciators

Falcon 900 A/B

CABIN



Falcon 900 A/B

with “CABIN” audio

1. Cabin altitude > 10,000 feet.

without “CABIN” au

1. Main entry door not locked.

2. Forward lavatory service pane

CABIN

CABIN



Annunciators

Falcon 900 A/B

REAR

DOOR



Falcon 900 A/B

1. Baggage compartment door n

2. Rear compartment door not c

REARDOOR



Annunciators

Falcon 900 A/B

BLEED

OVHT



Falcon 900 A/B

LP/HP Temperature > 33

BLEEDOVHT



Annunciators

Falcon 900 A/B

MACH

TRIM



Falcon 900 A/B

Mach trim system is not e

MACHTRIM



Annunciators

Falcon 900 A/B

BAG

ACCESS



Falcon 900 A/B

The baggage access door

BAGACCESS



Annunciators

Falcon 900 A/B

L/R AOA



Falcon 900 A/B

Voltage to stall vanes i

L/R AOA



Annunciators

Falcon 900 A/B

L/R PITOT



Falcon 900 A/B

Voltage to supplied items

L/R PITOT



Annunciators

Falcon 900 A/B

STBY PITOT



Falcon 900 A/B

Voltage to standby pitot pro

STBY PITOT



Annunciators

Falcon 900 A/B

AOA

PROBE



Falcon 900 A/B

Voltage to angle-of-attack p

AOAPROBE



Annunciators

Falcon 900 A/B

HOT

BAT



Falcon 900 A/B

Battery 1 or 2 temperature > 1

HOTBAT



Annunciators

Falcon 900 A/B

AP



Falcon 900 A/B

Autopilot has disconnected (

AP



Annunciators

Falcon 900 A/B

REV

UNLOCK



Falcon 900 A/B

DISAGREEMENT

Thrust Reverse Lever s

and

Thrust Reverser Claw Locks

REVUNLOCK



SMOKE REMOV

PHASE ONE

Emergency

Falcon 900 A/B



Falcon 900 A/B

Phase One

1. Crew Oxygen Masks andSmoke Goggles ...............................

2. Microphone Selector ........................

3. “NO SMOKING” Sign ......................

Only If there are no flames in the cab

4. Passenger Oxygen Controller and

Passenger Masks ........................OV



Emergency

Falcon 900 A/B

LOSS OF BOT

HYDRAULIC SYSTPHASE ONE



Phase One

1. Auto-Pilot and Yaw Damper ............

2. Airspeed..............................260 KIAS

Falcon 900 A/B



Emergency

Falcon 900 A/B

INADVERTENT TH

REVERSER DEPLOIN FLIGHT



Falcon 900 A/B

Phase One

1. No. 2 Engine ....................................

2. Thrust Reverser NORM/STOW Swit

3. Airspeed ..........................................2



Emergency

Falcon 900 A/B

ALL ENGINES INOPE

PHASE ONE



Falcon 900 A/B

Phase One

1. Communications ...............................2. Establish the aircraft within the airsta

(See Abnormal Tab 3, page A-5.)

3. Reduce the electrical load to the lowe

(See Emergency Tab 6, page E-13.)4. Relight the engines using the airstart

(See Abnormal Tab 3, page A-8 and



Emergency

Falcon 900 A/B

ENGINE FIRE IN F

PHASE ONE



Falcon 900 A/B

Phase One

1. Power Lever......................................

2. FUEL SHUTOFF Switch .................

Light ................................

3. Airspeed ..........................................

4. Fire Extinguisher DISCH Switch .....

If fire warning persists:

5. Fire Extinguisher DISCH Switch .....

TRANS



Emergency

Falcon 900 A/B

AIR CONDITIONING

PHASE ONE



Falcon 900 A/B

Phase One

Smoke at Air Conditioning Outlets

1. Crew Oxygen Masks and

Smoke Goggles ..............................


3. “NO SMOKING” Sign ......................

4. Passenger Oxygen Controller .........

5. Passenger Masks ............................



Emergency

Falcon 900 A/B

ELECTRICAL SMOKE

PHASE ONE



Falcon 900 A/B

Phase One

Smoke and unusual odors

1. Crew Oxygen Masks andSmoke Goggles ..............................


3. “NO SMOKING” Sign........................Only if there are no flames in the

4. Passenger Oxygen Controllerand Passenger Masks ................OV



Emergency

Falcon 900 A/B

FIRE IN BAGGAGE COM

PHASE ONE



Falcon 900 A/B

Phase One

1. BAG Switch...................................................................

In flight below 41,000 ft and with the baggage compartme

NOTE

This procedure is different for airplanes registered in the Uin the U.K., the baggage compartment is not accessible

2. The co-pilot dons the smoke hood and fights the fire w

The lavatory door must be properly closed prior to openindoor to prevent smoke from entering the passenger cabin

During takeoff or landing, or in flight above 41,000 ft or wnot pressurized:

1. BAG Switch ..................................................................

Light ...............................................................

