flight safety learjet 20 series pilot training manual volume 2

FlightSafetyinternational

LEARJET 20 SERIESPILOT TRAINING

MANUALVOLUME 2

AIRCRAFT SYSTEMSSECOND EDITION

FlightSafety International, Inc.Marine Air Terminal, LaGuardia Airport

Flushing, New York 11371(718) 565-4100

www.flightsafety.com

Courses for the Learjet 20 and other Learjet aircraft are taught at the followingFlightSafety learning centers:

Wichita Learning CenterTwo Learjet WayWichita, KS 67209(800) 491-9807Fax: (316) 943-0314

Tucson Learning Center1071 E. Aero Park Blvd.Tucson, AZ 85706(800) 203-5627Fax: (520) 918-7111

Copyright © 2005 by FlightSafety International, Inc.All rights reserved.

Printed in the United States of America.

INSERT LATEST REVISED PAGES, DESTROY SUPERSEDED PAGES

LIST OF EFFECTIVE PAGES

Dates of issue for original and changed pages are:

Second Edition ....... 0 .......... April 2005Revision............... .01........... July 2005

NOTE:For printing purposes, revision numbers in footers occur at the bottom of every pagethat has changed in any way (grammatical or typographical revisions, reflow of pages,and other changes that do not necessarily affect the meaning of the manual).

THIS PUBLICATION CONSISTS OF THE FOLLOWING:

*Zero in this column indicates an original page.

LEARJET 20 SERIES PILOT TRAINING MANUAL


LEP-1Revision .01 FOR TRAINING PURPOSES ONLY

Page *RevisionNo. No.

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Page *RevisionNo. No.

NOTICE

The material contained in this training manual is based on information obtained from theaircraft manufacturer’s pilot manuals and maintenance manuals. It is to be used for famil-iarization and training purposes only.

At the time of printing it contained then-current information. In the event of conflict betweendata provided herein and that in publications issued by the manufacturer or the FAA, thatof the manufacturer or the FAA shall take precedence.

We at FlightSafety want you to have the best training possible. We welcome any sugges-tions you might have for improving this manual or any other aspect of our training program.

FOR TRAINING PURPOSES ONLY


CONTENTS

Chapter 1 AIRCRAFT GENERAL

Chapter 2 ELECTRICAL POWER SYSTEMS

Chapter 2A ELECTRICAL POWER SYSTEMS

Chapter 3 LIGHTING

Chapter 4 MASTER WARNING SYSTEM

Chapter 5 FUEL SYSTEM

Chapter 5A FUEL SYSTEM

Chapter 6 AUXILIARY POWER UNIT

Chapter 7 POWERPLANT

Chapter 8 FIRE PROTECTION

Chapter 9 PNEUMATICS

Chapter 10 ICE AND RAIN PROTECTION

Chapter 10A ICE AND RAIN PROTECTION

Chapter 11 AIR CONDITIONING

Chapter 12 PRESSURIZATION

Chapter 12A PRESSURIZATION

Chapter 13 HYDRAULIC POWER SYSTEMS

Chapter 14 LANDING GEAR AND BRAKES

Chapter 15 FLIGHT CONTROLS

Chapter 16 AVIONICS

Chapter 17 MISCELLANEOUS SYSTEMS

WALKAROUND

APPENDIX

ANNUNCIATOR PANEL

INSTRUMENT PANEL POSTER

1-i

CHAPTER 1AIRCRAFT GENERAL

CONTENTS

Page

INTRODUCTION ................................................................................................................... 1-1

GENERAL............................................................................................................................... 1-1

STRUCTURES........................................................................................................................ 1-2

General ............................................................................................................................. 1-2

Fuselage ........................................................................................................................... 1-7

Wing............................................................................................................................... 1-15

Empennage..................................................................................................................... 1-15

AIRPLANE SYSTEMS ........................................................................................................ 1-16

Electrical Power systems ............................................................................................... 1-16

Lighting.......................................................................................................................... 1-16

Fuel System.................................................................................................................... 1-16

Powerplant ..................................................................................................................... 1-16

Ice and Rain protection .................................................................................................. 1-17

Air Conditioning and Pressurization.............................................................................. 1-17

Hydraulic Power Systems .............................................................................................. 1-17

Landing Gear and Brakes .............................................................................................. 1-17

Flight Controls ............................................................................................................... 1-17

Pitot-Static System......................................................................................................... 1-17

Oxygen System.............................................................................................................. 1-18

Static or Lightning Strike Protection ............................................................................. 1-18




1-iii

ILLUSTRATIONS

Figure Title Page

1-1 Learjet 25.................................................................................................................. 1-2

1-2 Airplane Dimensions................................................................................................ 1-3

1-3 Airplane Turning Radii............................................................................................. 1-4

1-4 Danger Areas ............................................................................................................ 1-5

1-5 Fuselage Sections ..................................................................................................... 1-6

1-6 Nose Section............................................................................................................. 1-7

1-7 Cockpit Layout (Typical) ......................................................................................... 1-8

1-8. Cabin Interior (Typical) ............................................................................................ 1-9

1-9 Passenger-Crew Door ............................................................................................... 1-9

1-10 Door Locking Pin Hole .......................................................................................... 1-10

1-11 Emergency Exit ...................................................................................................... 1-11

1-12 Windshield (Typical) .............................................................................................. 1-12

1-13 Passenger Windows (Typical) ................................................................................ 1-13

1-14 Tailcone Entry Door ............................................................................................... 1-13

1-15 Wing ........................................................................................................................ 1-14

1-16 Empennage............................................................................................................. 1-15

1-17 Static Wicks (Typical) ............................................................................................ 1-18

TABLE

Table Title Page

1-1 Learjet 20 Series Models and Serialization.............................................................. 1-2




INTRODUCTIONThis training manual provides a description of the major airframe and engine systemsinstalled in the Learjet 20 series airplanes. The information contained herein is intendedonly as an instructional aid. This material does not supersede, nor is it meant to substi-tute for, any of the manufacturer’s maintenance or operating manuals. The material pre-sented has been prepared from the basic design data. All subsequent changes in airplaneappearance or system operation will be covered during academic training and subsequentrevisions to this manual.

Chapter 1 covers the structural makeup of the airplane and gives an overview of the sys-tems. The “Walkaround” section also contains a pictorial walkaround of the airplane.

The “Annunciator Panel” section (located at the back of Volume 2) displays all light in-dications and should be folded out for reference while reading this manual.

GENERALThe Learjet 20 series airplanes are all-metal,low-wing, twin-engine jet airplanes with re-tractable landing gear. They are powered byGeneral Electric CJ610 series turbojet en-gines rated at either 2,850 or 2,950 pounds of

thrust. The airplanes are pressurized and seat8 to 10 people including the crew.

The 20 series airplanes are certificated under FARPart 25 (except Model 23 which is certificated

CHAPTER 1AIRCRAFT GENERAL

1-1FOR TRAINING PURPOSES ONLY



under CAR 3) as two-pilot transport categoryaircraft. They are certified for operation to analtitude of 41,000, 45,000, or 51,000 feet, de-pending on serial number and equipment in-stallation. Figure 1-1 shows a Learjet 25.

The terms “early” and “late” are frequentlyused in the chapters of this training manualwith reference to airplane models. This is tosimplify explanations for the two basic groupsof airplanes. Table 1-1 lists the early- andlate-model groups by airplane serial numberfor reference.

STRUCTURES

GENERALMost of the airplane structures are fabricated ofhigh-strength aluminum alloy, with steel, glassfiber, and other materials used as needed. Duringtesting, all load-bearing members and surfacesdemonstrated the capability to carry 90% of the“G” forces with an adjacent structural compo-nent failed. The airplane structure consists offuselage, wings, and empennage. The discussionon the fuselage includes the doors and windows.General dimensions are shown in Figure 1-2.

1-2 FOR TRAINING PURPOSES ONLY



Figure 1-1. Learjet 25

Table 1-1. LEARJET 20 SERIES MODELS AND SERIALIZATION

MODEL MODELSERIAL BLOCK SERIAL BLOCK

LATEEARLY

23

24

24B

25

24D

24E

24F

25B

25C

25D

23-003 THROUGH 23-099

24-100 THROUGH 24-180

24-181 THROUGH 24-229 EXCEPT 24-218

25-002 THROUGH 25-064 EXCEPT 25-061

24-218, 24-230 THROUGH 24-328

24-329 AND SUBSEQUENT


25-061, 25-070 THROUGH 25-205

25-061, 25-070 THROUGH 25-205





41 FT

43 FT 3 IN.

14 FT 8 IN.

8 FT 3 IN.

34 FT 1 IN.

35 FT 7 IN.

45 FT

47 FT 7 IN.

12 FT3 IN.

12 FT3 IN.

MODELS 23 AND 24

ALL MODELS

MODEL 25

Figure 1-2. Airplane Dimensions




A

B

C

MODELS23 AND 24

MODEL25

A 27 FT 10 IN. 31 FT 3 IN.

B 34 FT 4 IN. 37 FT 6 IN.

C 31 FT 11 IN. 35 FT 1 IN.

NOTE: THE VALUES SHOWN REFLECT NORMAL NOSEWHEEL STEERING. SHORTER TURNS CAN BE EFFECTED WITH THE USE OF A TOW BAR.

Figure 1-3. Airplane Turning Radii

Figures 1-3 and 1-4 show the airplane turningradii and the danger areas from the weatherradar and engine intakes and exhausts.


45°

30 FT

30 FT

WEATHERRADAROPERATING

12 FT

12 FT

600°F

600°F 150°F

40 FT

40 FT150°F

WEATHER RADAR OPERATING

EXHAUST DANGER AREAS SHOWNFOR IDLE RPM. VALUES APPROXIMATELYDOUBLE FOR TAKEOFF RPM.

Figure 1-4. Danger Areas






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FUSELAGE

GeneralThe fuselage is constructed of stressed all-me t a l sk in w i th t r an sve r se f r ames andstringers. It employs the area rule design to re-duce aerodynamic drag and has four basic sec-tions (Figure 1-5). They are as follow:

• The nose section extends from the radomeaft to the forward pressure bulkhead.

• The pressurization section, which in-cludes the cockpit and passenger areas,extends aft to the rear pressure bulkhead.

• The fuselage fuel section starts just aftof the rear pressure bulkhead and ex-tends to the tailcone.

• The tailcone section extends aft of thefuel section.

The fuselage incorporates attachments forthe wings, tail group, engine support pylons,and the nose landing gear. In addition to thepressurized cockpit and passenger compart-ments, the fuselage includes the nose wheelwell, an unpressurized nose compartment, anda tailcone compartment used for equipmentinstallation.

Nose SectionThe nose of the fuselage (Figure 1-6) is formedby the radome. Aft of the radome is the nosecompartment.

The nose compartment access panels are on topof the fuselage forward of the windshield. Thepanels must be removed for access to variouselectronic components, oxygen bottle (wheninstalled in the nose), emergency air bottle, andthe alcohol anti-icing reservoir.


Figure 1-6. Nose Section



Pressurized SectionGeneralThe pressurized center section includes thecockpit and passenger cabin. Interior arrange-ments vary with airplane model and customerpreference. Typically, the cockpit is separatedfrom the cabin by a partition. A typical instru-ment panel is shown in Figure 1-7.

Seating ArrangementA typical seating arrangement in a Learjet istwo crew seats, four reclinable adjustable pas-senger seats, and a three-passenger seat at therear of the cabin. Each individual passengerseat is equipped with an adjustable headrestand two armrests. The outboard armrest isfixed, and the aisle armrest is of a swing-awaydesign for ease of entry and exit. An ashtrayand a reading light are also standard for eachpassenger seat. The standard seats are trackmounted to allow movement fore and aft and

are reversible for a foursome arrangement.Optional lateral tracking seats that also moveinboard are available. Safety belts are in-stalled on all seats. Figure 1-8 depicts a typ-ical cabin interior arrangement.

Safety belts and shoulder harnesses with self-adjusting inertia reels are installed on the pilotand copilot seats. The inertia reels allow theshoulder harness to extend or retract during nor-mal movement; however, the strap locks se-curely in place under sharp forward force. Tocheck the function of the reel, tug sharply onthe strap. The reel should lock under this testand prevent the strap from extending.

Immediately forward of the aft pressure bulk-head is the baggage compartment. It is limitedto 500 pounds in all models. Optional foldingtables, storage cabinets, and refreshment cen-ters are available, as well as toilet facilities.




Figure 1-7. Cockpit Layout (Typical)

Passenger-Crew DoorThe primary entrance and exit for passengersand crewmembers is through the clamshelldoor located on the left side of the forwardfuselage (Figure 1-9). The standard entrance




Figure 1-8. Cabin Interior (Typical)

door is 36 inches wide, but an optional 24-inchdoor is available. The upper door serves as anemergency exit, and the lower door has integralentrance steps.

Figure 1-9. Passenger-Crew Door

The upper portion of the door has both outsideand inside locking handles connected to acommon shaft through the door. Rotating ei-ther of these handles to the closed positiondrives six locking pins into holes in the fuse-lage frame (Figure 1-10) (three pins forwardand three aft) and two pins through interlock-ing arms that secure the two door halves to-gether.

The lower door has a single locking handle onthe inside. Rotating the lower door handle tothe closed (forward) position drives two pinsinto holes in the fuselage frame (one forwardand one aft). There are a total of 10 lockingpins on the two door sections.

To facilitate engagement of the upper doorlocking pins during closing, an electric actu-ator motor, torque tube assembly, and one ortwo hooks are installed in the lower door whichengage rollers installed on the upper door. Theactuator motor is operated from inside the air-plane by a toggle switch on the lower door, andfrom the outside by a key switch. Excludingthose airplanes subsequent of SNs 24-294 and25-170 and those with replacement motors,should the motor fail, the hook(s) can still beoperated manually from inside. Access is pro-vided to the torque-tube mechanism through apanel in the lower door, and a ratchet handleprovided in the airplane tool kit can be used tooperate the torque tube manually.

NOTEOne hook is used on 24-inch doors,while two hooks are used on 36-inch doors.

A secondary safety latch is installed on thelower door and is separate from the door lock-ing system. It consists of a notched pawl at-tached to the door. The pawl engages a notchedstriker plate attached to the frame when thedoor is closed. This engagement holds the lowerdoor closed while the locking handle is beingpositioned to the locked position. Additionally,it prevents the door from falling open as soonas the door handle is opened. The latch is re-leased by depressing the pawl.

When closing the doors from the inside, closeand latch the lower door first. Then, close theupper door and actuate the door motor switchto the closed position. This engages the hook(s)over roller(s) in the upper door, and cinchesthe upper door down tight while allowing thelocking pins to line up properly and meet themicroswitches as the upper door handle is ro-tated to the closed position. The DOOR lightwill remain illuminated until the hook(s) arebacked away from the upper door rollers by re-verse operation of the door motor switch.

When the door handles are in the closed posi-tion, the pins all contact microswitches. If anyof the switches is not actuated, a red DOORlight illuminates on the annunciator panel (see“Annunciator Panel” section). If the light illu-minates while the door is closed, the pilot canvisually check through inspection ports forproper alignment between the white lines onthe latch pins and on the door structure. Thetwo latch pins which connect the upper andlower doors are visible through the upholsterygap at the interface and do not have white lines.

Cables and hydraulic dampers are provided tostabilize the lower door when lowering it andwhen using it as a step. The 24-inch door hasone cable and a hydraulic damper. The 36-inch door has two cables and may have an op-t ional hydraul ic damper. The cables areconnected to takeup reels in the lower door and




Figure 1-10. Door Locking Pin Hole

are also used to pull the door closed from in-side the airplane.

The key switch is used to secure the door fromthe outside. By inserting a key into the switchand turning it in one direction, the actuatormotor drives the hooks to engage the upperdoor. Turning it in the other direction releasesthe hooks to permit opening the door.

NOTEAnytime the airplane is occupiedwith the entry doors locked, thehooks must be released. This per-mits opening the upper door foremergency egress.

The red DOOR light illuminated means:

• Any one of the 10 latch pins is not en-gaged with its respective microswitch.

• The hooks are not disengaged from thedoor and fully retracted.

• The door may be unsafe for takeoff.

A hollow neoprene seal surrounds the door-frame; the seal has holes to allow the entry ofpressurized cabin air, forming a positive sealaround the door.

Emergency ExitA hatch near the right rear of the cabin servesas an emergency exit on all Learjet 20 models(Figure 1-11).

The hatch can be opened from the inside on allmodels by a latch handle located at the top cen-ter of the window. The latch must be pulled in-ward to unlock; a continued inward pullreleases the hatch from the lower retainers.




Figure 1-11. Emergency Exit

The method of opening the hatch from the ex-terior varies with airplane model:

• On 23 models, exterior opening requiresthat the top center screw be removedand punched out. The hatch then can bepushed inside.

• On 24, 24B, and 25 models, the exteriorlatch must be raised up, then pushed into release the hatch. The hatch then canbe pushed inside.

• On 24D, E, F, and 25B, C, and D mod-els, exterior opening is accomplishedby depressing a PUSH button above thewindow, which releases a handle. Thehandle must then be turned in the direc-tion of the arrow stamped on the handle.The hatch then can be pushed inside.

Windows

Windshield—23 Model

The windshield is divided into two sections—the pilot’s and copilot’s halves—and consistsof two panes of acrylic plastic (Figure 1-12).The outer pane is .460 inches thick and theinner pane is .188 inches thick, with an air sep-aration between the panes. A dehydrator lo-cated under the copilot’s seat removes moisturefrom conditioned cabin air before it is routedbetween the panes. The dehydrator and plasticline must be kept in proper operating conditionto prevent moisture, dust, or smoke from col-lecting on the inner surfaces of the panes.

Windshield—All 24 and 25 Models

The windshield is divided into two sections—the pilot’s and copilot’s halves—and consistsof three laminated layers of acrylic plastic. Thewindshield is approximately one-inch thick.It is impact resistant and was tested againstfour-pound bird strikes at 350 knots.

Passenger Windows—23, 24, 24BModels

Three large, dual-pane, acrylic plastic win-dows are installed in the airplane. The right rearwindow serves as an emergency exit. The panesare held apart and sealed airtight by a spacer.

Passenger Windows—All Other Models

Six to nine small, dual-pane, acrylic plasticwindows are installed in the airplane, includingthe emergency exit hatch window (Figure 1-13).

Cleaning of Windows

The following precautions should be takenwhen cleaning airplane windows:

• Remove loosely adhering dirt and grit fromthe window by flushing with clean water.

• Wash with nonabrasive soap and water.A soft, thoroughly clean cloth, sponge,or chamois may be used in washing, butonly as a means of carrying the soapywater to the plastic. Go over the surfaceonly with the bare hand so that any abra-sive can be quickly detected and re-moved before it scratches the plasticsurface.

• Remove oil and grease by rubbing lightlywith a cloth wetted with aliphatic naptha.

• Dry the surface with a clean, damp cham-ois. A clean, soft cloth or tissue may beused if care is taken not to rub the plas-tic after it is dry.




Figure 1-12. Windshield (Typical)

NOTE• All rubbing operations on acrylic

plastics shall be done with as lighta pressure as possible.

• Rubbing the plastic surface with adry cloth will cause scratches andbuild up an electrostatic chargewhich attracts dust particles.

• Do not use the following materialson acrylic plastics: gasoline, alco-hol, benzene, hexane, xylene, ace-tone, carbon tetrachloride, fireextinguisher or deicing fluids, lac-quer thinners, or window cleaningsprays because they soften the plas-tic and/or cause crazing.

Fuel SectionThe fuselage fuel section is located immedi-ately aft of the rear pressure bulkhead. (SeeFigure 1-5.)

The fuel section is located as follows:

• Frames 22 to 24: Models 23, 24, 24B, D,F, and 24E, when installed

• Frames 22 to 25: Models 25, 25B, and D

• Frames 18 to 25: Model 25C, which haslong-range tanks installed

Tailcone SectionThe tailcone section extends aft from the fuelsection to the empennage. The tailcone entrydoor (Figure 1-14) is located at the bottom ofthis section. On some airplanes, the door ishinged at the forward edge and drops downwhen released by quick-release thumb latches,allowing access to the batteries, electricalcomponents, fuel filters, refrigeration equip-ment, engine fire extinguishers, and hydrauliccomponents. On airplanes without hinges, thedoor is secured by camlock fasteners.

There is no cockpit indicator to warn the pilotif the door is open.




Figure 1-13. Passenger Windows (Typical)

Figure 1-14. Tailcone Entry Door




Figure 1-15. Wing

WINGThe Learjet 20 has a sweptback, cantilevered,all-metal wing (Figure 1-15) which is mountedto the lower fuselage. Most of the wing issealed to form an integral fuel tank.

Eight fittings attaching the wings to the fuse-lage are designed to prevent wing deflectionsfrom inducing secondary loads in the pressur-ized fuselage. Ailerons are attached to the rearspar at three hinge points. The single-slottedflaps are attached to the inboard rear spar bytracks, rollers, and hinges. The spoilers are at-tached to the top of the wing surface by twohinges just forward of the flaps. The tip tanksare secured to the wing by two attach points.

The Learjet 20 wing is fitted with either vor-tex generators or boundary layer energizers.Whichever is used, their function is to delayairflow separation over the ailerons at highMach numbers.

Model 23 airplanes have two rows of vortexgenerators bonded to the upper and lower wingsurface forward of the ailerons. Each of thefour rows has 26 vortex generators. A maxi-mum of three vortex generators on each wingare allowed to be missing before flight.

All 24 and 25 models (except those withboundary layer energizers) have two rows ofvortex generators bonded to the upper wingsurface only. The forward row has 26 vortexgenerators installed; the rear row has 19. Amaximum of three vortex generators on eachwing are allowed to be missing before flight.

On a i rp lanes incorpora t ing AMK 83-4 ,Softflite I, and AMK 83-5 Softflite wing mod-ifications, the vortex generators have been re-placed with the following:

• Two rows of boundary layer energizers(BLEs) on each wing which perform thesame function as vortex generators butare more efficient. If any are missing,MMO is reduced to 0.78 MI.

• A full-chord fence on each wing, in-board of the aileron, which reduces span-wise flow of air.

• A stall strip, affixed to the inboard sec-tion of each wing leading edge, whichgenerates a buffet at high angle of attackto warn of an impending stall and ensuresthat the inboard wing stalls first

• Aileron gap seals

On airplanes equipped with the Dee HowardMKII wing, the vortex generators have beenremoved. Wing-to-tip fuel tanks strakes havebeen enlarged and air gap seals have been in-stalled forward of the ailerons. Stall stripshave been added to the leading edges of thewings. Airplanes equipped with the DeeHoward XR wing have elongated and curvedstrakes, stall strips, boundary layer energiz-ers, two span flow limiters per wing, and athickened wing root area formed by two glovetanks containing 40 gallons of fuel per side.

EMPENNAGEThe T-tail empennage (Figure 1-16) includesa vertical stabilizer with an attached rudder anda variable incidence horizontal stabilizer withattached elevators. An electrically operatedactuator pivots the stabilizer about the rearspar attach point for longitudinal trim.


Figure 1-16. Empennage



The sweptback vertical stabilizer is formed byfive spars, which are securely connected in thetailcone. The vertical stabilizer is the mount-ing point for the rudder and horizontal stabi-lizer. At the lower leading edge of the verticalstabilizer is a dorsal fin which houses a ram-air scoop. Model 25 series airplanes have theoxygen bottle located within the dorsal fin.

The horizontal stabilizer is a sweptback, full-span unit, constructed around five spars. It isattached to the vertical stabilizer at two points:

• The center aft edge attaches to a heavy-duty hinge pin.

• The center leading edge attaches to anelectrically operated jackscrew to pro-vide pitch axis trim.

AIRPLANE SYSTEMSThe following is a brief introduction to themajor airplane systems on the Learjet 20 se-ries airplanes. Detailed descriptions of thesesystems are contained within the individualchapters of this training manual.

ELECTRICAL POWERSYSTEMSDC electrical power is provided by two engine-driven generators. Backup power is suppliedby two 24-volt batteries. The airplane isequipped with one or more emergency batter-ies. The airplane also has the capability of ac-cepting DC power from a ground power unit.

AC power is provided by either two or threesolid-state static inverters.

LIGHTINGInterior lighting is supplied for general cock-pit use and for instrument illumination. Cabinlighting consists of overhead lighting, individ-ual passenger positions, and cabin baggagecompartment lighting.

Exterior lighting includes the combinationlanding-taxi light on each main gear, naviga-tion lights, anticollision lights, strobe lights,and a recognition light. A second recognitionlight and wing ice inspection light are avail-able as options.

The glareshield warning light system consistsof three horizontal rows of red, amber, andgreen lights (see “Annunciator Panel” sec-tion), which alert the pilots to various malfunc-tions or switch positions. They are located onthe center portion of the glareshield just abovethe center instrument panel. Airplanes with-out glareshield warning lights have a readoutpanel on the instrument panel. Unless retro-fitted, these include airplanes SNs 23-003through -099, 24-100 through -155, and 25-003through -009.

FUEL SYSTEMFuel is contained in integral wing tanks, tiptanks, and in a bladder fuselage tank just aftof the rear pressure bulkhead on all models ex-cept the 24E. The 25C model is a long-rangeversion with a larger fuselage tank. Fueling isaccomplished through filler caps in the top ofeach tip tank. To fill the fuselage tank, fuelmust be transferred from the wing tanks.

POWERPLANTThe airplane is powered by two GeneralElectric CJ-610 series, single-rotor, axial-flow turbojet engines. The engines are in-stalled in pylon-mounted nacelles on eachside of the aft fuselage. An engine-drivenstarter-generator is installed in each engine toprovide engine starting and furnish DC powerto the electrical system. Engine bleed air isused for cabin pressurization and heating,windshield defogging, wing anti-ice, and en-gine front frame anti-icing.

All Learjet 20 series airplanes are equippedwith engine fire detection and fire-extinguish-ing systems except model 23 which has detec-tion only. The systems include detection circuitswhich give visual warning in the cockpit and




controls to activate one or both fire-extinguisherbottles. There is a test function for the fire de-tection system. One or two portable fire extin-guishers are stowed inside the airplane.

The airplane pneumatic system uses bleed airextracted from the engine compressor sec-tions. It includes controls for regulation anddistribution of air for heating, cooling, venti-lation, pressurization, and anti-icing.

ICE AND RAIN PROTECTIONThe anti-icing systems use engine bleed air,electric heating, and alcohol.

Bleed air is used to heat the wing leadingedge, the windshields, and engine front frame.Bleed air is also used to remove rain from thewindshield.

Electrically heated systems include the hori-zontal stabilizer leading edge, nacelle lips,pitot tubes, static ports, and the stall warningvanes. Airplanes not certified for flight inknown icing do not have stabilizer heat.

An alcohol system is used for radome anti-icing and to back up the pilot’s windshieldbleed-air anti-icing. Airplanes that are notcertified for flight in known icing supply al-cohol to both the pilot and copilot windshieldsand not to the radome. The Learjet 23 does nothave an alcohol system.

AIR CONDITIONING AND PRESSURIZATIONThe Freon refrigeration system is used forground cooling, in-flight cooling, and cabinair dehumidification. Cabin heating is pro-vided by electric cabin heat (optional) andengine compressor bleed air used for cabinpressurization.

Pressurization is provided by engine bleedair. Cabin altitude is maintained by regulat-ing the outflow of air from the cabin. Thepressurization module provides either auto-matic or manual control of cabin pressuriza-tion while airborne.

HYDRAULIC POWER SYSTEMSThe hydraulic system supplies pressure forthe operation of the landing gear, gear doors,brakes, flaps, spoilers, and thrust reversers, ifinstalled. A single reservoir supplies fluid tothe two engine-driven pumps through the hy-draulic shutoff valves. An electric auxiliarypump in the tailcone is used for hydraulicpressure on the ground when the engines arenot operating and in flight during a hydraulicfailure. It will supply pressure to all hydrauli-cally operated systems on the airplane.

LANDING GEAR AND BRAKESLearjet 20 models have retractable tricyclelanding gear which is electrically controlledand hydraulically operated.

An emergency air bottle, located in the right sideof the nose compartment, can be used to extendthe landing gear or for emergency braking, orboth, in case of hydraulic or electrical failure.

The self-centering nose gear has a single wheeland incorporates an electrical nosewheel steer-ing system which has either variable or non-variable authority, depending upon airplanemodel and serial number.

Each main gear has dual wheels, each equippedwith multiple-disc brakes. Hydraulic brakingis controlled from either the pilot’s or copilot’sstation. An antiskid system provides maxi-mum braking performance while protectingagainst skids.

FLIGHT CONTROLSLearjet 20 models use manually actuated pri-mary flight controls. Pilot inputs are transmit-ted via cables, bellcranks, and pushrods to theailerons, rudder, and elevators. There are nohydraulic or electric power boosts for thesesystems. Primary control trim is electricallycontrolled and operated.

Secondary flight controls (spoiler and flaps) areelectrically controlled and hydraulically operated.




PITOT-STATIC SYSTEMThere are two heated tubes installed—one oneach side of the nose compartment.

There are five static ports—two on the pilot’sside and three on the copilot’s side of the nose.All are heated except for the rearmost port onthe copilot’s side.

The rearmost static port on the copilot’s sideprovides a static source for the pressurizationmodule.

This port has a backup port in the nose com-partment should the external port becomeclogged.

Two shoulder static ports for the air data sen-sor, if installed, or the autopilot altitude con-troller are located on top of the nose, forwardof the windshield.

Two nonheated static ports are located insidethe nose compartment. One is used by the cabinaltitude controller. The other is the alternatestatic source for the pilot’s instruments.

A static port in the tailcone compartment pro-vides ambient air pressure for operation ofthe cabin safety valve.

OXYGEN SYSTEMThe oxygen system consists of the crew andpassenger distribution systems connected toa high-pressure oxygen storage cylinder lo-cated in either the nose compartment or thedorsal fin of the vertical stabilizer. The flightcrew is provided with pressure-demand or di-luter-demand masks. Constant-flow masks areprovided for the passengers. The passengermasks may be deployed either manually atany cabin altitude or automatically if cabin al-titude increases to 14,000 feet.

STATIC OR LIGHTNING STRIKEPROTECTIONThe radome is protected with four aluminumtapes bonded and grounded to the fuselage.Static dischargers are installed on the trail-ing edges to dissipate static electricity witha minimum amount of radio interference(Figure 1-17). The static dischargers are in-stalled on the tip tank fins, tailcone verticalfin, navigation light fairing, and elevatortrailing edge.




Figure 1-17. Static Wicks (Typical)

2-i

CHAPTER 2ELECTRICAL POWER SYSTEMS

CONTENTS

Page

INTRODUCTION ................................................................................................................... 2-1

GENERAL............................................................................................................................... 2-1

DC POWER............................................................................................................................. 2-2

Batteries ........................................................................................................................... 2-2

Optional Emergency Battery System............................................................................... 2-4

Generators and Regulators............................................................................................... 2-6

Ground Power .................................................................................................................. 2-7

Circuit Components ......................................................................................................... 2-8

Distribution .................................................................................................................... 2-11

AC POWER........................................................................................................................... 2-14

Inverters ......................................................................................................................... 2-14

Controls.......................................................................................................................... 2-14

Indicators ....................................................................................................................... 2-14

AC Operation ................................................................................................................. 2-14

Distribution .................................................................................................................... 2-15

QUESTIONS......................................................................................................................... 2-16




2-iiiFOR TRAINING PURPOSES ONLY



ILLUSTRATIONS

Figure Title Page

2-1 Battery Installation ................................................................................................... 2-2

2-2 Battery Switch and DC Voltmeter............................................................................ 2-3

2-3 Battery Switches....................................................................................................... 2-3

2-4 Emergency Battery No. 1 Switch and Schematic .................................................... 2-5

2-5 Generator Cooling Scoop......................................................................................... 2-6

2-6 Generator Switches .................................................................................................. 2-7

2-7 DC Ammeters........................................................................................................... 2-7

2-8 Ground Power Receptacle........................................................................................ 2-7

2-9 Pilot’s Circuit-Breaker Panel (Typical) .................................................................... 2-9

2-10 Copilot’s Circuit-Breaker Panel (Typical) ............................................................. 2-10

2-11 DC Power Distribution—SNs 24-260, -264, and Subsequent,and 25-090, -103, and Subsequent......................................................................... 2-12

2-12 DC Power Distribution—SNs 24-230 through -259, -261 through -263, 25-061 through -089, and -091 through -102 ................................. 2-13

2-13 AC Power Distribution (Typical) ........................................................................... 2-15

2-14 Inverter Controls .................................................................................................... 2-14

The electrical system incorporates a multiplebus system for power distribution. The systemis interconnected by relays, current limiters,overload sensor, and circuit breakers which au-tomatically react to isolate a malfunctioningbus. It is possible to power the entire DC andAC electrical systems from the airplane bat-teries, a single engine-driven generator, or aGPU. In the event of a double generator fail-ure the batteries may be used for a limitedtime. An optional emergency battery systemcan be used to power a standby attitude indi-cator and the gear, flaps, and spoilers. A sec-ond emergency battery may be installed topower a standby directional gyro.

There are certain circuits which are poweredfrom the batteries with the battery and gener-ator switches turned off. The “hot-wired”items that are connected to the batteries are:

• Left stall warning

• Cabin door motor

• Entry lights

• Right stall warning

• Firewall shutoff valve warning lights(pinhead)

CHAPTER 2ELECTRICAL POWER SYSTEMS




INTRODUCTIONThis chapter covers the electrical power systems for SNs 24-230 through 24-357, 25-061,and 25-070 through 25-373.

DC electrical power on the Lear 20 series airplanes is provided by two engine-drivenstarter generators. Backup DC electrical power is supplied by two batteries. A groundpower receptacle allows connection of the ground power unit. Ground power may be usedfor system operation or engine starting. AC electrical power is provided by two or threesolid-state inverters located in the tailcone.

GENERAL

#1 S

ERVO

SYSTEM

BATT HOT

BAT OFF

AC

GEN

#1 D

C

GEN

#1 E

NG

OIL PL

DC POWER

BATTERIESTwo batteries (Figure 2-1), located in the tail-cone, provide the secondary source of DCpower. Each battery has a removable coverand a case which is vented and cooled by over-board connections. Lead-acid batteries areenclosed in a plastic case. Nickel-cadmium(nicad) batteries are enclosed in a stainlesssteel case. The batteries are of sufficient ca-pacity to supply normal electrical require-ments for a short period of time. The batteriesmay be used for engine starting when exter-nal power is not available. Minimum voltagefor engine starting is 23 VDC for nicad bat-teries and 24 VDC for lead-acid batteries whenthe outside temperature is less than 70°F.

Either two 24-VDC nickel-cadmium or two 24-VDC lead-acid batteries may be installed.Nickel-cadmium batteries are standard equip-ment on airplanes through SN 24-355 and

through SN 25-235. Lead-acid batteries maybe installed on these airplanes with ECR 1663.Lead-acid batteries are standard equipmenton SNs 24-356 and subsequent, and 25-236and subsequent. Nickel-cadmium batteries areoptional.

Because of the poor regulation of most groundpower units, it is not recommended that eithernicad or lead-acid batteries be charged in theairplane. Should it become absolutely neces-sary to charge a nicad battery in the airplane,the battery voltage must not be less than 20.5volts, and the temperature must be less than100°F (37.8°C). All airplane switches must beoff and the ground power unit must be limitedto 28 volts and 1,000 amperes maximum. Onairplanes with dual battery switches, bothswitches must be in ON. On airplanes with asingle battery switch, electrical power is ap-plied directly to the batteries when the groundpower unit is connected. Voltage, amperage,and temperature must be constantly monitoredduring charging. Maximum permissible tem-perature rise during charging is 15°F (9.4°C).




VENT HOSE(TO LH VENT)


LH BATTERY

NEOPRENECOVERING

BATTERY TEMPERATUREINDICATOR PLUG (OPTIONAL)

BATTERY OVERHEATWARNING SYSTEMPLUG

QUICKDISCONNECT(TYPICAL)

VENT HOSE(TO RH VENT) RH BATTERY

Figure 2-1. Battery Installation

Battery Overheat Warning—Nicad Batteries OnlyThere are two red battery overheat lights(Appendix B), one labeled “BAT 140” and theother “BAT 160.” The BAT 140 light indicatesthat either or both batteries have reached 140°F.Check the battery temperature indicators, if in-stalled, to determine which battery is over-heating. The BAT 160 light indicates that eitheror both batteries have reached 160°F.

Dual battery temperature indicators, (if in-stalled), one for each battery, are provided togive a readout of battery temperature so thatcorrective action may be taken in the event ofbattery overheat.

Main Battery Switch and RelaysSNs 24-230 through -259, -261, -262,-263, and 25-061, -070 through -089,and -091 through -102One two-position ON–OFF battery switch(Figure 2-2) is provided to interconnect the bat-tery bus to the battery charging bus via the bat-tery relay. This allows charging of the batteryfrom the generators or provides battery andGPU power to the airplane systems. The bat-tery switch, when in ON, provides a ground tothe battery relay.

When the battery switch is in ON, battery volt-age is indicated on the DC voltmeter if noother power source is connected to the batterycharging bus.

NOTEBatteries are connected in parallel atall times, including engine starts.

SNs 24-260, -264, and Subsequent,and 25-090, -103, and SubsequentTwo battery switches (Figure 2-3) are locatedon the center instrument panel between thestarter-generator switches. Each switch hasan ON and OFF position. If a switch is on andthe respective battery has approximately 16volts available, the battery relay closes and the




Figure 2-2. Battery Switch and DCVoltmeter

Figure 2-3. Battery Switches

battery is connected to the battery bus andbattery charging bus. Individual battery volt-age may be read by turning on the respectivebattery switch if no other power source is con-nected to the battery charging bus.

NOTEThe batteries are connected in par-allel, including engine starts, onlywhen both battery switches are on.

Emergency Operation—Airplane Batteries OnlySee Section III of the appropriate AFM.

OPTIONAL EMERGENCYBATTERY SYSTEMOne or two emergency batteries may be in-stalled in the airplane in either the tailcone ornose section. The batteries are dry cell nicadand are rated at 25 VDC. A switch and anamber light are provided with each battery. A100-VA inverter powered by the battery sup-plies 115-VAC, 400-Hz, single-phase powerfor the standby attitude indicator and emer-gency directional gyro, if installed. The bat-tery receives a trickle-charge from the mainDC bus. The emergency battery or batteriesmay operate the following equipment:

• Emergency attitude gyro

• Gear

• Flaps

• Spoilers

• Emergency directional gyro

• Emergency communications radio

Operation—Airplanes Equippedwith a Single EmergencyBatteryThe system is controlled by a single ON–OFFswitch or, on 25-209 and subsequent, by anON–OFF–STBY emergency power switch.Both switch installations (Figure 2-4) are lo-cated on the pilot’s instrument panel. If air-plane power is lost, the emergency batteryprovides DC, 115-VAC, and 4.6-VAC powerthrough a self-contained inverter. The emer-gency attitude indicator is powered and illu-minated with AC power and the emergencybattery indicator light, and the gear, flaps andspoilers are powered with DC power. Some air-planes have an emergency communicationsradio wired to the emergency battery. Airplaneswith a two-position switch may or may notpower the gear, flaps, and spoilers.

Airplanes Equipped with DualEmergency BatteriesThe dual battery system is controlled with anON–OFF and an ON–OFF–STBY emergencypower switch. When airplane electrical poweris lost, the two-position switch provides powerto its associated avionics and emergency bat-tery indicator light. The three-position switchprovides power to its associated circuits, nor-mally the gear, flaps and spoiler systems, theemergency attitude indicator and light, andthe emergency battery indicator light. TheSTBY position is used to conserve batterypower when the expected flight time exceeds30 minutes by removing power to the gear,flaps, and spoilers.

With either a dual or single emergency batterysystem, the battery indicator lights illuminateonly when the emergency battery is providingpower. In this case, the spoiler light and flapindicator are inoperative but the gear indica-tor lights still function.







EM

ER

GE

NC

YB

ATT

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Y C

B

28 V

DC

INP

UT

BAT

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VD

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R

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ER

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YB

ATT

ER

Y

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PU

T

ATT

ITU

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GY

RO

115

VA

C O

UT

PU

T

4.6

VA

C O

UT

PU

TRH

ES

S D

C B

US

INP

UT

GE

AR

S

FLA

PS

SP

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ER

S

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Fig

ure

2-4

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mer

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Sw

itch

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chem

atic

GENERATORS ANDREGULATORS

GeneratorsTwo independent, engine-driven DC genera-tors, rated at 30 volts, 400 amperes, are pro-vided to supply the airplane with DC power.The generators are air cooled by ram air froma scoop on the engine nacelle (Figure 2-5).

Generator Control PanelThe generator control panel, installed in thetailcone, serves as the generator control unitfor the DC electrical system. The generatorcontrol panel incorporates the battery, exter-nal power, right and left generator control,overvoltage control, and equalizer circuit re-lays. Battery, external power, and generatoroutputs are connected to the generator controlbox which distributes DC power through cur-rent limiters and circuit breakers to the vari-ous electrical systems.

Voltage RegulatorsTwo solid-state voltage regulators installed inthe tailcone maintain a constant output volt-age under varying engine speeds and loadconditions. The voltage regulator is factory-adjusted, and no adjustment is allowed. Thedesign of the regulators precludes voltage drift

due to temperature. Each regulator is set to 28.5±0.3 VDC.

Should generator output exceed 32 VDC, acircuit within the voltage regulator completesa ground circuit and reduces generator outputto zero.

ControlsThe generator switches (Figure 2-6) locatedon the lower portion of the instrument panelhave three positions labeled “OFF,” “START,”and “GEN.” OFF position deenergizes thestarter-generator. START position energizesthe engine start cycle. GEN position suppliesvoltage to the generator relay, and throughthe voltage regulator, to the field windings, en-ergizing the generator to supply power to theairplane electrical system.

Generator Reset SwitchesTwo generator reset switches, one left and oneright, are installed on the switch panel adja-cent to its respective generator switch. Whenin the RESET position, contacts complete aground circuit to reset the respective genera-tor overvoltage relay.

Warning SystemL GEN and R GEN warning lights (AppendixB) are installed on the glareshield. The warn-ing lights illuminate if a generator has failedor the switch is in OFF.

IndicatorsA DC ammeter is installed on the center instru-ment panel for each generator (Figure 2-7). Theammeter indicates the load being carried bythe respective generator. The ammeter red lineis 400 amperes and has a yellow arc from 300to 400 amperes. The limit when OAT is 60°For above is 300 amperes.

A single DC voltmeter indicates DC voltagepresent on the battery-charging bus. The DCvoltmeter red line is 35 VDC.




Figure 2-5. Generator Cooling Scoop

GROUND POWERA ground power unit (GPU) can be connectedto the airplane through a power receptacle(Figure 2-8) located on the left side of the air-plane adjacent to the tailcone door. GPU capa-bility should be limited to 28 VDC at a maximumof 1,000 amperes. The external power overvolt-age cutout circuit limits the maximum voltageto 33 ±2 VDC. The 1,000-amp limit is to pre-vent exceeding starter shaft torque limits.

Airplanes with a Single BatterySwitch(SNs 24-230 through -263, excluding-260, and 25-070 through -102,excluding -090)To charge both batteries using external power,plug in the GPU and regulate the voltage out-put to 28 VDC. Ground power unit output shouldbe limited to a maximum of 1,000 amperes.

Airplanes with Dual BatterySwitches(SNs 24-260, -264, and Subsequent,and 25-090, -103, and Subsequent)To charge both batteries using external power,turn on both battery switches. Plug in the GPUand regulate the voltage to 28 VDC. Ground

power unit output should be limited to a max-imum of 1,000 amperes.




Figure 2-6. Generator Switches

Figure 2-7. DC Ammeters

Figure 2-8. Ground Power Receptacle

CIRCUIT COMPONENTS

Current LimitersA current limiter panel is installed on the gen-erator control panel in the tailcone. A currentlimiter is comparable to a slow blow fuse and,once blown, must be replaced. There is a sparecurrent limiter box in the tailcone with onespare current limiter for each size used in theairplane. This spare current limiter box has theinventory printed on the top cover. Currentlimiters are used to provide power to somesystems, to back up circuit breakers, and to in-terconnect buses.

Circuit BreakersAll circuit breakers are located in the cockpiton two circuit-breaker panels (Figures 2-9and 2-10), one on each side of the cockpit. Allcircuit breakers are the push-pull type. The DCcircuit breakers operate thermally, and the ACcircuit breakers operate magnetically. Theamperage ratings are stamped on the top ofeach circuit breaker.

RelaysRelays are used at numerous places through-out the electrical distribution system, partic-ularly in circuits with heavy electrical loads.The relays function as remote switches to“make” or “break” power circuits. This arrange-ment allows the control circuit and wiring tothe control switches to be a much smaller gagewire. Battery relays, external power relays,starter relays, generator relays, and inverter re-lays are used to connect/control the power cir-cuits. Instrument panel switches complete thecontrol circuits and operate the relays. Relaysare also used to connect the power circuits tothe left and right main DC buses.

Overload SensorsThe DC overload sensors are located in the tail-cone. An overload control sensor consists ofa 70-ampere thermal circuit breaker mechan-ically connected to a set of switch contacts.

During normal operation, the main bus powerrelay is energized through the closed contactsof the overload control sensor and the main buscircuit breaker. If an overload condition exists,the thermal breaker contained within the sen-sor positions the contacts to ground the circuitfrom the applicable bus circuit breaker. Theopened circuit breaker removes power from themain bus relay. The overload sensor automat-ically resets when the thermal breaker hascooled. However, power is not restored to thepower relay and bus until the DC bus circuitbreaker is reset.

Reverse Current DiodesTwo reverse current diodes are installed on theaft side of the engine beam in the tailcone(one for each generator system). The reversecurrent diodes isolate the generator systemsfrom each other and prevent the starter frommotoring when not selected.