2. BAG COMP Extinguisher Switch .................................

Do not open the baggage compartment door until the

BAG ISOL

CAUTION



Maximum Tire Rotatio

Limitations

Falcon 900 A/B



Falcon 900 A/B

Maximum Tire Rotatio

195 KIAS (225 mph



CG Limits

Limitations

Falcon 900 A/B



Falcon 900 A/B

CG Limits

14% MAC Forwa

31% MAC Aft



DATUM

Limitations

Falcon 900 A/B



Falcon 900 A/B

DATUM

25% MAC (FS 420 in

0% MAC (FS 392 inc



Maximum Speed Total Hy

Limitations

Falcon 900 A/B



Falcon 900 A/B

Maximum Speed Total Hy

260 KIAS/.76 MAC



Demonstrated Cros

Limitations

Falcon 900 A/B



Falcon 900 A/B

Demonstrated Cros

30 Knots



Maximum Tailwi

Limitations

Falcon 900 A/B



Falcon 900 A/B

Maximum Tailwi

10 Knots



AIRSPEED

MMO:

1. 25,000 Ft to 33,000 Ft: ?

2. 33,000 Ft to 37,000 Ft: ?3. Above 37,000 Ft: ?

Limitations

Falcon 900 A/B



Falcon 900 A/B

AIRSPEED

MMO:

1. 25,000 Ft to 33,000 Ft: .87 Ma

2. 33,000 Ft to 37,000 Ft: .87 Ma3. Above 37,000 Ft: .84 Ma



AIRSPEEDTurbulent Air Penetratio

Limitations

Falcon 900 A/B



Falcon 900 A/B

AIRSPEEDTurbulent Air Penetratio

280 KT/.76 M



AIRSPEEDVMO:

Limitations

Falcon 900 A/B



AIRSPEEDVMO:

350 KIAS increasing wit370 KIAS at 10,000

Falcon 900 A/B

Li it ti



AUTOPILOT

1. Minimum height during FMS appro

2. Minimum height radio altimeter ope

3. Minimum height radio altimeter ino

4. Minimum decision height?

5. Minimum height except during app

Limitations

Falcon 900 A/B



AUTOPILOT

1. Minimum height during FMS appro

2. Minimum height radio altimeter ope

3. Minimum height radio altimeter ino

4. Minimum decision height?

5. Minimum height except during app

Falcon 900 A/B

Limitations



STALL SPEED

Limitations

Falcon 900 A/B



Do not intentionally fly slowestall warning onset.

Falcon 900 A/B

Limitations



MINIMUM CONTROLVMCA:

Limitations

Falcon 900 A/B



MINIMUM CONTROLVMCA:

85.5 KCAS

Falcon 900 A/B

Limitations



Windshield Wiper Opera

Direct Vision Window Ope

Limitations

Falcon 900 A/B



Windshield Wiper Opera

Direct Vision Window Ope

215 KIAS

Falcon 900 A/B

Limitations



HIGH LIFT DEVICES OP

OR EXTENDED LIMIT S

Limitations

Falcon 900 A/B



HIGH LIFT DEVICES OP

OR EXTENDED LIMIT S

Slats + Flaps 7°: 2

Slats + Flaps 20°:1

Slats + Flaps 40°: 1

Falcon 900 A/B

Limitations



MANEUVERING SP

Limitations

Falcon 900 A/B



MANEUVERING SP

228 KIAS

Falcon 900 A/B

Limitations



MANEUVERING FLIGHT LOA

Limitations

Falcon 900 A/B



MANEUVERING FLIGHT LOA

Flaps Up: +2.5

Flaps Down: +2 t

Falcon 900 A/B

Limitations



Maximum Landing Gear Ope(VLO)

Maximum Landing Gear Exte(VLE)

Falcon 900 A/B



VLO: 190 KIAMLO: 0.70

VLE: 245 KIA

MLE: 0.75

Falcon 900 A/B



MAXIMUM RUNWAY

+/– 2.5%

Falcon 900 A/B

Limitations



Minimum Flight W

Falcon 900 A/B



Minimum Flight W

20,700 LB

Falcon 900 A/B

Limitations



Baggage Compartment Weigh

Falcon 900 A/B



Baggage Compartment Weigh

2,866 LB

41,000 ft baggage acceclosed and latche

Falcon 900 A/B

Limitations



Maximum Zero Fuel

Falcon 900 A/B



Maximum Zero Fuel

30,870 LB

Falcon 900 A/B

Limitations



Maximum Landing W

Falcon 900 A/B



Maximum Landing W

42,000 LB

Falcon 900 A/B

Limitations



Maximum Take-off W

Falcon 900 A/B



Maximum Take-off W

46,500 LB

Falcon 900 A/B

Limitations



Maximum Ramp W

Falcon 900 A/B



Maximum Ramp W

46,700 LB

Falcon 900 A/B

Limitations



ALTITUDEMaximum Take-off and

Falcon 900 A/B



ALTITUDEMaximum Take-off and

–1,000 to +14,000 F

Falcon 900 A/B

Limitations



ALTITUDEMaximum Operati

Falcon 900 A/B



Limitations



ENGINE OPERATIO

TFE-731-5BR-1C

Falcon 900 A/B

STARTING TIME

Ground Start and Starter Assist Airstart10% N2 to light off

Windmilling AirstartWindmilling N2 to 60% N2

Ground StartLight off to idle



Falcon 900 A/B

ENGINE OPERATIO

TFE-731-5BR-1CSTARTING TIME

Ground Start and Starter Assist Airstart10% N2 to light off

Windmilling Airstartwindmilling N2 to 60% N2

Ground StartLight off to idle

Limitations



TFE-731-5BR ENG

Engine Starter Motor-

Falcon 900 A/B

First Attempt:

40 seconds ON 3 minutes OFF



40 seconds ON 3 minutes OFF(If required after 1 minute off motor 15 seconds

Second Attempt:

30 seconds ON 3 minutes OFF(If required after 1 minute off motor 15 seconds

Third Attempt:

30 seconds ON then 30 minutes shutdownNOTE

In temperatures > 30° C time b2 successive starts must be d

Falcon 900 A/B

Limitations



Falcon 900 A/B

MAXIMUM INTERSTAGE TURBINE

ITTTFE-731-5BR-1C

Starting:

Take-off:Without Increased Thrust (RPR):With Increased Thrust (RPR):

Maximum Continuous:

MAXIMUM INTERSTAGE TURBINE

ITT



ITT

TFE-731-5BR-1CStarting:

Take-off:Without Increased Thrust (RPR):

Normal (5 minutes max.)Transient (5 seconds max.)

With Increased Thrust (RPR):Normal (5 minutes max.)Transient (5 seconds max.)

Maximum Continuous:

Falcon 900 A/B

Limitations



Falcon 900 A/B

OIL PRESSURE

MINIMUM THRUST SETTING PRESSURE

Idle ?

Takeoff ormaximum continuous ?

Transient



Limitations



Falcon 900 A/B

OIL TEMPERATU

Sea level to 30,000 ft ?

Above 30,000 ft ?

Transient all altitudes ?le

Minimum for exceeding idle power ?

OIL TEMPERATU



Sea level to 30,000 ft 1

Above 30,000 ft 1

Transient all altitudes 1

le

Minimum for exceeding idle power 30

Falcon 900 A/B

Limitations

MAXIMUM ENGINE ROTO



Falcon 900 A/B

MAXIMUM ENGINE ROTO

N

Take-off – Maximum Continuous

Transient (5 seconds max.)



Limitations



Falcon 900 A/B

TFE-731-5BR-1C

THRUST RATING

• Takeoff ?

• Maximum Continuous ?



TFE-731-5BR-1C

THRUST RATING

• Takeoff 4

• Maximum Continuous 4

Falcon 900 A/B



BATTERY TEMPERA



Before SB 94(Warm) Amber Light 120(Hot) Red Light 150

After SB 94(Warm) Amber Light 120(Hot) Red Light 160

Falcon 900 A/B

Limitations



CABIN PRESSURIZ

Falcon 900 A/B

CABIN PRESSURIZ



CABIN PRESSURIZ

Maximum Differential P

9.6 psi

Falcon 900 A/B

AUXILIARY POWER

Limitations



AUXILIARY POWERGTCP 36-150 (F

Maximum N1 Speed: ?

Starting T5 Limit: ?

Stabilized ?

Falcon 900 A/B



ELECTRICAL

Limitations



ELECTRICAL

Maximum Voltage ?

Maximum Amperage

Transient (1 minute max) ?Up to 43,000 Ft. ?

Above 43,000 Ft. ?

Falcon 900 A/B

ELECTRICAL



Maximum Voltage 3Maximum Amperage

Transient (1 minute max) 3

Up to 43,000 Ft. 3Above 43,000 Ft. 2

Falcon 900 A/B

APU

Limitations



APU• Ground use only• Operation of APU with passengers in the cab

monitoring is not authorized

1. Maximum N1 Speed

2. Starting EGT Temp

3. Stabilized EGT

4. Maximum Generator Output

a. Transient (1 minute max)b. Stabilized

Falcon 900 A/B

APU• Ground use only



• Operation of APU with passengers in the cabmonitoring is not authorized

1. Maximum N1 Speed

2. Starting EGT Tempma

3. Stabilized EGT4. Maximum Generator Output

a. Transient (1 minute max)b. Stabilized

Falcon 900 A/B

Limitations



LIMITATION ON NOSE GE

LIMITATION ON AIRBR

Falcon 900 A/B

Limitation On Nose Ge



Limitation On Nose Ge

Must be chined

Limitation On Airbra

Must not be actuated within 300 ft AG

Falcon 900 A/B

ENGINE ANTI-IC

Limitations



ENGINE ANTI IC

1. Must not be used with total airtemperature above ?

WING ANTI-ICE2. Must not be used with total air

temperature above ?

Falcon 900 A/B

ENGINE ANTI-IC



1. Must not be used with total airtemperature above 1

WING ANTI-ICE

2. Must not be used with total airtemperature above 1

Falcon 900 A/B

Limitations



WINDSHIELD CRACOR BUBBLES

Falcon 900 A/B

WINDSHIELD CRAC



Falcon 900 A/B

OR BUBBLES

Airspeed: Max

Cabin Differential Pressure: Max