L IGNST-GEN

L STALLWRN

L ESSBUS

L AIRIGN

L FIREDET

L FIREEXT

L FWSOV

L JETPMP VAL

L STBYPMP

L ICEDET

RAMAIRTEMP

PITCH

ROLL

YAW

AIRBL

OXYVAL

FUSLGVAL

FUELQTY

DCVM

FUELJTSN

AUD 1

COMM 1

NAV 1

ATC 1

TURNCOORD

WRNLTS

INSTRLTS

L PITHT

PRI FLTDIR

WHL MSTRMNVR

PRIAFCS

AFCSPITCH

AFCSROLL

AFCSYAW

S WRNHT

ARMCONT

PRILCH

REMERSTOW

INDLTS

AUXLCH

DOORACTR

ENTLTS

ADF 1

DME 1

L LDG &TAXI LT

NAVLTS

STROBELTS

CABLTS

FREONCONT

CABBLO

L ECSVAL

NOSESTEER

ANTISKID

STEREO

SQUATSW

HFCOMM

EMERLTS

HT VALIND

HFCOMM

ALCPMP

AUXCAB HT

WINGINSP

LT

L MAINBUS

PRIINV

L ACBUS

PRIVM

PRI DIRGY

PRIVERT GY

LEPR

PRIAFCS

RADARGY

FLT DRATTD

FLT DRCMD

FLT DRHD

NOSESTEER

AIRDATA

26 VACBUS

DMEREAD

L OILPRESS

PRIRMI

NAV 1

ADF 1

ALTM& ROC

ANTISKID

TONEGEN

L AUXAC BUS

THRUST

REVERSER

MAIN BUS

ESSENTIAL BUS

HOT-WIRED ITEMS

LEGEND

AC BUS

Figure 2-9. Pilot’s Circuit-Breaker Panel (Typical)




R AIRIGN

R ESSBUS

R STALLWARN

EMERBAT 2

AC BUSTIE

ESSBUS TIE

MAINBUS TIE

EMERBAT 1

R IGNST-GEN

R MAINBUS

SECINV

R ACBUS

SECVM

SECDIR GY

SECVERT GY

REPR

SECAFCS

26 VACBUS

FLT DIRATTD

FLT DIRCMD

FLT DIRHD

DMEREAD

FLTRCDR

R OILPRESS

SECRMI

FLTRCDR

R AUXAC BUS

NAV 2

ADF 2

ADF 2

DME 2

R LDG &TAXI LT

BCNLTS

RADALTM

ALTM

WSHLDHT

WINGHT

L NACHT

R NACHT

STABHT

FUSLGPMP

FILL &XFER

HRMTR

ENGSYN

TOILET

BATTTEMP

CLOCK

R ECSVAL

OVEN

INTPH

RADAR

SOVLTS

AUD 2

COMM 2

NAV 2

EMER PTRIM

ATC 2

TURN& BANK

WRNLTS

INSTRLTS

R PITHT

CABPRESS

MKRBCN

SECAFCS

SEC FLTDIR

S WRNHT

AUXINV

ARMCONT

PRILCH

LEMERSTOW

INDLTS

AUXLCH

R FIREDET

R FIREEXT

R FWSOV

R JETPMP VAL

R STBYFUS PMP

R ICEDET

OILTEMP

FUELQTY

TAB &FLAP PN

FLAPS

GEAR

SPOILER

AUXCOMM

RECOGLT

THRUST

REVERSER

MAIN BUS

ESSENTIAL BUS

HOT-WIRED ITEMS

LEGEND

AC BUS

Figure 2-10. Copilot’s Circuit-Breaker Panel (Typical)

DISTRIBUTIONA multiple bus, multiple conductor system isused (Figures 2-11 and 2-12). Buses and majorcircuits are protected by strategically placedrelays, current limiters, overload sensors, andcircuit breakers to preclude total failure andto simplify the isolation of any partial systemmalfunction. All circuit breakers are accessi-ble to the crew during flight.

Battery BusAirplanes with a Single BatterySwitchThe battery bus is connected directly to bothbatteries at all times. The battery switch ener-gizes a relay that connects the battery bus tothe battery charging bus when turned on. Ifvoltage on the small pin of the GPU recepta-cle does not exceed 33 ±2 VDC, external poweris applied to both batteries and the battery bus.This permits power from a GPU or the batter-ies to supply the airplane electrical system. Italso permits the generators to charge the bat-teries. During engine start, a relay connects thebattery bus to the starter. Hot-wired items areconnected directly to the battery bus.

Airplanes with Dual BatterySwitchesThe battery bus is connected to the two bat-teries via the battery relays when the corre-sponding battery switch(es) is turned on. Thebattery bus may be powered by external powerif the voltage from the ground power unit isless than 33 ±2 VDC on the small pin of theGPU receptacle, and at least one battery switchis on. As long as both battery switches are on,either engine DC generator charges both bat-teries. The engine generators are able to sup-ply DC power to all DC-powered equipmenton the airplane with both battery switches inOFF except for the hot-wired items. All hot-wired items are connected directly to the bat-teries. The battery bus is connected to the leftor right starter through the associated startrelay when the respective generator switch isplaced in the START position.

Battery Charging BusThe battery charging bus is located in the cur-rent-limiter panel in the tailcone. It distributespower from the batteries or GPU to the airplaneelectrical system and generator power back tothe batteries for charging. The battery charg-ing bus distributes power to the generatorbuses, essential DC buses, and the battery bus.

Generator BusesThe left and right generator buses are a part ofthe current-limiter panel in the tailcone. Thegenerator bus may be powered by the batterycharging bus or by the output of the generator.

Main DC Power BusesThe two main DC power buses receive powerfrom the respective generator bus through a 10-ampere current limiter. The power buses sup-ply their respective starter switches, inverterswitches, and main bus relays.

Main DC BusesThe two main DC buses receive power fromtheir respective generator buses. Circuit pro-tection for each bus is provided by a 70-am-pere overload sensor. The left and right mainbuses are interconnected by a 50-ampere DCbus-tie circuit breaker.

Essential DC BusesThe two essential DC buses receive powerfrom the battery charging bus. Each essentialDC bus is protected by a 50-ampere currentlimiter and a 40-ampere left or right bus cir-cuit breaker. The buses are interconnected bya 20-ampere DC bus-tie circuit breaker. OnSNs 25-368 through -373 and airplanes mod-ified by AMK 85-1, the essential buses arepowered by their respective generator buses.







PRIMARYINVERTER

START-GEN START-GEN

AUXINVERTER

SECONDARYINVERTER

LMAIN BUS

LGEN BUS

RGEN BUS

BATCHG BUS

BATTERY BUS

RMAIN BUS

L ESS

L PWR BUS R PWR BUS

R ESS

RBATTERY

LBATTERY EXTERNAL

POWER

EXT POWEROVERVOLTAGE

LSTARTER

GEN

RSTARTER

GEN

DCVOLTS

AMPSAMPS

70 A60 A

60 A

50 A

40 A

275A

40 A

70 A 60 A

50 A

275A

275A

STARTRELAY

STARTRELAY

LBATTERY

RELAY

RBATTERY

RELAY

R BATTERY BUSL BATTERY BUS

10 A10 A

20 A 20 A

100A 100A

50 A

20 A

FREON SYSSTAB HEAT

**

LNACELLEHEAT

RNACELLE

HEAT

OFF OFF

GENFIELD COIL

GENFIELD COIL

ON

ON

33 ± 2 VDC

* SNs 25-368 AND SUBSEQUENT AND AIRPLANES WITH AMK 85-1

Figure 2-11. DC Power Distribution—SNs 24-260, -264, and Subsequent, and 25-090,-103, and Subsequent


PRIMARYINVERTER

START-GEN START-GEN

AUXINVERTER

SECONDARYINVERTER

LMAIN BUS

LGEN BUS

RGEN BUS

BATCHG BUS

BATTERY BUS

RMAIN BUS

L ESS

L PWR BUS R PWR BUS

R ESS

RBATTERY

LBATTERY EXTERNAL

POWER

EXT POWEROVERVOLTAGE

LSTARTER

GEN

RSTARTER

GEN

DCVOLTS

AMPSAMPS

70 A60 A

60 A

50 A

40 A

275A

40 A

70 A 60 A

50 A

275A

275A

STARTRELAY

STARTRELAY

R BATTERY BUSL BATTERY BUS

10 A10 A

20 A 20 A

OFF

ON100A 100A

50 A

20 A

FREON SYS

STAB HEAT

**

LNACELLEHEAT

RNACELLE

HEAT

GENFIELD COIL

GENFIELD COIL

33 ± 2 VDC

*AIRPLANES WITH AMK 85-1

BATTERYRELAY

Figure 2-12. DC Power Distribution—SNs 24-230 through -259, -261 through -263,25-061 through -089, and -091 through -102



AC POWER

INVERTERSAlternating current for the AC electrical instru-ments and electronic equipment is provided bytwo or three 1,000-VA solid-state static invert-ers located in the tailcone. Figure 2-13 illustratesthe typical inverter system. The auxiliary inverteris an optional item. The primary and secondaryinverters power their respective buses througha paralleling control box. Inverter output is 115VAC, 400-Hz, single-phase. The 26-VAC poweris provided by two stepdown transformers lo-cated, one on each side, in the cockpit aft of thecircuit-breaker panels. These two transformersstep down 115 VAC to 26 VAC. Other compo-nents in the system include power relays, par-alleling box, overload sensors, circuit breakers,and inverter failure lights. The primary and sec-ondary fail lights are red, and the auxiliary faillight is amber. All fail lights are located on theglareshield.

CONTROLSTwo (or three) inverter on-off switches (Figure2-14), one for each inverter, are installed onthe lower center instrument panel. If the op-tional auxiliary inverter is installed, anotherswitch labeled “LH” or “RH AC BUS” is usedto direct the auxiliary inverter output to the se-lected bus.

INDICATORSTwo or three inverter warning lights (Appen-dix B) labeled “PRI INV,” “SEC INV,” and“AUX INV” are installed on the glareshield.If an inverter fails, its respective warning lightilluminates.

On SNs 24-230 through -277, 25-070 through25-134, and 25-136 through -140, turning offthe inverter switch extinguishes the inverterfail light.

On SNs 24-278 and subsequent, and 25-135 and-141 and subsequent, if the inverter fails andthe switch is turned to OFF, the PRI INV andSEC INV fail lights remain illuminated. If the

auxiliary inverter fails, turning off the inverterswitch extinguishes the fail light.

A single AC voltmeter with a two-position se-lector switch is provided to monitor the volt-age on the selected bus.

AC OPERATIONWhen the primary inverter switch is set toPRI, the control circuit is completed to the in-verter power relay which closes and connectsthe inverter to the generator bus through theoverload sensor. The inverter AC output is fedto the paralleling control box and on to the leftAC bus through the 10-ampere L AC BUSfeeder circuit breaker. Operation of the sec-ondary inverter is identical.

An overload on either 115-VAC bus causesthe affected AC BUS feeder circuit breaker toopen, activating the BUS TIE circuit breaker,and thus isolating the bus.

If a red INV light illuminates with the inverterturned on, refer to the “AC Power Loss” check-list in the “Emergency Procedures” section ofthe approved AFM. An illuminated INV lightdoes not necessarily indicate inverter failure.




Figure 2-14. Inverter Controls

DISTRIBUTIONAlternating current from the inverters is dis-tributed through the left and right AC buses.Primary inverter output goes to the left bus,secondary inverter output to the right bus.These two buses are tied together by the ACbus-tie circuit breaker, enabling either inverterto power all AC buses. Auxiliary inverter out-put, if installed, may be directed to either theleft or right bus.

Two step-down transformers draw 115-VACpower from their respective 115-VAC buses,




50 A 10 A10 A

26 VAC

BUS 2 A

10 A10A

7.5 A 2 A

10 A10A

26 VAC

BUS

L ACBUS

R ACBUSTRANS

PARALLELING BOX

TRANS

SECONDARYINVERTER

AUXINVERTER

PRIMARYINVERTER

L PWRBUS

R PWRBUS

R ESS

L GENBUS

R GENBUS

BAT CHG BUSL

GENR

GEN

AC VOLTS

60 A

L AC BUSCB

R AC BUSCB

L AUXAC BUS CB

R AUXAC BUS CB

AUX INVPRI-SEC SW

L BUS

R BUS

POWERRELAY

POWERRELAY

POWERRELAY

AUX INVCB

SEC INVSW

SEC INVCB

PRI INVSW

PRI INVCB

AUX INVON-OFF SW

60 A

60 A

275A

275A

BATTERY POWER

GENERATOR POWER

LEGENDGROUNDPOWER

PRIMARY INVERTER POWER

AUXILIARY INVERTER POWER

SECONDARY INVERTER POWER

P S

reducing the voltage to 26 VAC. The 26 VACis then distributed to the installed equipmentrequiring 26 VAC.

Each 115-VAC and 26-VAC bus is protectedby its respective bus circuit breaker.

The 26-VAC BUS circuit breaker applies to SNs24-315 and subsequent, and 25-189 and subse-quent, except 25-191. On prior SNs the trans-formers were wired directly to the bus.

Figure 2-13. AC Power Distribution (Typical)

1. The systems powered directly from thebattery bus with the airplane batteryswitches in OFF are:A. Fuel shutoff valves, door motor, and

entry lightsB. Shutoff valve (pinhead) lights, door

motor, left and right stall warning sys-tems, and entry lights

C. Fuel and hydraulic shutoff valves,shutoff valve (pinhead), door motor,and entry lights

D. Fi re-ext inguish ing sys tem, doormotor, entry lights, and shutoff valve(pinhead) lights

2. It is possible to charge the batteries in theairplane if:A. A GPU is plugged in.B. A GPU is plugged in with less than 33

±2 VDC on the small pin.C. Both battery switches are on and a

GPU is plugged in with less than 33±2 VDC on the small pin.

D. The battery switches are off and thestarter-generator switches are in GEN.

3. GPU output should be limited to a max-imum of:A. 600 amperesB. 800 amperesC. 1,000 amperesD. 1,100 amperes

4. Each of the ammeters indicates the cur-rent produced by its respective generator.The DC voltmeter indicates:A. Voltage on the battery busB. Voltage on the essential busC. Voltage on the main busD. Voltage on the battery charging bus

5. The maximum voltage (red line) valueon the DC voltmeter is:A. 24 VDCB. 28 VDCC. 35 VDCD. 33 ±2 VDC

6. The maximum (red line) amperage on theammeter is:A. 400 amperesB. 600 amperesC. 800 amperesD. 1,000 amperes

7. If a generator trips off the line, the cock-pit indications are:A. Battery voltage on the DC voltmeterB. Zero amperes and an illuminated gen-

erator fail lightC. Both an illuminated generator and in-

verter fail lightD. An illuminated generator fail light

and indicated battery voltage

8. Minimum battery voltage for engine startis:NicadA. 16 C. 23B. 24 D. 28Lead Acid (up to 70°F)A. 16 C. 24B. 23 D. 28

9. If both generators fail in flight, two fullycharged batteries supply power for:A. Only DC equipmentB. Approximately 1 hourC. Approximately 30 minutesD. A short period of time




QUESTIONS

10. The number, rating, and type invertersfor your airplane are:A. Two, 750 VA, rotaryB. Three, 400 VA, solid-stateC. Three, two 400 VA and one 750 VA,

solid-stateD. Two or three, 1,000 VA, solid-state

11. The output of the inverters is:A. 115 VAC, 400 Hz, single-phaseB. 115 VAC, 60 Hz, three-phaseC. 26 VAC, 400 Hz, single-phaseD. 26 VAC, 60 Hz, single-phase

12. The 26-VAC power is obtained from:A. Separate inverters in the tailconeB. A s t ep -down t r ans fo rmer in the

tailconeC. A step-down transformer in each

inverterD. Two step-down transformers aft of

the pilot and copilot circuit-breakerpanels

13. The inverters for your particular airplaneare located:A. In the nose compartmentB. Beneath the divan seatC. In the tailconeD. Behind the copilot

14. An inverter failure is indicated by:A. Lack of AC voltageB. Lack of AC cyclesC. An illuminated red fail light (with the

inverter switch in ON)D. Red flags on the instruments supplied

by the respective inverter

15. The first action to take after an AC powerloss is:A. Turn the failed inverter switch to OFF.B. Check for open circuit breakers.C. Pull the AC bus-tie circuit breaker.D. Re fe r t o t he “AC Power Los s”

checklist.




2A-i

CHAPTER 2AELECTRICAL POWER SYSTEMS

CONTENTS

Page

INTRODUCTION ................................................................................................................ 2A-1

GENERAL............................................................................................................................ 2A-1

DC POWER.......................................................................................................................... 2A-2

Batteries......................................................................................................................... 2A-2

Generators ..................................................................................................................... 2A-4

Circuit Components ...................................................................................................... 2A-5

Distribution ................................................................................................................. 2A-14

Ground Power ............................................................................................................. 2A-17

AC POWER........................................................................................................................ 2A-17

Inverters....................................................................................................................... 2A-17

Controls....................................................................................................................... 2A-18

Indicators..................................................................................................................... 2A-20

Operation..................................................................................................................... 2A-20

Emergency Modes ...................................................................................................... 2A-21

Emergency Battery System......................................................................................... 2A-21

QUESTIONS ...................................................................................................................... 2A-24




2A-iii

ILLUSTRATIONS

Figure Title Page

2A-1 Battery Installation ................................................................................................ 2A-2

2A-2 Battery Switch ....................................................................................................... 2A-3

2A-3 Electrical Indicators............................................................................................... 2A-3

2A-4 Generator Switches................................................................................................ 2A-4

2A-5 Circuit-Breaker Panels—Model 23 ....................................................................... 2A-6

2A-6 Tail Compartment Circuit-Breaker Panel—Model 23 .......................................... 2A-6

2A-7 Circuit-Breaker Panels—SNs 24-100 through 24-129 ......................................... 2A-7



2A-10 Circuit-Breaker Panels—SNs 25-003 through 25-010 ....................................... 2A-10

2A-11 Circuit-Breaker Panels—SNs 25-011 through 25-024 ....................................... 2A-11

2A-12 Circuit-Breaker Panels—SNs 24-181 through 24-199,and 25-025 through 25-034................................................................................. 2A-12

2A-13 Circuit-Breaker Panels—SNs 24-200 and Subsequent,and 25-035 and Subsequent ................................................................................ 2A-13

2A-14 DC Power Distribution—SNs 23-003 through 24-189,and 25-003 through 25-029................................................................................. 2A-14

2A-15 DC Power Distribution—SNs 24-190 through 24-229,...................................... 2A-15and 25-030 through 25-064 except 25-061......................................................... 2A-15

2A-16 External Power Receptacle ................................................................................. 2A-17

2A-17 AC Power Distribution—SNs 23-003 through 24-129....................................... 2A-17

2A-18 AC Power Distribution—SNs 24-130 through 24-180,and 25-003 through 25-024................................................................................. 2A-18

2A-19 AC Power Distribution—SNs 24-181 through 24-229, and 25-025 through 25-063, except 25-061 ........................................................................... 2A-19

2A-20 Auxiliary Battery System—SNs 24-100 through 24-129................................... 2A-21

2A-21 Emergency Battery System ................................................................................. 2A-23




INTRODUCTIONThis chapter covers the electrical power systems for SNs 23-003 through 24-229 and 25-003 through 25-064, excluding 25-061.

DC electrical power on the Learjet 20 series airplanes is provided by two engine-drivenstarter-generators. Backup DC electrical power is supplied by two batteries. A groundreceptacle allows connection of the ground power unit. Ground power may be used forsystem operation or engine starting. AC electrical power is provided by two or three solid-state inverters located in the tailcone.

GENERALThe electrical system incorporates a multiplebus system for power distribution. The systemis interconnected by relays, current limiters,overload sensors, and circuit breakers which au-tomatically react to isolate a malfunctioning

bus. Manual isolation is also possible by open-ing the appropriate circuit beakers. It is possi-ble to power the entire DC and AC electricalsystems from the airplane batteries, a single en-gine-driven generator, or a GPU.

#1 S

ERVO

SYSTEM

BATT HOT

BAT OFF

AC

GEN

#1 D

C

GEN

#1 E

NG

OIL PL

CHAPTER 2AELECTRICAL POWER SYSTEMS

2A-1FOR TRAINING PURPOSES ONLY



DC POWER

BATTERIESTwo batteries located in the tailcone (Figure2A-1) provide the backup source of DC power.Each battery has a removable cover and anairtight case which is vented overboard. Thebatteries are of sufficient capacity to supplynormal ground electrical requirements andmay be used for engine starting when exter-nal power is not available. Either lead-acid ornickel-cadmium (nicad) batteries are certi-fied for use in 20 series Learjets. They arerated at 24 VDC and vary in amperage de-pending on size.

Lead-acid batteries are enclosed in a plasticcase. Nickel-cadmium (nicad) batteries are

enclosed in a stainless steel case. Chargingnicad batteries with a GPU is not recom-mended, but if absolutely necessary, the fol-lowing Maintenance Manual restrictionsshould be observed:

• Battery voltage must not be less than20.5 volts (no load).

• Initial battery temperature must not ex-ceed 100ºF (37.8ºC).

• The GPU must be precisely regulated to28 ±0.5 volts/1,000 amperes maximum.

• Voltage and battery temperature mustbe continuously monitored.

• Maximum temperature r ise is 15ºF(9.4ºC) from the starting temperature.

2A-2 FOR TRAINING PURPOSES ONLY



Figure 2A-1. Battery Installation



LH BATTERY

BATTERY HEATINDICATOR PLUG (OPTIONAL)

BATTERY OVERHEATWARNING SYSTEMPLUG

QUICKDISCONNECT(TYPICAL)

VENT HOSE(TO RH VENT) RH BATTERY

FWD

Charging lead-acid batteries in the airplane isnot recommended because of poor GPU out-put regulation, but if absolutely necessary,the GPU should be regulated to 28 ±0.5 volts,all airplane electrical equipment should beturned off, and both batteries and the GPUshould be monitored continuously.

ControlsOne two-position ON–OFF battery switch(Figure 2A-2) is provided to interconnect thebattery bus to the battery charging bus. Poweris routed through the battery relay for batterycharging from the generators or to supply thebattery or GPU power to the airplane systems.The battery switch, when in ON, provides aground to the battery relay. Unless the batteryrelay closes, the batteries cannot be connectedto the airplane electrical system and operateequipment except those hot-wired to the bat-tery buses. The batteries are always connectedin parallel.

IndicatorsElectrical system gages are grouped in a clus-ter on the instrument panel (Figure 2A-3). A

single DC voltmeter indicates the voltage onthe battery charging bus from the highest volt-age input (batteries, generator, or GPU).Airplane generators and GPUs normally pro-duce a higher voltage than the batteries; there-fore, when either of these is powering thebattery-charging bus, battery voltage can nolonger be read. Nickel-cadmium batteries areequipped with an overheat warning light sys-tem, and some airplanes have battery temper-ature indicators. Two red warning l ightslabeled “BAT 140” and “BAT 160” are in-stalled on the instrument panel or glareshield.The warning lights are controlled by thermalswitches within each battery.

The BAT 140 light indicates that one or bothbatteries have reached 140ºF. To determinewhich battery is hot, check the battery temper-ature indicators, then turn off the battery switchand land as soon as practical. The BAT 160light indicates that one or both batteries havereached 160ºF. To determine which battery ishot, check the temperature indicators. Thenturn off the battery switch and land as soon aspossible.




Figure 2A-2. Battery Switch

Figure 2A-3. Electrical Indicators

GENERATORSTwo engine-driven starter-generators, one oneach engine, provide the primary source ofairplane DC power. Each generator is rated at30 volts DC, 400 amperes. Cooling air is routedfrom a scoop on the engine nacelle to the as-sociated generator. During normal operation,both generators operate in parallel throughthe voltage regulators located in the tailcone.As long as the battery switch is in ON, eithergenerator charges both batteries through theassociated 275-ampere current limiter. Thegenerators supply DC power to all DC equip-ment installed on the airplane.

The generator control panel contains relays forthe batteries, starters, GPU overvoltage con-trol, generator overvoltage control, and anequalizer circuit for load sharing.

Voltage RegulatorsSNs 23-003 through 24-189, and SNs 25-003 through 25-029Two Bendix carbon pile voltage regulatorsare installed in the tail compartment on top ofthe engine beam. Each unit consists of a car-bon pile regulator, a resistor, a rheostat assem-bly, and a voltage regulator base. The regulatoris set to maintain voltage at 28.5 VDC.

SNs 24-190 through 24-229, andSNs 25-003 through 25-064Two Phoenix Aerospace solid-state voltageregulators are installed in the tail compart-ment on the left side of the engine beam or elec-trical equipment tray. The regulator maintainsa constant output voltage of 29.0 VDC undervarying engine speeds and load conditions.The voltage regulator is factory adjusted, andno regulator adjustment is allowed.

ControlsTwo starter-generator switches (Figure 2A-4)are installed on the switch panel. Each three-position switch is marked “GEN–OFF–START.”Moving the starter-generator switch to GENallows the generator relays to energize andcomplete three circuits: a circuit for 28 VDC

to the voltage regulator, a circuit for power tothe Freon system compressor motor, and a cir-cuit for the voltage regulator equalizer bus.

When placed to START position, the respec-tive fuel motive flow control valve closes, thestandby fuel pump turns on, and the engine ig-nition system arms. After the motive flow con-trol valve closes, a circuit is completed toenergize the start relays. With the start relaysenergized, battery power is applied to thestarter winding of the starter-generator.

On SNs 23-003 through 23-009 (without jetpumps), there are no motive flow valves andthe pilot must turn on the main fuel pumps. Onairplanes through SN 23-009, the pilot mustalso turn on the individual ignition switches.

Reverse Current DevicesReverse Current Cutout Relays—SNs 23-003 through 24B-189, andSNs 25-003 through 25-029Undervoltage, reverse current, reverse polar-ity, and differential protection in each gener-ator are provided by the reverse current cutoutrelay. The relays are located on the generatorcontrol panel in the tail compartment.




Figure 2A-4. Generator Switches

Reverse Current Diodes—SNs 24-190 through 24-229, andSNs 25-030 through 25-064Two reverse current diodes are installed on theaft side of the engine beam (one for each gen-erator system). The reverse current diodes iso-late the generator systems from each other.

Generator Reset SwitchOperation—SNs 24-190 through 24-229, and SNs 25-030 through 25-064Two generator reset switches (Figure 2A-4),one for each generator, are installed on theswi tch pane l ad jacent to the respec t iveGEN–OFF–START switch. Each switch is atwo-position momentary-type switch. When inthe RESET position, the switch completes aground circuit that resets the generator over-voltage relay. On prior airplanes, the genera-tor reset switch is not installed. Resetting isaccomplished by placing the GEN switch toOFF, then back to ON.

IndicatorsA red generator off warning light above thegenerator switch on SNs 23-003 through 24-142, 24-145, 24-147, 24B-188, and 25-003indicates the generator is off the line. An amberGEN OFF warning light on the glareshieldwarning light panel on all other airplanes in-dicates the generator is off line. These lightsalso illuminate if the generator switch is inOFF.

Two ammeters, one for each generator, indi-cate the load, in amperes, being carried by thegenerators. The load indication is taken fromthe generator field coil. Generator voltage isdisplayed on the DC VOLTS meter and meas-ures the voltage on the battery charging bus.

CIRCUIT COMPONENTS

Current LimitersA current limiter panel is installed on the gen-erator control box in the tail compartment. A

current limiter is like a slow-blow fuse; oncea current limiter has blown, it must be re-placed. Current limiters are used to providepower to certain systems, to back up circuitbreakers, and to interconnect buses.

RelaysRelays are used at numerous places through-out the electrical distribution system, partic-ularly in circuits with heavy electrical loads.The relays function as remote switches tomake or break power circuits. This arrange-ment allows the control circuit wiring to be amuch smaller gage and requires much lesscurrent to operate the relay.

Overload Sensors—SNs 24-190and Subsequent, and SNs 25-030 and SubsequentOverload sensors are used in the power circuitsto the left and right feeder buses. These over-load sensors react thermally to electrical loadsin excess of 70 amperes. They electricallyground the relay control circuit and cause theassociated control circuit breaker to trip, caus-ing the relay to open and break the power cir-cuit. Once the overload condition has beenremoved, the overload sensor cools and re-sets automatically; however, the failed controlcircuit breaker must be reset manually.

Circuit BreakersPush-pull circuit breakers protect all electri-cal systems in the airplane. On SNs 23-003through 24-129, the circuit breakers are in thetail compartment on the generator controlpanel and on the copilot’s armrest in the cock-pit. On SNs 24-130 through 24-229, and 25-003 through 25-064, circuit breakers arelocated on the left side wall of the baggagecompartment and at the copilot’s armrest areaof the cockpit. The DC circuit breakers arethermal, and the AC circuit breakers are mag-netic. The amperage ratings are stamped on thetop of each circuit breaker. Figures 2A-5through 2A-13 illustrate the circuit-breakerpanels by model and serial number.







AFCS DAMPER

RMI SECONDARY

VERTISYN DIRECTISYN

FUEL PUMPS

FUEL FLOW

FIRE DETECT

NOSE STEERING

LIGHTS

TRIM TAB

FIRE WALL PRESSURE

OIL PRESS

TURN & BANK

RADAR

AC DC GS NAV COMM AUDIO

CARD POINTER GS NAV COMM AUDIO

ADF ATC MKR BCN DME GYRO O PWR

PRI SEC PRI SEC

L R

FUELQTY

RAM AIRTEMP

AC BUSTIE

WARNLIGHTS GEAR

FLAPSTAB

OXYGENVALVE

ANTISKID

L TIP L ENG FUSELAGE

CROSSFLOW

R ENG R TIP

IGN &START

OILTEMP L R L R

L R L SHUT OFF R L RRATIO

RAM AIRTEMP

TURN & BANK

FUEL PUMPS

FUEL FLOW OIL PRESS OIL PRESS

AIR IGN

HYD PUMP

L TIP L ENG FUSELAGE

L

L L R RBUS TIE

L RRNAC ICEDETECT

CABINPRESS

CABINHEAT

WINDSHIELDDE-ICE

MIC

R ENG R TIP R TIPFUELQTY

ANTISKID

HYDPRESS

AC DC SPOILER YAW PITCH

WHEELMASTER

CABINHEAT

WINDSHIELDDE-ICE

ROLL

INST NAV CABIN

PITOT HEAT

MAIN DC POWER

NACELLE HEAT

FLOOD BCN

L RNAC ICEDETECT

CABINPRESS L R

L L BUS TIE R RMIC

PHONE

VOL

PRIMARY

2 5 2 2 10 2

2 2 2 2 10 2

2 2 2 10 2 7

2 2 2 2 2 2

2 2 2 2 2 5

10 10 10 10 10

10 2 2 2 2 2 2

10 10 10 10 10 10 10 2

IGN &START

OILTEMP L R L R L R

10

15 15 10

45 45 50 45 45

2 7 7

2 2 2 2 2 2 2

2 25

7 7 7 7 2 2

2 7 2 2 7 2

7 7 7 7 7

15 15 10 2 7 7

45 45 50 45 45

PITOT HEAT

MAIN DC POWER

NACELLE HEAT

1

1 2 4

2

3

4

5

A

A

3 C

5 D

3 B

A

BCDE

E

AIRPLANES SNs 23-003 THROUGH 23-014

AIRPLANES SNs 23-025 THROUGH 23-069AIRPLANES SNs 23-070 THROUGH 23-099


L R

Figure 2A-5. Circuit-Breaker Panels—Model 23

HYDPUMP

BLANK

BLANK BLANK BLANK BLANK BLANKDCVOLTMETER

L R

FUEL PUMPLH STANDBY OUTER VOLTAGE

LOCKOUT

EMERGENCYPITCH TRIM

DOORACTUATOR

FREONCONTROL

CABINBLOWER STALL WARNING

LEFT RIGHT

CABINDUMP

RADARDC

FUEL PUMPRH STANDBY

Figure 2A-6. Tail Compartment Circuit-Breaker Panel—Model 23




Figure 2A-7. Circuit-Breaker Panels—SNs 24-100 through 24-129

PRIMARY

PRISTEER

ACBUS TIE

115 VAC PRI

VERT

VERT DIRNOSESTEER

CABINPRESS

ACVOLTMETER

DIR AFCSRADARGYRO O

115 VAC SEC

COPILOT’S SUBPANEL

ARMREST CIRCUIT-BREAKER PANEL

AFT CIRCUIT-BREAKER PANEL MAIN CIRCUIT-BREAKER PANEL

AURALWARN

BLEEDAIR

CABINDUMP

CABINAIR FLOW

CABINHEAT

ALCOHOLPUMP

FL

7

FL

9

FL

29

FL

11

FL

15

FL

19

FL

20

FL

14

FL

12

FL

18

FL

10

FL

8

HYDPUMP

EMERGENCYPITCH TRIM

RADARAC

DCVOLTMETER

DOORACTUATOR

FREONCONTROL

CABINBLOWER

LEFT RIGHTSTALL WARN

TAXI

L R

RADARDC

OVER VOLTAGELOCKOUT

COPILOT GYROS

RMI SECONDARY

ATT

CARD POINTER GS NAV COMM AUDIO

ADF ATC MKR BCN

AIR IGN TURN & BANK

DMEEMERG

PWR

DE-FOG RAT L R L R

WARNLIGHTS

WARNLIGHTS GEAR FLAPS

OXYGENVALVE

ANTISKID

L STDBYPUMP

L MOTIVEVALVE

FUSPUMP FILL

XFER

R MOTIVEVALVE

R STDBYPUMP

IGN &START

OILTEMP

FUELFLOW

FUELQTY

PRESSURERATIO

L R

L R L RSHUT-OFF

FIRE DETECT FIREWALL OIL PRESSURE

TRIM TAB

L R

AFCSNOSE

STEERING SPOILER

LIGHTS

PITOT HEAT NACELLE HEAT

YAW PITCH

WHEELMASTER

ROLL

INST NAV CABIN FLOOD BCNTRIM TAB

POS

L RNAC ICEDETECT

MAX CABINAIR FLOW

MAIN DC POWER

L R

L L BUS TIE R RCABINHEAT

W/SDEICE

DIR GS NAV COMM AUDIO




115 VAC

115 VAC

PRIVOLT

METER

PILOT’SVERT

PILOT’SDIRECT

RADARGYRO O

AFCS NOSESTEER

CABINPRESS

RMI LEPR

LOIL

PRESS

26 VAC

26 VAC

SECVOLT

METER

COPILOTVERT

COPILOTDIRECT

RADAR REPR

ROIL

PRESS

COPILOT’S SIDE CIRCUIT-BREAKER PANEL

BAGGAGE COMPARTMENTCIRCUIT-BREAKER PANEL

MAIN CIRCUIT-BREAKER PANEL

5

5

5 5 2.5 2.5 2.5 2.5

7.5 10 7.5 10 2.5 2.5 2.5

5

DEFUEL

FUELL TIP

RADARDC

DOORACT

DCVOLTMETER

FREONCONTROL

CABINBLOWER

CABINLIGHT

NAC ICEDETECT

A/P STATICDRAIN

RAMAIR TEMP

CABINHEAT

L R & TAXIWING

ANTI-ICE

JETTISONR TIP

LDT LT

A

B

C

ABC



AIRPLANES 24-130 THROUGH 24-169

AUDIO

LEFT & PRIMARY RIGHT & SECONDARY

COMM NAV NAV COMM AUDIO

GS ATC DME MKR BCN ADF GS

TURN& BANK

AFCSDAMPER

WARNLIGHTS

WARNLIGHTS

TAB & FLAPPOSITION

TURN& BANK

FIREDETECT

FIREWALLSHUTOFF

ENG FIREEXT

ENG FIREEXT

FIREWALLSHUTOFF

FIREDETECT

NACELLEHEAT

PITOTHEAT

JET PUMPVALVE

JET PUMPVALVE

PITOTHEAT

STALL WARNHEAT

STALL WARNHEAT

NACELLEHEAT

FUELQTY

XFER& FILL

STDBYPUMP

STDBYPUMP

FUSPUMP

FUELFLOW

OILTEMP INSTR

LIGHTSNAV FLOOD BEACON

IGNSTART

ANTISKID DE-FOG

AIRBLEED

OXYVALVE

MAX CABINAIRFLOW

ALCOHOLPUMP

EMERPITCHTRIM PITCH

TRIM TABROLL YAW

BAT BUS

SHUTOFFVALVE LT

CABINDUMP

SPOILER GEAR FLAPSNOSESTEER L R

STALL WARN

L L BUS TIE

MAIN DC POWER

R RWING

ANTI-ICE

AIRIGN

AIRIGN

LIGHTS






AIRPLANES 24-170 THROUGH 24-180

AUDIO




TURN& BANK

AFCSDAMPER

WARNLIGHTS

WARNLIGHTS

TAB & FLAPPOSITION

TURN& BANK

FIREDETECT

FIREWALLSHUTOFF

ENG FIREEXT

ENG FIREEXT

FIREWALLSHUTOFF

FIREDETECT

NACELLEHEAT

PITOTHEAT

JET PUMPVALVE

JET PUMPVALVE

PITOTHEAT

STALL WARNHEAT

STALL WARNHEAT

NACELLEHEAT

FUELQTY

XFER& FILL

STDBYPUMP

STDBYPUMP

FUSPUMP

FUELFLOW

OILTEMP INSTR


IGNSTART

ANTISKID DE-FOG

AIRBLEED

OXYVALVE

MAX CABINAIRFLOW

ALCOHOLPUMP

EMERPITCHTRIM PITCH

TRIM TABROLL YAW

BAT BUS

SHUTOFFVALVE LT

CABINDUMP


STALL WARN

L L BUS TIE

MAIN DC POWER

R RWINGS

ANTI-ICE

AIRIGN

AIRIGN

LIGHTS




AUDIO



EMERPITCHTRIM PITCH ROLL

TRIM TAB

STALL WARN

BAT BUS

SPOILER GEAR FLAPS L R

YAWCABINDUMP


TURN& BANK

AFCSDAMPER

WARNLIGHTS

WARNLIGHTS

AFCSDAMPER

TURN& BANK

FIREDETECT

FIREWALLSHUTOFF

ENG FIREEXT

ENG FIREEXT

FIREWALLSHUTOFF

FIREDETECT

NACELLEHEAT

PITOTHEAT

JET PUMPVALVE

JET PUMPVALVE

PITOTHEAT

STALL WARNHEAT

STALL WARNHEAT

NACELLEHEAT

FUELQTY

XFER& FILL

STDBYPUMP

STDBYPUMP

FUSPUMP

FUELFLOW

OILTEMP INSTR


IGNSTART

ANTISKID DE-FOG

AIRBLEED

OXYVALVE

MAX CABINAIRFLOW

ALCOHOLPUMP

EMERPITCHTRIM PITCH

TRIM TABROLL YAW

BAT BUS

CABINDUMP

L L BUS TIE

MAIN DC POWER

R RWING

ANTI-ICE

AIRIGN

AIRIGN

LIGHTS


STALL WARN

AUDIO




TURN& BANK

AFCSDAMPER

WARNLIGHTS

WARNLIGHTS

TAB & FLAPPOSITION

TURN& BANK

FIREDETECT

FIREWALLSHUTOFF

ENG FIREEXT

ENG FIREEXT

FIREWALLSHUTOFF

FIREDETECT

NACELLEHEAT

PITOTHEAT

JET PUMPVALVE

JET PUMPVALVE

PITOTHEAT

STALL WARNHEAT

STALL WARNHEAT

NACELLEHEAT

FUELQTY

XFER& FILL

STDBYPUMP

STDBYPUMP

FUSPUMP

FUELFLOW

OILTEMP INSTR


IGNSTART

ANTISKID DE-FOG

AIRBLEED

OXYVALVE

MAX CABINAIRFLOW

ALCOHOLPUMP

EMERPITCHTRIM PITCH

TRIM TABROLL YAW

BAT BUS

CABINDUMP

SHUTOFFVALVE LT

L L BUS TIE

MAIN DC POWER

R RWING

ANTI-ICE

AIRIGN

AIRIGN

LIGHTS


STALL WARN

AIRPLANE SN 25-003 ONLY

AIRPLANES SNs 25-004THROUGH 25-009 AIRPLANE SN 25-010 ONLY






AUDIO




TURN& BANK

AFCSDAMPER

WARNLIGHTS

WARNLIGHTS

AFCSDAMPER

TURN& BANK

FIREDETECT

FIREWALLSHUTOFF

ENG FIREEXT

ENG FIREEXT

FIREWALLSHUTOFF

FIREDETECT

NACELLEHEAT

PITOTHEAT

JET PUMPVALVE

JET PUMPVALVE

PITOTHEAT

STALL WARNHEAT

STALL WARNHEAT

NACELLEHEAT

FUELQTY

XFER& FILL

STDBYPUMP

STDBYPUMP

FUSPUMP

TABS & FLAPPOSITION

OILTEMP INSTR


IGNSTART

ANTISKID DE-FOG

AIRBLEED

OXYVALVE

MAX CABINAIRFLOW

ALCOHOLPUMP

EMERPITCHTRIM PITCH

TRIM TABROLL YAW

BAT BUSSHUTOFFVALVE LT

CABINDUMP


STALL WARN

L L BUS TIE

MAIN DC POWER

R RWINGS

ANTI-ICE

AIRIGN

AIRIGN

LIGHTS

115 VAC

115 VAC

PRIVOLT

METER

PILOT’SVERT

PILOT’SDIRECT

RADARGYRO O

AFCS NOSESTEER

CABINPRESS

RMI LEPR

LOIL

PRESS

26 VAC

26 VAC

SECVOLT

METER

COPILOTVERT

COPILOTDIRECT

RADAR AFCS REPR

TONEGEN

ROIL

PRESS

COPILOT’S SIDE CIRCUIT-BREAKER PANEL(TYPICAL)


(TYPICAL)

5

5

5 5 2.5 2.5 2.5

7.5 10 7.5 10 2.5 2.5 2.5

5

DEFUEL

2.5

STEREO

FUELL TIP

RADARDC

DOORACT

DCVOLTMETER

FREONCONTROL

CABINBLOWER

CABINLIGHT

NAC ICEDETECT

AUTOCABINHEAT

RAMAIR TEMP

MAINCABINHEAT

L R & TAXIL ICEDET

R ICEDET

JETTISONR TIP

LDG LT

1 1 5 2.5

PRI SECDMESEC

ADFSEC

10

STROBELIGHTS

FLT DIR





SNs 24-181 THROUGH 24-189SNs 24-025 THROUGH 25-029

SNs 24-190 THROUGH 24-199SNs 25-030 THROUGH 25-034

AUDIO




TURN& BANK

AFCSDAMPER

WARNLIGHTS

WARNLIGHTS

TAB & FLAPPOSITION

TURN& BANK

FIREDETECT

FIREWALLSHUTOFF

ENG FIREEXT

ENG FIREEXT

FIREWALLSHUTOFF

FIREDETECT

NACELLEHEAT

PITOTHEAT

JET PUMPVALVE

JET PUMPVALVE

PITOTHEAT

STALL WARNHEAT

STALL WARNHEAT

NACELLEHEAT

FUELQTY

XFER& FILL

STDBYPUMP

STDBYPUMP

FUSPUMP

FUELFLOW

OILTEMP INSTR


IGNSTART

ANTISKID DE-FOG

AIRBLEED

OXYVALVE

MAX CABINAIRFLOW

ALCOHOLPUMP

EMERPITCHTRIM PITCH

TRIM TABROLL YAW

BAT BUSSHUTOFFVALVE LT

CABINDUMP


STALL WARN

L L BUS TIE

MAIN DC POWER

R RWINGS

ANTI-ICE

AIRIGN

AIRIGN

LIGHTS

AUDIO




TURN& BANK

AFCSDAMPER

WARNLIGHTS

WARNLIGHTS

TAB & FLAPPOSITION

TURN& BANK

FIREDETECT

FIREWALLSHUTOFF

ENG FIREEXT

ENG FIREEXT

FIREWALLSHUTOFF

FIREDETECT

NACELLEHEAT

PITOTHEAT

JET PUMPVALVE

JET PUMPVALVE

PITOTHEAT

STALL WARNHEAT

STALL WARNHEAT

NACELLEHEAT

FUELQTY

XFER& FILL

STDBYPUMP

STDBYPUMP

FUSPUMP

NOSESTEER

OILTEMP INSTR


WINGHEAT

WSHLDHEAT

AIRBLEED

OXYVALVE

MAX CABINAIRFLOW

ALCOHOLPUMP

ANTISKID

SHUTOFFVALVE LT

TRIM TABEMERPITCHTRIM YAWROLLPITCH

SPOILER L R GEAR FLAPSTAB & FLAPPOSITION

LEFT IGNSTART-GEN

LEFTBUS BUS TIE

RIGHTBUS

RIGHT IGNSTART-GEN

MAIN DC POWER

AIRIGN

AIRIGN

LIGHTS

STALL WARN

BAT BUS

Figure 2A-12. Circuit-Breaker Panels—SNs 24-181 through 24-199,and 25-025 through 25-034




MAIN CIRCUIT-BREAKER PANEL

GS ATC DMEMKRBCN ADF GS

AUDIO COMM NAV NAV COMM AUDIO

AIRIGN

AIRIGN

STABHEAT

TURN& BANK

AFCSDAMPER

WARNLIGHTS

WARNLIGHTS

AFCSDAMPER

TURN& BANK

FIREDETECT

FIREWALLSHUTOFF

ENG FIREEXT

ENG FIREEXT

FIREWALLSHUTOFF

FIREDETECT

NACELLEHEAT

PITOTHEAT

STALL WARNHEAT

STALL WARNHEAT

JET PUMPVALVE

JET PUMPVALVE

PITOTHEAT

NACELLEHEAT

FUELQTY XFER & FILL

LIGHTS

STDBYPUMP

STDBYPUMP FUS. PUMP

NOSESTEER

OILTEMP

WINGHEATINSTR INSTR BEACONNAV

WSHLDHEAT

AIRBLEED

STALL WARN

MAX CABINAIRFLOW

OXYVALVE

ALCOHOLPUMP

SPOILER LEFT RIGHT GEAR FLAPS

BUS TIE

MAIN DC POWER

RIGHTBUS

RIGHT IGNSTART-GEN

LEFTBUS

LEFT IGNSTART-GEN

TAB & FLAPPOSITION

LIGHTS

BATTERYSHUTOFFVAL LTS

BUSEMERP TRIM

PITCH

ANTISKID ROLL YAW

NORMAL TRIM

115 VAC

115 VAC

PRIVOLT

METER

PILOT’SVERT

PILOT’SDIRECT

RADARGYRO O

AFCS NOSESTEER

CABINPRESS

LEPR

RAM LOIL

PRESS

26 VAC

26 VAC

SECVOLT

METER

COPILOTVERT

COPILOTDIRECT

RADAR AFCS TONEGEN

REPR

ROIL

PRESS

COPILOT’S SIDE CIRCUIT-BREAKER PANEL(TYPICAL)


(TYPICAL)

STEREO PCIHF

COMMHP

COMMATCSEC

CIGARLIGHTER

AUXCOMM

FUELL TIP

JETTISONR TIP

EMERBAT CHG

DMESEC

ADFSEC

STROBELIGHTS

RADARDC

DOORACT

DCVOLTMETER

FREONCONTROL

CABINBLOWER

CABINLIGHT

RAMAIR TEMP

CABINHEAT

L R & TAXIL ICEDET

R ICEDET

LDG LT

AIRPLANES SNs 24-200 AND SUBSEQUENTAIRPLANES SNs 25-035 AND SUBSEQUENT

Figure 2A-13. Circuit-Breaker Panels—SNs 24-200 and Subsequent,and 25-035 and Subsequent

Figure 2A-14. DC Power Distribution—SNs 23-003 through 24-189,and 25-003 through 25-029

50A 40A

LH FEEDERBUS

AMMETER

AMMETER

RHGEN

V

LHGEN

RHBATTERY

LHBATTERY

EXTERNALPOWER

RH GENBUS

BATTERYBUS

BATTERYCHARGING

BUS

LH GENBUS

275A

275A

STARTRELAY

STARTRELAY

50A 40A

NACELLEHEAT

BATTERYRELAY

REVERSECURRENT

RELAY

REVERSECURRENT

RELAY

FREON150A

NACELLEHEAT

RH FEEDERBUS

BUS TIE50A

APPLICABLE: SNs 23-003 THROUGH 24-189 SNs 25-003 THROUGH 25-029

EXCEPT NAC HT ON: SNs 24-181 THROUGH 24-189 SNs 25-025 THROUGH 25-029




DISTRIBUTIONThe power distribution system functions as asingle system, but it is really separate systems in-terconnected by a series of current limiters andcircuit breakers. Figures 2A-14 and 2A-15 illus-trate power distribution. It is possible to poweralmost all DC electrical equipment from onegenerator through the interconnected buses oronly certain buses by operation of the bus and/orbus-tie circuit breakers. Starting with SNs 24-130through 24-229, and 25-003 through 25-064, allcircuit breakers are accessible to the crew dur-ing flight. There are certain hot-wired circuitsconnected to the batteries and are as follows:

• Left stall warning (all models)

• Right stall warning (all models)

• Cabin door motor (all models)

• Entry lights (door step light, center aislel ight[s] , and baggage compartmentlights) (all models)

• Firewall shutoff valve warning lights(airplanes equipped with glareshieldwarning lights)

• Left standby fuel pump (23 models with-out jet pumps)


• Right standby fuel pump (23 modelswithout jet pumps)

• Freon and cabin fan systems (SNs 23-002 through 23-069)

• No smoking/fasten seat belt light (someairplanes)

• Emergency pitch tr im (SNs 24-130through 24-229, and 25-003 through25-064)

• Cabin dump valve (SNs 24-130 through24-180, and 25-003 through 25-026, and25-028)

Battery BusSNs 23-003 through 24-189, and 25-003 through 25-029The battery bus is connected directly to theGPU receptacle and both batteries. The bat-teries may be charged by a GPU with the bat-t e ry sw i t ch i n OFF. The ba t t e ry bus i sconnected to the battery charging bus by turn-ing on the battery switch. This also permits ei-ther generator to charge the batteries. Withthe battery switch in OFF, the generators sup-ply power to all airplane equipment exceptthe batteries and hot-wired items.



Figure 2A-15. DC Power Distribution—SNs 24-190 through 24-229,and 25-030 through 25-064 except 25-061

AMMETER

AMMETER

RHGEN

DCVOLTMETER

OVERLOAD

SENSOR

OVERLOAD

SENSOR

LHGEN

RHBATTERY

LHBATTERY

EXTERNALPOWER

EXT. POWEROVERVOLTAGE

RH GENBUS

BATTERYBUS

BATTERYCHARGING

BUS

RH FEEDERBUS RELAY

LH FEEDERBUS RELAY

LH GENBUS

275A

2ARH IGN START

2ALH IGN START

2ARH BUS

2ALH BUS

275A

STARTRELAY

STARTRELAY

10A

REVERSECURRENT

DIODE

REVERSECURRENT

DIODE

33 – 2 VDC

NACELLEHEAT

BATTERYRELAY

FL3275A

NACELLEHEAT

+

START/GENSWITCH

FREON150A

START/GENSWITCH

STAB HEAT130A

RHFEEDER

BUS

LHFEEDER

BUS

BUS TIE50A

EFFECTIVE: SNs 24-190 AND SUBSEQUENT SNs 25-030 AND SUBSEQUENT

SNs 24-190 through 24-229,25-030 through 25-064 except 25-061The battery bus may be powered by externalpower if the voltage from the GPU is 33 ±2VDC or less. The batteries are connected to thebattery charging bus when the battery switchis placed to ON. As long as the battery switchis in ON, either engine generator charges bothbatteries. The engine generators supply allDC equipment on the airplane when the bat-tery switch is in OFF except the batteries andthe hot-wired items.

Battery Charging BusThe battery charging bus is the central distri-bution point of the DC power from the batter-ies , a GPU, or the generators to variousairplane systems. The DC voltmeter reads di-rectly from this bus. The battery switch con-nects the batteries to the battery charging busthrough the battery relay.

With the battery switch turned on, the entireelectrical system may be powered for a lim-ited time. Exceptionally high amperage equip-ment shortens this time considerably. A relayisolates the Freon air conditioner and auxil-iary heater until a GPU is connected or a gen-erator is on and operating.

Generator BusesThe left and right generator buses are in thecurrent limiter panel in the tailcone. The gen-erator buses may be powered by the batterycharging bus or by the output of the respectivegenerator.

The 275-ampere current limiters connect thegenerator buses to the battery charging bus.

Feeder BusesThe left and right feeder buses receive powerfrom the respective generator buses or from theopposite generator and feeder bus through the50-amp DC bus-tie circuit breaker. The leftfeeder bus provides power through circuitbreakers for primary navigation, communi-cations, instruments, and one-half the DCloads. The right feeder bus provides powerthrough circuit breakers for secondary navi-gation, communication, instruments, and theremaining half of the DC loads.

Each bus is protected by two circuit breakersand two current limiters. On airplanes SNs24-190 and 25-030 and subsequent, the feederbuses are protected by a single bus circuitbreaker and overload sensor.




GROUND POWERThe external power receptacle (Figure 2A-16)is located immediately forward of the tail com-partment access door on SNs 23-003 through24-129. On all other Learjets, it is located onthe left side of the fuselage adjacent to the leftengine tailpipe. It is not necessary to operateany airplane switches to apply external powerto the batteries or battery bus. On SNs 24-190through 24-229, and 25-030 through 25-064,the external power system incorporates an ex-ternal power overvoltage cutout circuit con-tained within the generator control box.

The overvoltage circuit prevents the GPU frompowering the airplane electrical system if GPUvoltage exceeds 33 ±2 volts. GPU capabilityshould be limited to 28 VDC and a maximumof 1,000 amperes. The 1,000-ampere limit isset because of starter shaft torque limits. Whenusing a GPU for engine starting, a minimumof 800 amperes is required.

AC POWER

INVERTERS

SNs 23-003 through 24-129Two JET solid-state static inverters, locatedin the tail compartment, provide 115-VAC,400-Hz, single-phase power (Figure 2A-17).Through a separate stepdown transformer, 26-VAC power is also provided. Each inverter israted at 750 VA. During normal operation, themain inverter, through the closed AC bus-tiecircuit breaker, operates all AC-powered equip-ment except the radar. When the radar is in ei-ther standby position or a range setting, thestandby inverter automatically comes on tosupply power just for the radar.




Figure 2A-17. AC Power Distribution—SNs 23-003 through 24-129

Figure 2A-16. External Power Receptacle

LIMITERS

LIMITERS

DC BUS

DC BUS

RADARFUEL FLOW R

EPR RHYD PRESSURE

EPR LOIL PRESS ROIL PRESS L

RMI CARDRMI POINTERFUEL FLOW L

AC VOLTMETER

26 VACBUS

115 VAC(PRI)

115 VAC(SEC)

AFCSPRI DIRECTISYNPRI VERTISYNRADAR GYRO O

AC BUS TIE

FUEL QUANTITYSEC DIRECTISYNSEC VERTISYNNOSEWHEEL STEERINGCABIN PRESSUREADF

MAININVERTER

STDBYINVERTER

POWER RELAY

SWITCHINGRELAY

MAIN

STANDBY

OFF

SNs 24-130 through 24-229, and 25-003 through 25-064Three JET solid-state static inverters locatedin the tail compartment provide 115-VAC and26-VAC, 400-Hz, single-phase power (Figures2A-18 and 2A-19). PRI and SEC inverters arerated at 400 VA, and the standby inverter israted at 750 VA. During normal operation, thePRI inverter supplies the left AC equipment,the secondary inverter supplies the right ACequipment, and the standby inverter suppliesthe radar. The standby inverter automaticallypicks up the load in case of a primary or sec-ondary inverter failure.

CONTROLS

SNs 23-003 through 24-129One three-position MAIN–OFF–STANDBYinverter switch is located on the switch panel.

SNs 24-130 through 24B-229,and 25-003 through 25-064Three ON–OFF inverter switches for the pri-mary, secondary, and standby inverters are lo-cated on the switch panel.




Figure 2A-18. AC Power Distribution—SNs 24-130 through 24-180,and 25-003 through 25-024

PRIINVERTER

STDBYINVERTER

SECINVERTER

DC PWRL GEN BUS

DC PWRR GEN BUS

BATCHG BUS

PRIINVERTER

SWITCH

STDBYINVERTER

SWITCH

SECINVERTER

SWITCH

PRIINVERTERPOWERRELAY

SECINVERTERPOWERRELAY

STDBYINVERTERPOWERRELAY

400VA

400VA

750VA

TO INDICATORLIGHT

TO INDICATORLIGHT

TO INDICATORLIGHT

115 VAC115 VAC

115 VAC

RADAR RELAY

115 VAC

26 VAC

26 VAC

AC SWITCHING

115 VAC

115 VAC

26 VAC

26 VAC

26 VAC

PRI VOLTMETERPILOT VERTPILOT DIRECTRADAR GYROAFCSNOSE STEERCABIN PRESSRMI

L EPRL OIL PRESS

RADAR

SEC VOLTMETERCOPILOT VERTCOPILOT DIRECT

R EPRR OIL PRESS




Figure 2A-19. AC Power Distribution—SNs 24-181 through 24-229,and 25-025 through 25-063, except 25-061

PRIINVERTER

STDBYINVERTER

SECINVERTER

DC PWRL GEN BUS

DC PWRR GEN BUS

BATCHG BUS

PRIINVERTER

SWITCH

STDBYINVERTER

SWITCH

SECINVERTER

SWITCH

PRIINVERTERPOWERRELAY

SECINVERTERPOWERRELAY

STDBYINVERTERPOWERRELAY

400VA

400VA

750VA

TO INDICATORLIGHT

TO INDICATORLIGHT

TO INDICATORLIGHT

115 VAC115 VAC

115 VAC

RADAR RELAY

115 VAC

26 VAC

26 VAC

AC SWITCHING

115 VAC

115 VAC

26 VAC

26 VAC

26 VAC

PRI VOLTMETERPILOT VERTPILOT DIRECTRADAR GYROAFCSNOSE STEERCABIN PRESSRMIEPR

L OIL PRESSANTI-SKID TEST

RADAR

SEC VOLTMETERCOPILOT VERTCOPILOT DIRECTAFCSEPRNOSE STEERING*

TONE GENR OIL PRESS

*NOSE STEERING EFFECTIVE SNs 24-209 AND 25-041 AND SUBSEQUENT IS NOW ON THE COPILOT’S 115-VAC BUS.

INDICATORSOn SNs 24-130 through 24-229, and 25-003through 25-064, three amber inverter warninglights are located on the switch panel or theglareshield. If an inverter fails, the respectivewarning light illuminates. If the inverter switchis OFF, the inverter fail light is extinguished.

OPERATION

SNs 23-003 through 24-129AC power is normally provided by the maininverter. A standby inverter is also installed tosupply AC power in case of a main inverter fail-ure. AC output from the main inverter is ap-plied via the inverter switching relay to a26-VAC stepdown transformer and to the pri-mary and secondary buses. The radar systemis operated by the standby inverter exclusivelywhen the main inverter and the radar are bothturned on. The 26-VAC stepdown transformersupplies voltage to the AC instrument bus.The primary and secondary buses are con-nected by the AC bus-tie circuit breaker.

Emergency OperationIn case of a main inverter failure, the inverterswitch must be set to standby. With the in-verter switch in this position, the standby in-verter relay and the inverter switching relayare energized. This allows the standby inverterto supply AC voltage to all systems with theexception of the radar. The radar is inopera-tive when the switch is in the standby position.

Buses—Primary and SecondaryThe primary and secondary AC buses are in-terconnected by the AC bus-tie circuit breaker.The primary AC bus powers the No.1 vertisynand directisyn yaw damper, AC voltmeter, andthe radar gyro phase detection. The secondarybus powers the No. 2 vertisyn and directisyn,nosewheel steering, cabin pressurization, andsometimes fuel quantity.

Stepdown TransformersThe stepdown transformers, located on theforward side of frame 28, reduce a portion ofthe 115 VAC to 26 VAC. The 26 VAC is usedfor the antiskid test (SNs 24-112 and subse-quent), oil pressure gages, EPR gages, andfuel flow on SNs 23-003 through 23-069.

Inverter Switching RelayThe inverter switching relay, located in the tailcompartment, is energized when the inverterswitch is set to STANDBY. When energized, therelay allows AC voltage from the standby in-verter to be distributed to the AC equipment.

IndicationThe AC voltmeter displays the voltage for theselected inverter.

SNs 24-130 through 24-229,and 25-003 through 25-063,except 25-061AC power for the pilot’s flight instruments isprovided by the primary inverter while thecopilot’s instruments are powered by the sec-ondary inverter. The radar is powered by thestandby inverter. A switching network allowsthe standby inverter to automatically supply115 and 26 VAC to the bus of a failed primaryor secondary inverter. An indicator light abovethe inverter switch (green on SNs 24-130,133,and 135 through 138) indicates that the in-verter is on. An amber indicator light on allother Learjets is located either above the in-verter switch or on the glareshield warninglights panel. It indicates that the inverter hasfailed when the control switch is in ON. TheAC voltmeter selector switch, installed justabove the AC voltmeter, is used to monitor the115-VAC output of the inverters.




EMERGENCY MODESIn emergency operation, three different modesof inverter operation are possible: (1) With aprimary inverter failure, the standby inverter,through a switching network, automaticallyprovides power to the primary 115-VAC and26-VAC buses. Secondary inverter operationis not affected. (2) With a secondary inverterfailure, the standby inverter, through a switch-ing network, automatically provides power tothe secondary 115-VAC and 26-VAC buses.Primary inverter operation is not affected. (3)With both primary and secondary invertersfailed, the standby inverter, through a switch-ing network, automatically provides power tothe pilot’s 115-VAC and 26-VAC buses and thecopilot’s 26-VAC bus. Operation of the radaris not affected by any of the emergency modes.With a primary and secondary inverter failure,the copilot’s 115-VAC bus is deenergized,and, because this bus powers the pilot’s RMIcard, it ,too, is inoperative.

EMERGENCY BATTERYSYSTEMAn emergency battery is probably installedin most model 23 airplanes. SNs 24-100through 24-129 have an auxiliary battery in-stalled. On all other airplanes, an optionalemergency battery is available.

Auxiliary Battery System—SNs 24-100 through 24-129Auxiliary battery power (Figure 2A-20) issupplied by a 2.6-ampere-hour, 25-volt, nickel-cadmium dry-cell battery. The battery, in-stalled beneath the divan seat, consists of asteel case containing 20 cells and a 50-VA in-verter. The inverter provides 115-VAC, 400-Hz, single-phase power. The battery receivesa trickle charge from the battery charging bus.Power is controlled by an ON–OFF switchand a momentary switch.




Figure 2A-20. Auxiliary Battery System—SNs 24-100 through 24-129

RIGHTFWDBUS

OFF ON (MOM)

GEAR, FLAPSPOILER SWITCH

TRICKLE CHARGE

ATTITUDE GYRO

DIRECTIONAL GYRO

AUXILIARYBATTERY

AND INVERTER

TOINVERTER

GEAR CONTROL

FLAPS

SPOILER

DIRECTIONAL GYRO POST LIGHTATTITUDE GYRO POST LIGHT

TRANSCEIVERTRANSCEIVER LIGHT

PILOT’S AUDIOCOPILOT’S MAP LIGHT

ADF

WHEEL MASTER

NOSE STEERING RELAY

YAW TRIM

ROLL TRIM

DC POWER (STANDARD ON: 24-100THROUGH 24-129 AC POWER (OPTIONAL ON ALLOTHER LEARJETS

LEGENDOFF

GYROSON

The battery may be used to power the following:

• Emergency attitude gyro and post light

• Gears

• Flaps

• Spoilers

• Emergency directional gyro and postlight

• Auxiliary communications and internallight

• Pilot’s audio panel

• Copilot’s map light

• ADF

• Norm pitch trim

• Roll trim

• Yaw trim

• Nose steering

OperationUnder normal conditions, 28 VDC is beingsupplied from the battery charging bus to theemergency battery. The 28 VDC provides atrickle charge for the emergency battery. Whenthe AUX. GYROS switch is placed to ON, 25

VDC from the emergency battery is applied tothe pilot’s audio panel, ADF, extra transceiver,nosewheel steering relay, wheel master switch,and trim systems (including the self-containedinverter). The inverter system provides ACpower to the attitude and directional gyros.

Emergency Operation Under an emergency mode of operation, theauxiliary battery, through a diode network,provides power to the pilot’s audio panel,ADF, transceiver, nosewheel steering relay,wheel master button, directional and attitudegyros, trim systems, copilot’s map light, etc.The momentary switch must be held to theON position when power is needed to operatethe gear, flaps, and spoilers. With all the aboveequipment operating, the auxiliary batterylasts approximately 15 minutes.

Optional EmergencyBattery SystemAn optional dry-cell nicad battery is providedto power the standby attitude indicator througha self-contained inverter and also to power anauxiliary communications transceiver and thegear, flaps, and spoiler solenoids (Figure 2A-21). When illuminated, an amber light next tothe emergency battery switch indicates that theemergency battery is supplying power.







GEAR

FLAPS

SPOILERS

AUX COMM

EMERGENCYBATTERY C/B

28 VDC25 VDC

25-VDC INPUT

115-VAC OUTPUT

4.6-VAC OUTPUT

ON

OFFBATTERY

INV

INSTRUMENT LIGHTS

FEEDER BUS

INPUT

ATTITUDE GYRO

INPUT POWER

INTERNAL LIGHTS

Figure 2A-21. Emergency Battery System

1. The voltage and amperage rating of theengine-driven generator is:A. 24 VDC and 400 amperesB. 28.5 VDC and 300 amperesC. 28.5 VDC and 400 amperesD. 30 VDC and 400 amperes

2. Each battery is rated at:A. 24 VDCB. 25 VDCC. 28 VDCD. 30 VDC

3. The maximum GPU output that may beused for an engine start is:A. 400 amperesB. 600 amperesC. 1,000 amperesD. 1,100 amperes

4. Generator current flow in amperes ismeasured from the:A. Battery charging busB. Generator field coilC. Generator busD. Feeder bus

5. The DC voltmeter draws its readingfrom the:A. Battery busB. Battery charging busC. Generator busD. Feeder bus




QUESTIONS

3-i

CHAPTER 3LIGHTING

CONTENTS

Page

INTRODUCTION ................................................................................................................... 3-1

GENERAL............................................................................................................................... 3-1

INTERIOR LIGHTING .......................................................................................................... 3-2

Cockpit Lighting .............................................................................................................. 3-2

Cabin Lighting ................................................................................................................. 3-3

Emergency Lighting......................................................................................................... 3-4

EXTERIOR LIGHTING ......................................................................................................... 3-4

General ............................................................................................................................. 3-4

Landing–Taxi Lights ........................................................................................................ 3-6

Recognition Lights (Not Installed on All Airplanes)....................................................... 3-7

Navigation Lights............................................................................................................. 3-7

Anticollision Lights ......................................................................................................... 3-8

Wing Inspection Light (Optional).................................................................................... 3-8

QUESTIONS ........................................................................................................................... 3-9




3-iii

ILLUSTRATIONS

Figure Title Page

3-1 Cockpit Lighting Controls........................................................................................ 3-2

3-2 Map Lights (Typical)................................................................................................ 3-2

3-3 Overhead Lights Control (Typical) .......................................................................... 3-3

3-4 Advisory Lights and Control.................................................................................... 3-3

3-5 Emergency Lights Switch ........................................................................................ 3-4

3-6 Exterior Lighting Locations ..................................................................................... 3-5

3-7 Exterior Lighting Switches ...................................................................................... 3-6

3-8 Landing–Taxi Lights ................................................................................................ 3-6

3-9 Recognition Light..................................................................................................... 3-7

3-10 Strobe and Navigation Lights................................................................................... 3-7

3-11 Anticollision Lights.................................................................................................. 3-8




INTRODUCTIONAirplane lighting for the Lear 20 series airplanes is divided into interior, exterior, andemergency (if installed) lighting packages. Interior lighting provides illumination of boththe cockpit and cabin areas under normal conditions. The cockpit area is provided withgeneral illumination and specific lighting for instruments and map reading. Cabin arealighting provides illumination for the standard warning signs and specific area illumi-nation for passenger safety and convenience. Exterior lighting consists of the standardrequired lights, and optional packages are available for the wing and tailcone areas.

An optional emergency lighting system may be installed to illuminate the cabin interiorand egress points in the event of airplane electrical power failure.

GENERALCockpit lighting consists of the instrumentlights, floodlight, panel lighting, and maplights, all adjustable for intensity with rheo-stat controls. The panel lighting illuminates thelettering on the various switch panels, pedestal,and circuit-breaker panels. Side panel maplights are installed for the pilot’s convenience.Cabin lighting consists of upper center panellights, door entry lights, baggage compart-ment lights, individual reading lights, and the

No Smoking/Fasten Seat Belts sign. The op-tional emergency lighting systems illuminatethe upper center panel lights and other lightsat the exits. Exterior lights include landing–taxilights, wing and tail navigation lights, anticol-lision beacons, recognition lights, and high-in-tens i ty s t robe l igh t s . An op t iona l winginspection light illuminates the right wing areato check for ice accumulation or egress.

CHAPTER 3LIGHTING




EXIT

INTERIOR LIGHTING

COCKPIT LIGHTING

GeneralControls for lighting are either on the deviceor as illustrated in Figure 3-1.

Instrument Panel Post andEyebrow LightsThe pilot’s and copilot’s instrument panel postlights and eyebrow lights are individually con-trolled by separate rheostat switches that havean off detent. The pilot’s switch is located onthe cockpit left side immediately aft of thepilot’s subpanel. The copilot’s switch is locatedon the cockpit right side. A third switch, on thecockpit left side, controls the center instrumentpanel lights. The lights are powered by lightdimmer assemblies through the rheostatswitches. The light dimmer assemblies are lo-cated in the nose compartment.

Instrument Panel FloodlightsThe pilot’s and copilot’s floodlights, locatedunder the glareshield, are individually con-trolled by separate on–off rheostat switches.

The pilot’s controls on the left side, and thecopilot’s, on the right side of the cockpit,control separate sections of the floodlights.The lights are powered through the INSTRLT circuit breakers.

Map Reading LightsThe map reading lights, located on the leftside of the pilot and right side of the copilot,are controlled by rheostat switches adjacentto the lights (Figure 3-2). The lights are pow-ered through the LH and RH INSTR LT cir-cuit breakers.




Figure 3-1. Cockpit Lighting Controls

Figure 3-2. Map Lights (Typical)

CABIN LIGHTING

GeneralPassenger compartment lighting consists ofreading lights, overhead lights, entry lights,and No Smoking/Fasten Seat Belt signs.

Upper Central Panel LightsTwo upper center panel lights are installedand controlled by the switch on the left serv-ice cabinet (Figure 3-3). The lights are pow-ered through the ENT LTS and CAB LTScircuit breakers.

Entry LightThe entry light, installed in the left service cab-inet, is controlled by the switch on the left serv-ice cabinet. The light is powered through theENT LTS circuit breaker. Battery power isavailable to the light even when the batteryswitches are turned off.

No Smoking/Fasten Seat BeltSignThe lights are located on the aft side of the leftservice cabinet (Figure 3-4) and are controlled

by a switch on the pedestal. In addition to en-ergizing the lights, the control switch applies28 VDC to the tone generator, which producesan audible tone through the aft cabin speaker.The lights are powered through the CAB LTScircuit breaker.

Baggage Compartment LightsTwo baggage compartment lights are installedin the airplane. They are located in the head-liner of the baggage compartment. The lightsare controlled by the switch located on theleft service cabinet. The lights are poweredthrough the ENT LTS circuit breaker.




Figure 3-3. Overhead Lights Control(Typical)

Figure 3-4. Advisory Lights and Control

EMERGENCY LIGHTING

GeneralOn airplanes Models 24-255 and subsequent,25-090, and 25-094 and subsequent, and thosemodified by AAK 73-6, emergency cabin light-ing is provided by utilizing the upper centralpanel lights. If a depressurization conditionshould occur, the oxygen system aneroidswitch completes a 28-VDC circuit to the de-pressurization relay control circuit. The con-trol circuit causes the upper center panel lightsto illuminate regardless of light control switchposition. An optional emergency exit and winginspection lighting system may be installed toprovide emergency lighting.

Egress LightingGeneralIf the optional emergency lighting package isinstalled, the airplane is equipped with twoemergency exit and wing inspection powersupplies. Each power supply consists of twosealed nickel-cadmium batteries and a controlcircuit module. One power supply is installedin the left service cabinet. The second powersupply is located on the cabin floor beneath thedivan seat. The wing inspection light is in-stalled on the right side of the airplane adja-cent to the lower forward corner of theemergency exit window. The emergency exitlight is installed in the upper cabin door. Thelight illuminates the lower cabin door and theimmediate area around the door. The coldcathode lights on the right side of the uppercenter panel and the optional forward left andright upper center panel lights are powered bythe emergency battery.

ControlThe EMER LIGHT TEST switch (Figure 3-5)provides the test function for the system and forautomatic illumination of the emergency lightsin the event of an interruption of normal DC elec-trical power. The switch has three positions:TEST, ARM, and DISARM. Setting the switchto TEST simulates a failure of normal DC elec-trical power and illuminates the upper cabinentry door light, the emergency exit light, and

the cabin overhead lights. Setting the switch toARM will arm the system to illuminate theemergency lights in the event of a failure ofnormal DC electrical power. With power on theairplane, setting the switch to DISARM iso-lates the emergency lights from the emergencybatteries. The switch should be set to ARMprior to takeoff. If the switch is in the DISARMposition and at least one BAT switch is on, theamber light adjacent to the switch will illumi-nate to remind the pilot that the switch shouldbe set to ARM. The switch should be set to DIS-ARM prior to setting the BAT switches to OFF.

The WING INSPECTION light switch (in-cluded as part of the emergency lighting sys-tem), located adjacent to the EMER LIGHTTEST–ARM–DISARM switch, may be used in-dependently of the emergency lighting systemto visually check for ice accumulation on theright wing leading edge. Turning the switch onilluminates the exterior wing inspection light.

EXTERIOR LIGHTINGGENERALThe exterior lighting systems consist of thelanding–taxi lights, navigation lights, anticol-lision lights, recognition lights, strobe lights, andan optional right wing inspection light (Figure3-6). The exterior lighting switches are locatedon the instrument panel (Figure 3-7).




Figure 3-5. Emergency Lights Switch




NAVIGATION LIGHTANTICOLLISION LIGHT

ANTICOLLISION LIGHT

WING INSPECTION LIGHT

STROBE LIGHT

STROBE LIGHT

LANDING-TAXI LIGHTSRECOGNITION LIGHT

NAVIGATION LIGHT

NAVIGATION LIGHT

STROBE LIGHT

Figure 3-6. Exterior Lighting Locations

LANDING–TAXI LIGHTSThe landing light system consists of one 450-watt lamp mounted on each main landing gearstrut (Figure 3-8), one 20-amp current limiterfor each side, relays, dimming resistors, andthe L and R LDG LT switches. The L and Rlanding light switches have three positions:OFF, TAXI LT, and LDG LT. DC power to op-erate the relays comes from the left and rightmain buses, respectively.

Setting the L or R LDG LT switch to TAXI LTcloses a relay which shunts DC power from therespective generator bus through a resistorwhich limits current to the lamp element.Moving the switch to LDG LT closes a secondrelay, allowing current flow to bypass the re-sistor, thereby increasing the brightness of thelamp. The 20-amp current limiters protect the




Figure 3-7. Exterior Lighting Switches

Figure 3-8. Landing–Taxi Lights

power circuits between the respective gener-ator bus and lamp filament.

On SNs 23-003 through 23-084, 24-140 andsubsequent, the landing lights will not illumi-nate unless the respective landing gear downand locked switch is closed and provides aground. On SNs 23-085 through 24-139, thelanding lights are disabled in flight by thegear-up trunnion switches.

NOTEIt is recommended that the lights beoperated in the L and R LDG LT po-sit ions as sparingly as possible.Lamp service life is shortened in theLDG LT position because of thehigher current flow.

RECOGNITION LIGHTS (NOTINSTALLED ON ALLAIRPLANES)A 250-watt recognition light is installed in thenose of the right tank or in both tanks (Figure3-9). The light is controlled with the RECOGswitch. When turned on, DC power, appliedthrough the RECOG LT circuit breaker, closesa control relay and connects power through a30-ampere current limiter to each light.

STROBE LIGHTS

The strobe light system consists of one strobelight mounted inside each wingtip navigationlight fixture (Figure 3-10) and one in the tail,a power supply for each strobe, a STROBEswitch, a DC STROBE LTS circuit breaker, anda timing circuit module that causes the strobesto flash. Each power supply is protected by aninternal 3-ampere fuse.

NAVIGATION LIGHTSThe navigation light system consists of onelamp in the outboard side of each tip tank(Figure 3-10), two lamps in the upper aft tailfairing, a NAV LT switch, and a NAV LTS cir-cuit breaker on the left main bus.

All three navigation lights are controlled bythe NAV LT switch. Additionally, position-ing the NAV LT switch on dims the landinggear safe/unsafe lights and the flight directorannunciator lights.




Figure 3-9. Recognition Light

Figure 3-10. Strobe and Navigation Lights

ANTICOLLISION LIGHTSAnticollision lights are installed on top of thevertical stabilizer and on the bottom of thefuselage (Figure 3-11). The lights are con-trolled by a BCN LT switch. Each light is adual-bulb light, and each bulb oscillates 180°at 45 cycles per minute. The beam is concen-trated by an integral lens, and an illusion of90 flashes per minute occurs due to the oscil-lation. The lights operate on DC power throughthe BCN LT circuit breaker.

WING INSPECTION LIGHT(OPTIONAL)A light is installed on the right side of thefuselage adjacent to the lower forward corner

of the emergency exit window. This light is de-signed to illuminate the leading edge of theright wing. The light may be installed as partof the emergency lighting system. The WINGINSPECTION control switch is located on theemergency lighting panel.

Models 23, 24, 24B, and 25 airplanes vary tosome extent in the location of switches andpower source for some lighting. The switch forthe center panel and baggage compartmentlights is on the aft side of the cabin door andis powered by the DOOR ACT circuit breaker.




Figure 3-11. Anticollision Lights

1. The instrument panel floodlight controlis located:A. Just forward of the warning panelB. On the pilot’s side panelC. On the copilot’s side panelD. Both B and C

2. The cockpit map lights are controlled:A. With an ON–OFF switch on the

copilot’s side panelB. With the overhead map light rheo-

stat on the copilot’s side panelC. With a collocated rheostatD. Automatically, relative to ambient

light

3. The entry light control is located onthe:A. Right forward service pedestalB. The entrance door thresholdC. Left forward service cabinetD. Light assembly

4. The lights that are illuminated by theemergency lighting system are the:A. Instrument panel floodlightsB. Cabin overhead lights, wing inspec-

tion light, and emergency exit lightsC. Navigation lightsD. Strobe lights

5. If installed, the emergency lightingswitch position used during normaloperation is:A. DISARMB. ARMC. TESTD. EMER LT

6. The lights that can be operated with theairplane batteries turned off are the:A. Entry light and baggage lightB. Overhead lightsC. Aisle lightsD. Reading lights




QUESTIONS

4-i

CHAPTER 4MASTER WARNING SYSTEM

CONTENTS

Page

INTRODUCTION ................................................................................................................... 4-1

GENERAL............................................................................................................................... 4-1

GLARESHIELD WARNING LIGHTS................................................................................... 4-2

TEST FUNCTION................................................................................................................... 4-2

INTENSITY CONTROL......................................................................................................... 4-2

BULB CHANGE ..................................................................................................................... 4-2

ILLUMINATION CAUSES .................................................................................................... 4-2

QUESTIONS ........................................................................................................................... 4-5




4-iii

ILLUSTRATION

Figure Title Page

4-1 Test Switch ............................................................................................................... 4-2

TABLE

Table Title Page

4-1 Annunciator Illumination Causes............................................................................. 4-3




INTRODUCTIONThe glareshield warning lights for the Lear 20 series provide a warning for airplane equip-ment malfunctions, an indication of an unsafe operation condition requiring immediateattention, or an indication that a system is in operation. Airplanes prior to 24-156 and25-010 did not originally have glareshield warning lights, but they may have beenretrofitted.

GENERALThe warning system consists of three hori-zontal rows of annunciator lights colored red,amber, and green (see “Annunciator Panel” section), which alert the pilots to various mal-functions or switch positions. The lights arelocated on the center portion of the glareshieldjust above the engine gages. These lights arereferred to as the glareshield warning lights.

Provision is made to test all warning systemlights with a switch located just beneath theglareshield lights panel.

Other annunciator lights are located on the instrument panel, center pedestal, or thrust re-verser control panel, which are not part of themaster warning system. These lights functionindependently as system advisory annunciators.

TEST

CHAPTER 4MASTER WARNING SYSTEM




GLARESHIELD WARNING LIGHTSThe red and amber lights illuminate when amalfunction is detected in the associated system and remain i l luminated unti l the malfunction is cleared. The green lights illu-minate to indicate that the associated systemis in operation.

TEST FUNCTIONDepre s s ing t he t e s t sw i t ch unde r t heglareshield (Figure 4-1) causes the followinglights to illuminate:

• All glareshield warning lights

• Shutoff valve lights

• Fire warning and fire extinguisher armedlights

• Decision height lights

• Antiskid generator lights (Model 25 series airplanes)

• Air ignition lights

• All fuel panel lights

• Pitot heat indicator lights (if installed)

• T/O TRIM and OVSPD lights (if in-stalled)

• LOW OIL , LOW HYD, and FUELXFLO (if installed)

INTENSITY CONTROLThe intensity of the landing gear warninglights and the flight director annunciator lightsis dimmed when the navigation lights switchis turned on. The gear lights are also con-trolled by a rheostat on the landing gear panel.

Photoelectric cells, inboard of each FIREswitch, automatically dim the lights on theglareshield to a level corresponding to exist-ing light conditions in the cockpit or to a min-imum preset level for a totally dark cockpit.

On some airplanes there is also a rheostatunder the left side of the glareshield that con-trols the intensity of the warning lights aboveor below that of the photo cell value.

BULB CHANGEIndividual lights may be “popped out” ap-proximately 1/4 inch by depressing and re-leasing. The light may then be removed bypulling out to replace either of the two bulbsin each light. In the extended position (poppedout), the light is inoperative.

ILLUMINATION CAUSESThe lights associated with the glareshield warn-ing system, their legends, colors, and causesfor illumination are shown in Table 4-1.

NOTESome lights are optional, and arrange-ments may vary between airplanes.




Figure 4-1. Test Switch




Table 4-1. ANNUNCIATOR ILLUMINATION CAUSES

Annunciator AnnunciatorCause for Illumination Cause for Illumination

L. FUEL PRESS INLET HTR

L GEN

R GEN

CAB ALT

ALC AI

PRI INV

SEC INV

PRI INV

SEC INV

STBY INV

AUX INV

LO HYD PRESS

L CAB AIR

R CAB AIR

L STALL

R STALL

WING OV’HT

FUEL XFR

R. FUEL PRESS

SPOILER

DOOR

STAB HEAT

FUEL FILTER

L. ENG ICE

R. ENG ICE

ENG SYNC

BAT 140

BAT 160

LOW FUEL

LO OIL PRESS

WSHLD OV’HT

WSHLD OV’HT

STEER ON

There is less than .25 psi of fuelpressure to the indicated engine.

Either nacelle inlet temperature isabove 190° F.

The indicated generator hasfailed or the switch is positionedto OFF.

Fuel transfer is in progress.

FUEL XFLO Fuel crossflow valve is open.

FUEL XFLO

On SNs 24-350 and subsequentand 25-227 and subsequent,cabin altitude is above 8,750 feet(±250 feet).

Steady—The switch is off or on,and the system has failed, or isin the pusher range.

Flashing— In shaker/nudgerrange.

Right wing structure temperature ismore than 215°F.

The alcohol system has failed orthe alcohol tank is empty.

Either or both spoilers are notdown and locked.

The entrance door is improperlyclosed or malfunctioning.

Either or both fuel filters arebypassing fuel.

Bleed-air pressure to the engine’sfront frame is less than 5 psi.

The engine sync system is turnedON with the nose gear extended.

One or both batteries’ tempera-ture is 140°F or more (nicadbattery-equipped airplanes only).

One or both batteries’ tempera-ture is 160°F or more (nicadbattery-equipped airplanes only).

Less than 450 pounds of fuel re-mains in either wing in levelflight.

There is less than 5 psi oil pres-sure to one or both engines.

Nosewheel steering is engaged.

In fl ight—The wing/stab heatswitch is on, and the system hasfailed.

On ground—The wing/stab heatswitch is on, but stab heat is dis-abled by the right squat switch.

In flight—Windshield duct tem-perature is 250°F.

On ground—Windshield ductheat is 215°F.

The inverter is off or failed.

Early Airplanes

Inverter has failed with the switchon.

The inverter has failed with theswitch on.

Hydraulic pressure is less than1,200 psi (approximately).

On SNs 24-350 and subsequentand 25-227 and subsequent,cabin air duct temperature ismore than 395°F.




Table 4-1. ANNUNCIATOR ILLUMINATION CAUSES (Cont)

Annunciator Cause for Illumination Annunciator Cause for Illumination

PITOT HEAT LOWHYD

FUELXFLO

L LOOIL

R LOOIL

UNSAFE

UNSAFE

LOCK DOWN

(25 SERIES)

L R ANTI-SKIDSYS TEST

WSHLD HEAT

ARMED

L PITOTHEAT

R PITOTHEAT

TAKE-OFFTRIM

PITCHTRIM OVSP

FIRE

The windshield heat defog valveis not closed.

Flashing—The temperature is480°F in the forward section ofnacelle, 510°F in the aft sec-tion.

LOW HYD—Hydraulic systempressure is less than 1,200 psi.

The nose gear is unsafe or intransit, and/or the inboard geardoors are not closed.

The indicated landing gear isdown and locked.

There is an antiskid failure inthe associated brake.

The green lights indicate anti-skid generator operative iftraveling faster than 5 mph. Thewhite light indicates a valid testof the antiskid system.

At or below altitude set in radioaltimeter.

Air ignition operating. Air igni-tion switch is on or system is instart cycle.

FUEL XFLO—Crossflow valve isopen.L LO OIL, R LO OIL—Indicatedengine oil pressure is below 5 psi.

The FIRE switch is in the closedposition. The fuel and hydraulicshutoff valves are closed if thebattery switches are on.

The indicated fire bottle isarmed.

The horizontal stabilizer is notproperly positioned for takeoff.

High-speed trim is available inthe slow-speed mode.

The pitot heat switch is off.

The pitot heat switch is off.

The pitot heat switch is on andsystem has failed.

The pitot heat switch is on andsystem has failed.

FIRE

ARMED

(24 SERIES)

L

DH

AIR IGN

R ANTI-SKIDSYS TEST

DH




QUESTIONS

1. All glareshield lights can be tested by:A. Individual system testingB. Depressing each individual capsuleC. Depressing the test switchD. Shutting the represented system off

2. An illuminated glareshield warninglight suddenly extinguishes, indicating:A. Five minutes have passed.B. The malfunction no longer exists.C. Three minutes have passed.D. The test switch has been reset.

3. The glareshield annunciator light intensity is adjusted:A. Automatically by a photoelectric cellB. By depressing the test switchC. By depressing each individual

capsuleD. By turning on the navigation

lights switch

5-i

CHAPTER 5FUEL SYSTEM

CONTENTS

Page

INTRODUCTION ................................................................................................................... 5-1

GENERAL............................................................................................................................... 5-1

FUEL TANKS AND TANK VENTING SYSTEM ................................................................ 5-3

General ............................................................................................................................. 5-3

Tip Tanks.......................................................................................................................... 5-3

Wing Tanks ...................................................................................................................... 5-3

Fuselage Tank .................................................................................................................. 5-3

Ram-Air Vent System ...................................................................................................... 5-4

FUEL INDICATING SYSTEMS ............................................................................................ 5-4

Fuel Quantity Indicating System/Low Fuel Warning ...................................................... 5-4

Fuel Flow Indicating System ........................................................................................... 5-7

FUEL DISTRIBUTION .......................................................................................................... 5-8

General ............................................................................................................................. 5-8

Boost Pumps .................................................................................................................... 5-8

Motive-Flow Fuel and Jet Pumps .................................................................................... 5-8

Filters ............................................................................................................................... 5-9

Main Fuel Shutoff Valves ................................................................................................ 5-9

Low Fuel Pressure Warning Lights.................................................................................. 5-9

Pressure-Relief Valves ................................................................................................... 5-10

Fuel Drain Valves........................................................................................................... 5-10




FUEL TRANSFER SYSTEM............................................................................................... 5-11

Crossflow Valve ............................................................................................................. 5-11

Transfer Valve/Xfer–Fill Switch (Normal Transfer Line) ............................................. 5-11

Fuselage Valve/Fus Valve Switch (Gravity-Flow Line) ................................................ 5-12

Float and Pressure Switches .......................................................................................... 5-12

Fuselage Fuel Fill-Transfer Operations ......................................................................... 5-13

TIP TANK FUEL JETTISON SYSTEM (OPTIONAL) ...................................................... 5-14

FUEL SERVICING.................................................................................................................5-14

General........................................................................................................................... 5-14

Approved Fuels.............................................................................................................. 5-16

Fuel Density Adjustments.............................................................................................. 5-16

Aviation Gasoline .......................................................................................................... 5-16

Anti-Icing Additive ........................................................................................................ 5-17

Refueling........................................................................................................................ 5-17

OPERATIONAL CONSIDERATIONS ................................................................................ 5-17

QUESTIONS......................................................................................................................... 5-19

Revision .015-ii FOR TRAINING PURPOSES ONLY



Revision .01 5-iii

ILLUSTRATIONS

Figure Title Page

5-1 Fuel System.............................................................................................................. 5-2

5-2 Ram-Air Vent and Vent Sump Drain (Typical) ........................................................ 5-4

5-3 Fuel Vent System...................................................................................................... 5-5

5-4 Fuel Control Panels .................................................................................................. 5-6

5-5 Fuel Flow Indicators................................................................................................. 5-7

5-6 Jet Pump Schematic ................................................................................................. 5-8

5-7 Fuel Drain Locations.............................................................................................. 5-10

5-8 Airplane Grounding Points .................................................................................... 5-15

5-9 Anti-icing Blending Apparatus .............................................................................. 5-15

5-10 Fuel Density Adjustment........................................................................................ 5-16

5-11 Refueling Filler Cap............................................................................................... 5-17




INTRODUCTIONThe Learjet fuel system consists of the fuel tanks, tank venting, indicating, distribution,transfer, and jettison systems.

This chapter covers the operation of the fuel system up to the engine-driven fuel pumps.At that point, fuel system operation becomes a function of the engine. Refer to Chapter7, “Powerplant,” for additional information.

GENERALThe fuel storage system consists of tip tanks,integral tanks in each wing, and a fuselagetank. A crossflow valve permits fuel transferbetween the wings for fuel balancing.

Each wing tank contains a jet pump and anelectric standby pump to supply fuel to the en-gine on the same side. Tip tank and fuselage tankfuel must be transferred into the wing tanks byjet pumps and an electric pump, respectively.

A ram-air system is used to vent all tanks.Drain valves are provided to remove conden-sation and contaminants from the low pointsin the fuel tanks and to drain the contents ofthe vent system sump.

Tip tank fuel can be jettisoned, when required,if the optional jettison system is installed.

Figure 5-1 depicts the Learjet fuel system.

CHAPTER 5FUEL SYSTEM




0

2

4 6

8

10

MAINFUEL

LBS X 100




FILLER CAP

QUANTITY PROBE

BOOST PUMP

JET PUMP

FILTER

ENGINE PUMP

PRESSURE-RELIEF VALVE

CHECK VALVE

FLAPPER VALVE

FLOAT SWITCH

SUPPLY

LOW PRESSURE

HIGH PRESSURE

GRAVITY (TRANSFER)LOW PRESSURE (FILL)

LEGEND

F

P

TOSUMP

FUEL JETTISONSHUTOFF VALVE

LOW FUELPRESSURE SWITCH

DIFFERENTIALPRESSURESWITCHFUEL

SHUTOFFVALVE

75-PSIRELIEFVALVE

*24 SERIES AIRPLANES**25 SERIES AIRPLANES

** *

CROSSFLOWVALVE

WINGPRESS SW

TRANSFERVALVE

FUSELAGETANK(NOT INSTALLEDON LR-24E)

TRANSFERLINE

EMPTY LIGHTPRESSURESWITCH

MOTIVE-FLOWVALVEMOTIVE-

FLOWFUEL

MODELS 24D, E, F,AND 25B, D

F

P

P

P

PP

TOSUMP

FUEL JETTISONSHUTOFF VALVE

LOW FUELPRESSURESWITCH

DIFFERENTIALPRESSURESWITCH

FUELSHUTOFF

VALVE

75-PSIRELIEFVALVE

CROSSFLOWVALVE

WINGPRESS

SW

TRANSFERVALVE TRANSFER

LINE

FUSELAGEVALVE

GRAVITY-FLOWLINE

EMPTY LIGHTPRESSURESWITCHFUSELAGE

TANK(MODEL 25CTANK SHOWN)

MOTIVE-FLOWVALVE

MOTIVE-FLOWFUEL

MODELS 25C AND 25DWITH OPTIONAL

GRAVITY-FLOW LINEF

P

P PP

P

Figure 5-1. Fuel System

FUEL TANKS AND TANKVENTING SYSTEMGENERALApproximately 23 gallons of fuel in the 20 se-ries fuel system is trapped. The weight of thisfuel is included in the airplane basic weight.The approximate total usable fuel is 4,800pounds in the 24E model, 5,600 pounds in the24D and F models, 6,050 pounds in the 25B andD models, and 7,350 pounds in the 25C model.More detailed information is provided in the“Weight and Balance” section of the AFM.

TIP TANKSEach tip tank capacity is 1,235 pounds of us-able fuel; capacity is reduced to 1,195 poundswith installation of a recognition light. Thetanks are permanently attached to the wingsand are positioned at 2° nose down relative tothe airplane centerline. Baffles are installedto minimize slosh and prevent adverse effectson the airplane center of gravity during ex-treme pitch attitudes.

A jet pump installed in each tip tank transfersfuel into the wing tank. Approximately one halfof the fuel will gravity-flow through two flap-per valves into the wing tank; however, any fuelat a level lower than one-half full must betransferred using the jet pump. A standpipe isinstalled in each jet pump transfer line to pre-vent fuel from being siphoned from the wingtank to the tip tank when the applicable engineis shut down.

The tip tank is vented through two vent floatvalves located in the forward and aft ends ofthe tank.

A fuel probe in each tip tank provides infor-mation to the fuel quantity indicating system.

All tip tank fuel can be jettisoned through avalve in the tank tailcone, if the optional jet-tison system is installed.

A filler cap on each tip tank is used to servicethe entire airplane fuel system.

WING TANKSEach wing tank extends from the airplane cen-terline to the tip tank and holds 1,160 poundsof usable fuel. Areas which are not part of thewing fuel cell are the main landing gear wheelwell, the leading edge forward of spar 1 (wingheat area), and the trailing edge between spars7 and 8 (flap, spoiler, and aileron areas).

The 2.5° wing dihedral makes the inboard por-tions of the wing tanks the lowest areas. In eachwing tank, a jet pump and an electric standbypump are located within these areas and willremain submerged in fuel until the tanks arenearly empty.

Wing ribs and spars act as baffles to minimizefuel shifting. Flapper valves located in thewing ribs allow unrestricted inboard flow offuel and limit outboard flow. Two pressure-re-lief valves at the centerline rib equalize inter-nal pressures between the two wing tanks. Thewing tanks begin to fill through the two tip tankflapper valves as tip tank fuel increases beyondone-half full.

Two fuel probes in each wing tank provideinformation to the fuel quantity indicatingsystem.

FUSELAGE TANKThe fuselage tank consists of rubber bladderfuel cells located between the aft pressurebulkhead and tailcone section. The 24D and Fand 25B and D models are equipped with twofuel cells with a capacity of 840 and 1,305pounds, respectively, of usable fuel. The model25C is equipped with four fuel cells with a ca-pac i ty o f 2 ,603 pounds o f u sab le fue l .Depending on the airplane, either one or twofuel lines connect the fuselage tank to the wingtanks for filling and transfer. This is explainedin sections entitled “Transfer Valve/XFER-FILL Switch.” Model 24E has no fuselage tank.One fuel probe provides information to thefuel quantity indicating system.




RAM-AIR VENT SYSTEMA NASA ram-air scoop located on the under-side of each wing (Figure 5-2) supplies posi-tive air pressure in flight to a manifold whichdirectly vents the fuselage tank and both tiptanks. Each wing tank is indirectly vented toits own tip tank through a length of tubing, theends of which extend to the uppermost area ofeach tank (Figure 5-3). The NASA scoops, bydesign, do not require heating to remain ice-free. Two vent float valves are located in eachtip tank and are in the fuselage tank except inmodel 25C. The float valves close when thefuel level reaches the vent ports, preventingfuel from entering the vent lines. A vacuum-relief valve in each tip tank and the fuselagetank opens to allow air to enter the tanks shouldvacuum conditions occur during fuel trans-fer. Each tip tank has two pressure-relief valveswhich protect the tanks from excessive pres-sure due to thermal expansion of fuel in the

tank when the vent float valves are closed.Thepressure-relief valves are set at 1.0 and 1.5 psi,the second valve providing a backup in casethe first valve fails.

Thermal expansion of fuselage fuel is accountedfor by a continuation of the fuselage vent linebypassing the fuselage vent float valve, therebyrelieving pressure overboard through the NASAscoops. A sump, installed in the vent manifold,located at the bottom center fuselage just aft ofthe main landing gear, collects any fuel whichmight enter the vent lines during refueling. Avent drain valve permits draining of the sump,which must be accomplished prior to everyflight to ensure that the vent line to the fuse-lage tank is unobstructed.

FUEL INDICATING SYSTEMSFUEL QUANTITY INDICATINGSYSTEM/LOW FUEL WARNINGThe fuel quantity indicating system includes anindicator and tank selector switch located onthe fuel control panel (Figure 5-4). A red LOWFUEL warning light (“Annunciator Panel” sec-tion) illuminates if either wing tank fuel levelis less than approximately 450 pounds.

The fuel quantity indicating system uses DCpower from the right and left essential busesthrough the FUEL QTY circuit breakers. Thesix-position rotary selector switch enables thepilot to check the fuel quantity in each of thefive tanks and the airplane total fuel quantity.The fuel quantity for the position selected isread on the fuel quantity indicator. The quan-tities printed beside each selector switch po-sition indicate usable fuel capacities in pounds.

There are seven capacitance fuel probes. Onefuel probe is located in each tip tank and in thefuselage tank. Each wing tank has two probeswired in parallel. The inboard probe in the leftwing contains a temperature compensator whichadjusts quantity readings for all switch selec-tions for fuel density change due to temperature.




Figure 5-2. Ram-Air Vent and Vent SumpDrain (Typical)




WINGVENT

FLOATVALVE

VACUUMRELIEF

FUELVENT

DRAIN

OVERBOARDDRAIN

FLAMEARRESTER

RAM-AIRSCOOP

FLOATVALVE

(TYPICAL)

PRESSURE RELIEF

VACUUMRELIEF

TO AMBIENT

VACUUM-RELIEFVALVE

1.5-PSIRELIEFVALVE

1.0-PSIRELIEFVALVE

Figure 5-3. Fuel Vent System




CLOSE

JET PUMPS

STANDBY PUMPS

0 0 0 0

0

1

2

3 4

5

6

FUELQUANTITY

LBS 1000

L TIP1235

L TIP1235

L WING1160

L WING1160

R WING1160

R WING1160

R TIP1235

R TIP1195

TOTAL5630LBS

TOTAL5590LBS

FUS840

L RON

L RON

OPEN

CROSS FLOWFULL

OFF

EMPTYXFER

FILL

FUS TANK

FUS840

L TIP1235

L WING1160

R WING1160

R TIP1195

TOTAL6055LBS

FUS1305

L TIP1220

L WING1160

R WING1160

R TIP1220

TOTAL6029LBS

FUS1305

L TIP1220

L WING1160

R WING1160

R TIP1180

TOTAL6025LBS

FUS1305

24D/F MODELS

CLOSE

JET PUMPS

STANDBY PUMPS

0 0 0 0

0

1

2

3 4

5

6

FUELQUANTITY

LBS 1000

L TIP1235

L WING1160

R WING1160

R TIP1235

TOTAL6095LBS

FUS1305

L RON

L

L R

RON

ON

FUEL JETTISON

OPEN

CROSS FLOWFULL

OFF

EMPTYXFER

FILL

FUS TANK

25B/D MODELS

CLOSE

JET PUMPS

STANDBY PUMPS

0 0 0 0

0

1

2

3 4

5

6

FUELQUANTITY

LBS 1000

L TIP1235

L WING1160

R WING1160

R TIP1235

TOTAL4790LBS

L RON

L RON

OPEN

CROSS FLOW

MODEL 24E

MODELS EQUIPPED WITH RECOGNITION LIGHT

25B/D MODELS WITHRECOGNITION LIGHT

25B/D MODELS WITHFUEL JETTISON

25B/D MODELS WITHRECOGNITION LIGHT AND

FUEL JETTISON

L TIP1235

L WING1160

R WING1160

R TIP1195

TOTAL7353LBS

FUS2603

L TIP1220

L WING1160

R WING1160

R TIP1220

TOTAL7363LBS

FUS2603

L TIP1220

L WING1160

R WING1160

R TIP1180

TOTAL7323LBS

FUS2603

CLOSEFUS VALVE

JET PUMPS

STANDBY PUMPS

0 0 0 0

FUELQUANTITY

LBS 1000

L TIP1235

L WING1160

R WING1160

R TIP1235

TOTAL7393LBS

FUS2603

L RON

L

L R

RON

ON

FUEL JETTISON

OPEN

CLOSE

CROSS FLOW

OPEN

FULL

OFF

EMPTYXFER

FILL

FUS TANK

25C/F MODELS25C/F MODELS WITHRECOGNITION LIGHT

25C/F MODELS WITHFUEL JETTISON

25C/F MODELS WITHRECOGNITION LIGHT AND

FUEL JETTISON

0

1

2

34

5

6

7

8

Figure 5-4. Fuel Control Panels

NOTEIf the lef t inboard compensatorprobe becomes dry, it will causeerroneous fuel quantity indicationsat all switch positions.

Each probe uses an electrical capacitance meas-uring system to sense the fuel level. It then trans-mits an electrical signal to the cockpit indicatorwhere it is read as pounds x 1,000 on the dial.

NOTEThe total position of the selectorswitch should be used for flight plan-ning purposes. The quantity indicat-ing system is set to indicate properlyin level flight.

Each wing tank has a fuel low-level floatswitch. When either wing tank fuel levelreaches 400 to 500 pounds remaining, the re-spective float switch actuates the red LOWFUEL light on the annunciator panel to indi-ca t ed l ow wing fue l quan t i t y. (See“Annunciator Panel” section.) When flyingwith the LOW FUEL light on, limit pitch at-titude and thrust to the minimum required.

When the fuel quantity gage indicates600 pounds or less remaining in eitherwing tank, prolonged noseup attitudeof 10º or more may cause fuel to betrapped in the aft area of the wingtank outboard of the wheel well. Fuelstarvation and engine flameout mayoccur. Reducing pitch attitude andthrust to the minimum required willprevent this situation.

FUEL FLOW INDICATINGSYSTEMA fuel flow indicator provides a readout ofpounds of fuel flow per hour. The indicator(Figure 5-5) may have one or two pointers (L andR) and uses DC power. Gages with one pointerhave a L–R selector switch. The fuel counter,located on the fuel control panel, provides a

WARNING




SINGLE POINTER

DUAL POINTER

Figure 5-5. Fuel Flow Indicators

four-digit readout (in pounds of fuel consumedby both engines) and uses DC power. It shouldbe reset to zero using the reset button adjacentto the counter before starting the first engine.

FUEL DISTRIBUTION

GENERALEach engine is supplied with fuel from its re-spective wing fuel system; there is no crossfeedcapability. Either the wing standby pumps or thewing jet pumps supply fuel under pressure to theengine-driven pumps. During engine start, therespective wing standby pump is automaticallyenergized and the motive-flow valve closedwhen the GEN–START switch is placed in theSTART position. After the engine is started andthe START switch is moved to the OFF position,the wing standby pump is deenergized, the mo-tive-flow valve opens, and the wing jet pumpthen provides fuel to the engine. The wing jetpumps and standby pumps have check valves onthe output side to prevent reverse flow when theyare inactive.

BOOST PUMPSSubmerged DC-powered boost pumps are in-stalled at three different locations—one standby

pump in each wing adjacent to the jet pump andone transfer pump in the fuselage tank.

NOTEBoth wing tank standby pumps aredeactivated when the XFER–FILLswitch is in the XFER position.

The wing tank standby pumps are used:

• For engine start (automatically ener-gized with starter switch activation)

• As a backup for the wing jet pumps

• For wing-to-wing crossflow

• For filling the fuselage tank (automati-cally energized with the XFER–FILLswitch in the FILL position)

The fuselage tank transfer pump is used totransfer fuselage tank fuel to the wing tanks.

The wing standby pumps are powered from theleft and right essential buses; the fuselagepump is powered from the right main bus.

MOTIVE-FLOW FUEL AND JETPUMPSHigh-pressure fuel from the engine-drivenfuel pumps is the source of motive-flow fuel




Figure 5-6. Jet Pump Schematic

INPUTOUTPUT

WING TANKSTRUCTURE

FUEL SUPPLY HIGH PRESSURE LOW PRESSURE

LEGEND

which causes the jet pumps to function. Thefuel is routed through the motive-flow controlvalves (one on each side) to the jet pumps,where it sprays through a small orifice into aventuri. The low pressure created in the ven-turi draws fuel from the respective tank, result-ing in a low-pressure, high-volume outputfrom the jet pump (Figure 5-6).

Motive-flow pressure varies with engine rpm.Consequently, jet pump discharge pressurealso varies with engine rpm. At idle, dischargepressure is approximately 10 psi, while at full-power settings, discharge pressure is approx-imately 12 psi.

There are four jet pumps—one in each wing tankadjacent to the standby pump, and one in eachtip tank. The wing tank jet pumps draw fuel fromthe wing tanks and supply low-pressure fuel tothe engine-driven, high-pressure fuel pumps.Wing jet pump output can be supplemented bythe wing standby pump to ensure positive pres-sure to an engine. The tip tank jet pumps drawfuel from the tip tanks and deliver it directly tothe cavities where the standby pumps and jetpumps are located.

Jet pumps require no electrical power andhave no moving parts. They are controlled bytwo jet pump switches which electrically openand close the motive-flow control valves. Theamber indicator lights next to the switches il-luminate when the motive-flow valves are intransit or are not in the position selected on theswitch. Each jet pump switch (and motive-flow control valve) controls both jet pumps(wing and tip) on that side.

NOTEShould operational requirementsmake it necessary to turn off a jetpump switch, the standby pump onthe same side must first be turned onto ensure against loss of pressure tothe engine-driven pump.

NOTEWith an engine shut down, both thewing and the tip tank jet pumps areinoperative on that side.

NOTEIf a jet pump indicator light remainsilluminated after a selection is made,the motive-flow valve is not in theposition selected by the switch.

The motive-flow control valves close auto-matically when the GEN–START switch isenergized, and reopen when the GEN–STARTswitch is moved to OFF.

FILTERSA fuel filter is installed in each engine feed lineto filter the fuel before it enters the engine-driven fuel pump. Should the filters becomeclogged from ice or contaminants, the fuel by-passes the primary element and is filteredthrough a secondary element. A differentialpressure switch installed in each filter assem-bly illuminates the amber FUEL FILTER lightwhen one or both filters are bypassing fuel.

MAIN FUEL SHUTOFF VALVESThe fuel shutoff valves are powered by essen-tial bus DC power through the L and R FWSOV circuit breakers and are controlled bythe FIRE switches on the glareshield. Pushingeither FIRE switch closes the associatedvalves, illuminates the SOV light, and arms thefire extinguisher switches. Pushing the FIREswitch a second time will open the valve.

NOTEWhen the FIRE switch is pushed,the fire extinguisher is armed. Don o t a c c i d e n t a l l y d e p r e s s t h eARMED buttons.

LOW FUEL PRESSUREWARNING LIGHTSA low fuel pressure switch is located in eachengine feed line. The switches cause illumi-nation of the appropriate red L or R FUELPRESS annunciator light when fuel pressuredrops below 0.25 psi. The light extinguisheswhen pressure increases above 1 .0 ps i .Illumination of a FUEL PRESS warning lightis an indication of loss of fuel pressure to the




engine. The probable cause is failure of the af-fected wing jet pump.

NOTEThe engine-driven pump is capableof suction-feeding enough fuel to sus-tain engine operation without eitherthe wing standby pump or jet pump.However, 25,000 feet is the highest al-titude at which continuous operationshould be attempted in this event; thelimiting factor is that possible dam-age to the engine-driven pump mayoccur due to reduced fuel flow.

PRESSURE-RELIEF VALVESA 75-psi relief valve is installed in each mainfuel line on the engine side of the main shut-off valve. The valves relieve pressure buildupcaused by thermal expansion of trapped fuel

when the engines are shut down by venting fueloverboard.

Two one-way pressure-relief valves are lo-cated at wing rib 0.0, which separates the leftand right wing fuel tanks. Each valve, reliev-ing in the opposite direction, opens at 1 psi toequalize fuel pressure between the wing tanksduring crossflow operation.

FUEL DRAIN VALVESDrain valves are located at low points through-out the fuel system for draining condensationor sediment. A small amount of fuel should bedrained from each valve during the exteriorpreflight inspection. The valves, spring-loadedto the closed position, are located as follows:one for each tip tank, one for the crossflowvalve, one for each wing, one for each line tothe engines, one for each fuel filter, and one, two,or three for the fuselage tank (Figure 5-7).




FLUSH SUMP DRAIN

WING SUMP DRAIN

RIGHT WING LEFT WING

LH ENGINE FUELLINE DRAIN

EPA CAN DRAIN

RH ENGINE FUELLINE DRAIN

EPA CAN DRAIN

FUEL VENT DRAIN

FUEL FILTER DRAIN FUEL FILTER DRAIN

FUSELAGELINE

DRAIN

FLUSH SUMP DRAIN

CROSSFLOW DRAIN

WING SUMP DRAIN

FUSELAGE TANK DRAIN(S)24 D, E, F25 B, D25 C

1 DRAIN2 DRAINS3 DRAINS

Figure 5-7. Fuel Drain Locations

NOTEModels 24D, E, and F have one fuse-lage drain valve; models 25B and Dhave two; and model 25C has three.

There is one drain valve located at the fuel ventsump. The valve must be completely drainedduring the exterior preflight inspection to pre-vent possible blockage of the fuselage ram-airvent line.

FUEL TRANSFER SYSTEMCROSSFLOW VALVEA DC motor-driven valve is installed in thecrossflow manifold connecting the two wingtanks. It is opened during fuselage fuel trans-fer and filling operations and for wing-to-wing fuel balancing. The valve is controlledby t he CROSS FLOW swi t ch o r t heXFER–FILL switch on the fuel control paneland is powered through the right main busFILL and XFER circuit breaker. Additionally,on those airplanes equipped with the gravity-flow transfer line, the valve is controlled withthe FUS VALVE switch, which is poweredfrom the FUS VALVE circuit breaker on theleft essential bus.

The amber light adjacent to the CROSS FLOWswitch illuminates when the valve is in tran-sit or is not in the position selected. An optionalgreen or amber FUEL XFLO annunciator light(see “Annunciator Panel” section) on the glare-shield illuminates continuously whenever thecrossflow valve is fully open.

If wing fuel imbalance occurs, crossflow is ac-complished by opening the crossflow valve andturning on the standby pump in the heavywing, while ensuring that the standby pump inthe light wing is off. The transfer rate is ap-proximately 50 pounds of fuel per minute.

Opening the crossflow valve to balance fuelshould not be attempted when a red FUELPRESS light is illuminated unless it can be ac-complished below 25,000 feet. To do so would

divert pressure being delivered to the affectedengine-driven pump by the operating standbypump on that side. Asymmetric power set-tings may be used to balance fuel, if necessary.

The above considerations do not apply to sin-gle-engine operations, and normal crossflowoperations may be performed as usual.

TRANSFER VALVE/XFER–FILL SWITCH (NORMALTRANSFER LINE)The fuel transfer line connects the fuselagetank with the crossflow manifold. A DC motor-driven transfer valve installed in the line con-trols fuel movement between the fuselage andwing tanks. The valve is controlled by theXFER–FILL switch located on the fuel con-trol panel. When the switch is positioned fromOFF to XFER, the transfer and crossflowvalves are sequenced open and the fuselagetransfer pump is energized automatically,while both wing standby pumps, if operating,are deactivated. When the switch is positionedfrom OFF to FILL, the transfer and crossflowvalves are sequenced open, and both wingstandby pumps are energized automatically.When the switch is positioned from eitherXFER or FILL to the OFF position, the trans-fer pump or wing standby pumps (whicheverthe case may be) are deenergized and the trans-fer valve and crossflow valves are sequencedclosed. The amber l ight adjacent to theXFER–FILL switch illuminates when thevalve is in transit or is not in the position se-lected. The valve is powered through the rightmain bus FILL and XFER circuit breaker.

NOTEThe transfer line is connected to theright side of the crossflow valve. Onall 25C models, and 25D models withthe optional gravity-flow line, thetransfer line is connected to the leftside of the crossflow valve. (SeeFuse lage Va lve /FUS VALVESwitch.) In either case, should thefuselage transfer pump fail, fuel canbe expected to gravity-flow throughthe transfer line.




FUSELAGE VALVE/FUS VALVESWITCH (GRAVITY-FLOW LINE)A DC motor-driven fuselage valve is installedin a second fuel line, connecting the fuselagetank with the crossflow manifold on the rightside of the crossflow valve (see Figure 5-1).The valve is controlled by the FUS VALVEswitch on the fuel control panel. When the FUSVALVE switch is positioned to OPEN, both thefuselage valve and the crossflow valve are si-multaneously sequenced open, allowing fuel togravity-flow from the fuselage tank to bothwings. The fuselage valve is also controlled bythe XFER–FILL switch. When placed to FILL,the transfer valve, fuselage valve, and crossflowvalve are sequenced to open and the wingstandby pumps are energized to pump wingtank fuel through both fuel lines into the fuse-lage tank. The fuselage valve remains closedwhen the XFER–FILL switch is positioned toXFER. The amber light adjacent to the FUSVALVE switch will illuminate when the fuse-lage valve is in transit or is not in the positionselected. The fuselage valve is powered throughthe left essential bus FUS VAL circuit breaker.

NOTEIf either wing standby pump switch ison, the FUS VALVE switch is ren-dered inoperative, and neither thefuselage valve nor the crossflow valvewill open if the FUS VALVE switchis moved to OPEN. Conversely, if theFUS VALVE switch is already in theOPEN position (fuselage valve andcrossflow valve open), turning eitherwing standby pump switch on willautomatically cause the fuselage valveand crossflow valves to sequenceclosed.

NOTEPositioning the XFER–FILL switchto XFER while also positioning theFUS VALVE switch to OPEN should

be avoided because it will cause bothfuselage transfer lines and the cross-flow line to be open and the fuselagepump to be energized. The result ofthis action is that some fuel will bereturned to the fuselage tank throughthe open gravity-flow line. The whitetank EMPTY light may also illumi-nate due to a drop in line pressure.

FLOAT AND PRESSURE SWITCHES

Fuselage Fuel Tank FloatSwitchWhen filling the fuselage tank, a float switchmounted inside the tank actuates when thetank is full. The switch:

• Turns on the green FULL light on thefuel control panel

• Deenergizes the wing standby pumps

• Closes the transfer and crossflow valves

• Closes the fuselage valve (all airplanesequipped with the gravity-flow trans-fer line)

The green FULL light on the fuel control panelremains illuminated until the XFER–FILLswitch is turned off.

Fuselage Tank Low-Pressure SwitchThe fuselage tank low-pressure switch is in-stalled in the fuselage transfer line to alert thepilot when fuselage fuel is depleted. With theXFER–FILL switch in the XFER position,the switch senses low pressure in the line andilluminates the white EMPTY light on thefuel control panel when either of two condi-tions exist:

• The tank is empty.

• The fuselage transfer pump fails.




Wing Fuel Pressure SwitchA wing fuel pressure switch is installed toprevent internal overpressurization of thewings during transfer of fuselage tank fuel. Theswitch, located in the right main wheel well,deenergizes the fuselage transfer pump whenwing fuel pressure activates the switch.

FUSELAGE FUEL FILL-TRANSFER OPERATIONS

Fill OperationFuel may be pumped from the wings to thefuselage tank using the FILL position onthe XFER–FILL switch. The FILL positionmay be used for CG considerations in flight;however, it is normally used only during fuelservicing.

When the XFER–FILL switch is placed to theFILL position:

• The transfer valve opens.

• The crossflow valve opens.

• The wing standby pumps are energized.

• The fuselage tank float switch is enabled.

NOTEOn 25C airplanes, automatic open-ing of the fuselage valve also occursas a result of selecting the FILL po-sition on the XFER–FILL switch.The filling time is shortened as bothlines are open.

Transfer OperationsTransferring fuselage fuel in flight is accom-plished by using the XFER position on theXFER–FILL switch. When the switch is placedin the XFER position:

• The wing standby pumps are disabled.

• The transfer valve opens.


• The fuselage transfer pump is energized.

• The whi te EMPTY l igh t (p ressureswitch) is enabled.

NOTEAs fuselage quantity approacheszero, the crew should be alert forsteady illumination of the whiteEMPTY light. Failure to positionthe XFER–FILL switch to OFF atthat time can cause damage to thetransfer pump.

When the XFER–FILL switch is placed in theOFF position:

• The transfer pump is deenergized.

• The transfer valve closes.

• The crossflow valve closes.

• Operation of the wing standby pumps isonce again possible.

An optional method of transferring fuselagefuel in flight is only possible on airplanesequipped with the gravity-flow line by usingthe OPEN position on the FUS VALVE switch.However, prior to doing so, it is essential tofirst assure that the XFER–FILL switch is off,and that both wing standby pump switchesare off. Then, when the FUS VALVE switch isplaced in the OPEN position:

• The fuselage valve opens.


When the amount of fuel in the wing tanks be-gins to decrease, the FUS VALVE switch maybe turned off, and the transfer process may becompleted using the normal transfer procedure.

NOTEThe approved AFM lists several re-s t r ic t ions concern ing when thetransfer of fuel from the fuselagetank may be initiated. See “FuelManagement,” Section II, of the ap-proved AFM.




TIP TANK FUEL JETTISON SYSTEM(OPTIONAL)A DC motor-driven valve in the tailcone ofeach tip tank provides the capability of jetti-soning tip tank fuel. One FUEL JTSN switchon the fuel control panel controls both tip tankjettison valves. When the FUEL JTSN switchis placed to the ON position, the jettison valvesopen and two amber lights illuminate contin-uously on the fuel control panel, indicating thatthe valves are fully open. The jettison tubesare scarfed, which creates a low-pressure areathat helps pull the fuel out of the tank(s). This,in combination with the force of gravity, en-ables the entire contents of both tanks to be jet-tisoned. Fuel jettison can be accomplishedwith flaps and landing gear in any position andat all speeds. Fuel jettison is faster while theairplane is in a noseup attitude. It takes approx-imately five minutes to jettison fuel from thetip tanks.

NOTEJettisoned fuel is not recorded onthe fuel counter. Therefore, the fuelcounter reading should not be usedto compute fuel remaining or air-plane weight after jettisoning fuel.

FUEL SERVICING

GENERALFuel servicing includes those procedures neces-sary for fueling and adding anti-icing additives.

Fueling is accomplished through a filler capin the top of each tip tank. Fuel then begins toflow by gravity from the tip tanks into thewing tanks as the tip tanks reach one-half full.The wing standby pumps fill the fuselage tankwhen the XFER–FILL switch is set to FILL.

At normal temperatures, some water is alwaysin solution (dissolved) with fuel. At high al-

titudes, fuel undergoes a cold-soaking processand small amounts of water come out of thesolution and subsequently freeze. The anti-icing additive specified for use in the Learjet20 series must conform to MIL-I-27686. Italso prevents the growth of microbiological or-ganisms in the fuel.

Safety PrecautionsRefueling should be accomplished only inareas which permit free movement of fireequipment.

Ground fuel truck to apron and nosegear uplatch spacer and ground fuelnozzle to tip tank ground jack priorto removing filler cap. This will pre-clude possible fire and/or explosiondue to static electricity or sparks.

Figure 5-8 shows the airplane ground points.

When adding anti-icing inhibitor (Figure 5-9),follow the instructions for blending found inthe AFM.

Hi-Flo Prist may be harmful if in-haled or swallowed. Use adequateventilation. Avoid contact with skinand eyes. If sprayed into eyes, flushwith large amounts of water and con-tact physician immediately.

WARNING

WARNING







MIL-I-

27686

FUEL ADDITIVE

TRIGGER

RING

HANDLE

BLENDER HOSE

FUEL NOZZLE

Figure 5-9. Anti-icing Blending Apparatus

Figure 5-8. Airplane Grounding Points

APPROVED FUELSThe “Limitations” section of the approvedAFM lists specific manufacturer’s fuels ap-proved for use.

FUEL DENSITY ADJUSTMENTSIf optimum engine acceleration is desired, thedensity adjustment knob on the front of the fuelcontrol unit should be readjusted to the fuel den-sity setting for the fuel type being used. Referto the AFM for recommended fuel control adjust-ment knob positions (density setting).

Engine stall margin may be reducedduring high-altitude operation if thefuel control density adjustment knobis positioned to a lower density set-ting than that recommended for typeof fuel being used.

Either of two types of fuel density adjustmentknobs may be installed. They are designatedType A and Type B (Figure 5-10). To make anadjustment, push the knob in, rotate the knobto the density setting desired, and release pres-sure. If necessary, jiggle the knob until it locks.The adjustment knob must be seated in its po-sitioning detent for satisfactory operation.

AVIATION GASOLINEAviation gasoline may be used as an emer-gency fuel and mixed, in any proportion, withthe various approved kerosene-base fuels.

Aviation gasoline may not be used in excessof 25 hours between engine overhauls.

When using any aviation gasolinein the fuel mixture, operation is lim-ited as follows:

• Do not take off with fuel temperaturelower than –54ºC (–65ºF).

• Restrict airplane flights to below 15,000feet.

• Both jet pumps and both wing standbypump switches must be on and the pumpsmust be operating.

NOTECertain aviation gas and jet fuel mix-ture ratios can produce greater flam-mability hazards under partiallyfilled aircraft tank or storage tankconditions encountered during re-fueling or mixing. Follow estab-lished safety standards to eliminateany possible sparks.

CAUTION

CAUTION




Figure 5-10. Fuel Density Adjustment

75 7 2

69

86813

78

FUEL ADJUST—P

US

H&

TU

RN

JP-5&

6

JP-4

PU

SH

TUR

N

NOMINAL

DEC

INC

–1 +1

+2+

3+4

–6–5

–4–3

–2

INDEXPIN

INDEXPIN

TYPE BTYPE A

TYPE APOSITION

TYPE BPOSITION

DENSITYSETTING

8 6 7 8 7 5 7 2 6 981 3

JP5 & JP6 JP4

86

87

–5 –4 –3 –2 –1 +1

84 81 78 75 72 69

813 78 75 72 69

INCREASEDECREASENOMINAL

ANTI-ICING ADDITIVEAny approved fuels not containing the addi-tive must have it blended during refueling.Prior to refueling, check with the fuel supplierto determine if the fuel contains anti-icingadditive meeting the requirements of MIL-I-27686. The additive must be used in amountsnot to exceed 0.15% by volume and a minimumconcentration of 0.06%. Not less than 20 fluidounces (1 can) of additive per 260 gallons(984.2 liters) of fuel nor more than 20 fluidounces per 104 gallons (393.7 liters) of fuelshould be used. Refer to Section II of the ap-proved AFM for the proper anti-icing additiveblending procedure.

Lack of anti-icing additive may causefuel filter icing and subsequent en-gine flameout.

REFUELINGRefueling is accomplished through the tiptank filler caps (Figure 5-11). The fuel beginsto flow by gravity into the wing tanks as thetip tanks reach one-half full. The wing standbypumps are used to fill the fuselage tank. (SeeFuel Transfer System, this chapter.) A groundpower unit should be used, if possible, be-cause of the requirement to operate the standbypumps. Refer to the approved AFM for de-tailed refueling procedures.

Do not completely fill one tank be-fore adding fuel to the opposite tank.Fill both tanks simultaneously, oralternately add 125 gallons (473.2liters) of fuel to each tank until de-sired amount is obtained. Failure tofollow this procedure will result inexcessive lateral unbalance.

On models 25C and D with the FUSVALVE switch, do not set the FUSVALVE switch to OPEN. Check toassure that it is set to CLOSE.

OPERATIONAL CONSIDERATIONSMaximum demonstrated fuel imbalance forlanding is 600 pounds in one tip tank.

Maximum tip tank fuel for landing is 800pounds in each tip tank.

A minimum of 600 pounds of fuel in eachwing tank is required for takeoff and inten-tional go-around.

CAUTION

WARNING

CAUTION




Figure 5-11. Refueling Filler Cap

If either wing tank is below 600 pounds, limitprolonged pitch attitude to 10° nose up.

With the LOW FUEL warning light illumi-nated, limit pitch attitude to the minimum re-quired for go-around.

Maximum fuel in tip tank to begin normaltransfer is 800 pounds in 24 models, 760pounds in 25 models.

With wing jet pumps and standby pumps in-operative, the engine-driven fuel pumps suc-tion-feed sufficient fuel to supply the enginesat altitudes of 25,000 feet or below.

Do not energize fuselage transfer system whenwing and tip tanks are full.

Do not crossflow with wing jet pump inoperative.

If the crossflow valve fails to open, fuselagefuel gravity-flow into the right wing tank (lefttank on 25C airplanes).

Minimum fuel temperature for takeoff is:

• Jet B/JP-4 ....................... –54ºC (–65ºF)

• Jet A/Jet A-1 .................. –29ºC (–20ºF)

• AVGAS ........................... –54ºC (–65ºF)

In level flight with gear and flaps down, fueljettison of full tip tanks takes approximatelyfive minutes.




1. Approximately 23 gallons of fuel in the20 series is trapped. This fuel weight:A. Must be added to the weight of fuel

taken on board when servicing theairplane

B. Is included in the airplane basicweight for airplanes certified in theUnited States

C. Must be accounted for in the fuse-lage tank for CG purposes

D. May be disregarded since it is lessthan 200 pounds

2. With the exception of the FUEL JTSNlights, all other amber lights on the fuelcontrol panel, when illuminated steady,indicate that the respective:A. Valves are cycling or the pumps are

properly operating.B. Valves are in the correct position;

the pumps are inoperative.C. Switch position agrees with the

valve position or pump operation.D. Valve position disagrees with the

switch position.

3. The red LOW FUEL light illuminateswhen:A. 350 pounds total fuel remains.B. 250 to 350 pounds remain in either

wing, depending on the airplane SN.C. 400 to 500 pounds total fuel remains.D. 400 to 500 pounds remain in either

wing.

4. The wing standby pumps are used forall the following functions except:A. Engine startB. As a backup for the main jet pumpsC. Fuselage-to-wing fuel transferD. Wing-to-fuselage transfer of fuel

5. The crossflow valve opens:A. Only when the CROSS FLOW

switch is set to OPENB. Only when the CROSS FLOW

switch is set to OPEN or theXFER–FILL switch is set to XFER

C. Anytime electrical power is lostD. Whenever the CROSS FLOW,

XFER–FILL, or FUS VALVEswitches are moved from the OFF orCLOSED position

6. Steady illumination of an amber trans-fer valve light indicates:A. The valve has failed closed.B. The valve has failed open.C. The valve has stuck in an intermedi-

ate position.D. The valve has failed to move to the

position commanded by theXFER–FILL switch.

7. Illumination of the red L or R FUELPRESS light indicates:A. Fuel pressure to the respective en-

gine-driven fuel pump is low.B. Fuel pressure to the respective en-

gine is too high for safe operation.C. A fuel filter is bypassing.D. Fuel pressure to the respective en-

gine is optimum for engine start.

8. When the XFER–FILL switch is placedto the FILL position, the:A. Float switch is disabled.B. Wing standby pumps are disabled.C. Fuselage valve closes.D. Crossflow valve opens.




QUESTIONS

9. Motive-flow fuel for the jet pumps issupplied by the:A. Engine-driven fuel pumpsB. Wing standby pumpsC. Fuselage transfer pumpD. Motive-flow control unit

10. The amber FUEL FILTER light indicates:A. Low fuel pressure to the engine-

driven pump; the standby pumpsshould be turned on

B. That both fuel filters are bypassingfuel; the light does not illuminate ifonly one filter is bypassing

C. That one or both fuel filters are by-passing fuel

D. That only the secondary fuel filtersare being bypassed

11. When actuated, the wing fuel pressureswitch turns:A. Off the fuselage transfer pumpB. On the fuselage transfer pumpC. Off the wing standby pumpsD. On the wing standby pumps

12. When using any aviation gasoline inthe fuel mixture, operation is limited tothe following:A. Do not take off with the fuel tem-

perature lower than –54ºC (–65ºF).B. Restrict flights to below 15,000 feet.C. Both main jet pump switches and

both wing standby pump switchesmust be on and the pumps must beoperating.

D. All the above

13. The Learjet 20 series requires anti-icingadditive:A. At all timesB. Only when temperatures of –37ºC

and below are forecastC. Only for flights above 15,000 feetD. Only for flights above FL 290




5A-i

CHAPTER 5AFUEL SYSTEM

CONTENTS

Page

INTRODUCTION ................................................................................................................ 5A-1

GENERAL............................................................................................................................ 5A-1

FUEL TANKS AND TANK VENTING SYSTEM.............................................................. 5A-5

Tip Tanks....................................................................................................................... 5A-5

Wing Tanks.................................................................................................................... 5A-5

Fuselage Tank................................................................................................................ 5A-5

Ram-Air Vent System ................................................................................................... 5A-5

FUEL INDICATING SYSTEMS ......................................................................................... 5A-6

Fuel Quantity Indicating System/Low Fuel Warning.................................................... 5A-6

Fuel Flow Indicating System......................................................................................... 5A-8

FUEL DISTRIBUTION ....................................................................................................... 5A-9

General .......................................................................................................................... 5A-9

Boost Pumps.................................................................................................................. 5A-9

Motive-Flow Fuel and Jet Pumps.................................................................................. 5A-9

Filters and Pressure Switches...................................................................................... 5A-10

Main Fuel Shutoff Valves ........................................................................................... 5A-10

Low Fuel Pressure Warning Lights............................................................................. 5A-10

Pressure-Relief Valves ................................................................................................ 5A-11

Fuel Drain Valves........................................................................................................ 5A-11

FUEL TRANSFER SYSTEM............................................................................................ 5A-12

Crossflow Valve .......................................................................................................... 5A-12




Fuselage Fuel Transfer................................................................................................ 5A-12

Float, Pressure, and Temperature Switches ................................................................ 5A-12

Fuselage Fuel Fill Operation....................................................................................... 5A-13

Tip Tank Fuel Jettison System (Optional) .................................................................. 5A-13

FUEL SERVICING ............................................................................................................ 5A-14

General ........................................................................................................................ 5A-14

Safety Precautions....................................................................................................... 5A-14

Approved Fuels and Fuel Density Adjustments ......................................................... 5A-14

Aviation Gasoline........................................................................................................ 5A-15

Anti-icing Additive ..................................................................................................... 5A-15

Refueling..................................................................................................................... 5A-16

Jet Fuel ........................................................................................................................ 5A-17

QUESTIONS ...................................................................................................................... 5A-18

5A-ii FOR TRAINING PURPOSES ONLY



5A-iii

ILLUSTRATIONS

Figure Title Page

5A-1 Fuel System—SNs 23-003 through 23-069 .......................................................... 5A-2

5A-2 Fuel System—SNs 23-070 through 23-099 .......................................................... 5A-3

5A-3 Fuel System—24, 24B, and 25 Models................................................................. 5A-4

5A-4 Fuel Vent System................................................................................................... 5A-6

5A-5 Fuel Control Panels ............................................................................................... 5A-7

5A-6 Fuel Flow Indicator ............................................................................................... 5A-8

5A-7 Jet Pump Schematic............................................................................................... 5A-9

5A-8 Fuel Drain Locations........................................................................................... 5A-11

5A-9 Airplane Grounding Points.................................................................................. 5A-14

5A-10 Fuel Density Adjustment..................................................................................... 5A-15

5A-11 Anti-ice Additive Blending Apparatus................................................................ 5A-16

5A-12 Refueling Filler Cap ............................................................................................ 5A-17




INTRODUCTIONThe Learjet 20 series fuel system consists of the fuel tanks, tank venting, indicating, distribution, and transfer systems.

This chapter covers Learjet 23-003 through 24-229, and 25-003 through 25-063. Majordifferences, where applicable, will be described by serial number.

The fuel in the Learjet is stored in the integralwet wing, wet tip tanks, and a bladder fuse-lage tank. The 23 models without jet pumpshave two electrically operated low-pressureboost pumps installed at the inboard side ofeach wing fuel tank. All other Learjets and 23models with jet pumps have one electricallyoperated, low-pressure boost pump and a jetpump installed at the inboard side of eachwing fuel tank. The total trapped and unusablefuel (included in the empty weight) is ap-proximately 23 gallons.

An electric boost pump in each tip tank on 23models without jet pumps is used to transferall tip tank fuel into the respective wing tank.This is primarily for the lower portion of thetank since the top portion of the tank willgravity flow into the wing tank. All other mod-els use a jet pump for the same purpose.

Depending on model and serial number, eitherthree or five pressure switches are locatedwithin the fuel system to indicate that trans-fer of fuel has been completed or boost pump

0

2

4 6

8

10

MAINFUEL

LBS X 100

CHAPTER 5AFUEL SYSTEM




GENERAL




FIREWALLSHUTOFF

LOWFUEL

CROSSFLOW

FUS

EMPTY

EMPTY

FULLOFF

FUS TANK FILL

ON

L FUEL PRESS

FUEL BYPASS

TIP TANKEMPTY LIGHT-LH

LHTIP

TANK

LH WING TANK

MAIN FUELSHUTOFF VALVE-LH

FUELFILTER

LH ENGINE

FUELQUANTITYGAGE

CROSSFLOWVALVE

REFUELLINE

TRANSFERVALVEREFUEL VALVE

(NC)

FUSELAGE FUEL TANK

TO FUELQUANTITYGAGE

FUSELAGETANKTRANSFERSWITCH

SUPPLY

LOW PRESSURE

FUSELAGE TANKTRANSFER

LEGENDFUEL PROBE FILLER CAP ELECTRICAL

CHECK VALVE BOOST PUMP FLOAT SWITCH

SHUTOFF VALVEPRESSURESWITCH

FUSELAGETANK FILL

*SNs 23-015 THROUGH 23-069

*

Figure 5A-1. Fuel System–SNs 23-003 through 23-069

failure has occurred. Four or five motorizedshutoff valves are installed throughout thefuel system for isolating various tanks andfuel lines. Drain valves for the sumps and fuelvent system are located at the low points in thefuel system to drain any accumulation of water,sediment, and/or fuel. Two high-pressure relief valves, one in each fuel line between the

eng ine main f i r ewal l shu toff va lve and engine, are installed to relieve any pressurebuildup caused by thermal expansion oftrapped fuel when the engine is shut down.

Figures 5A-1 through 5A-3 illustrate the earlyseries Learjet fuel system.




EMPTYLOWFUEL

CROSSFLOW

FUSFILL

ON

OFF

EMPTY

FUEL FILTER

L FUEL PRESS

FUS

FULL

TIP TANKEMPTY LIGHT-LH

CROSSFLOWVALVE

LHTIP

TANK

LH WING TANKFUSELAGETANKTRANSFERSWITCHTRANSFER

VALVEFUEL

FILTER

MAIN FUELSHUTOFF VALVE

LH ENGINE

FLOWMETER

FUSELAGE FUEL TANK

FUEL COUNTER

FIREWALL

SHUTOFF

0 0 0 0

SUPPLY

LOW PRESSURE

TRANSFER/FILL

LEGENDFUEL PROBE FILLER CAP ELECTRICAL

CHECK VALVE BOOST PUMP FLOAT SWITCH

SHUTOFF VALVEPRESSURESWITCH

Figure 5A-2. Fuel System–SNs 23-070 through 23-099




FIREWALL

SHUTOFF

L

DUMP

FIRE

**

***

*

(OPTIONAL)

L FUEL PRESS

START

GEN

STARTSTART

FILL

OFF

XFER

FUSTANK

EMPTY

OPEN

CLOSECROSSFLOW

GEN GEN

OFF

JET PUMPON

LH LH

FUSELAGEFUELTANK

COUNTER

FUELSHUTOFF VALVE-LH

TRANSFERVALVEFUEL

FILTER

LHENGINE

MOTIVE-FLOWCONTROL VALVE

RELIEFVALVE

0 1 3 4

FUEL FILTER

LOW FUEL

LH WING TANK

LOWFUEL

STANDBYPUMP

ON

LH

OFF

FILLERCAP

CROSSFLOWVALVE

SUPPLY

LOW PRESSURE

HIGH PRESSURE

LEGEND

EFFECTIVITY:

FUEL PROBE

PRESSURE RELIEF VALVE ELECTRICAL

CHECK VALVE

TRANSFER/FILL

* 24 AND 24B MODELS** AIRPLANES WITHOUT GLARESHIELD LIGHTS*** 25 SERIES

BOOST PUMP FLOAT SWITCH

SHUTOFF VALVE

JET PUMP

PRESSURESWITCH

Figure 5A-3. Fuel System—24, 24B, and 25 Models

FUEL TANKS AND TANKVENTING SYSTEM

TIP TANKSThe tip tanks are positioned at 2° nose up rel-ative to the wings. Two attach points securethe tank to the wing. Access covers on the topof the tank provide entry for inspection andmaintenance. Baffles are installed to mini-mize fuel shift and prevent adverse effects ofshifting fuel on the airplane center of gravityduring extreme flight attitudes. Each tip tankholds approximately 1,235 pounds of fuel.

An electric boost pump (23 models without jetpumps) or a jet pump (all other) in each tip tanktransfers fuel into the associated wing tank.Fuel gravity flows through flapper checkvalves into the wing tank, but fuel in the lowerhalf must be transferred by the pump. A vac-uum-relief valve is installed in the left tip tankon the airplanes SNs 24-190 and 25-030 andsubsequent. This is available on earlier 20 se-ries Learjets as a retrofit. This valve preventsa vacuum in the fuel vent system.

The tip tanks contain a fuel quantity probeand two sump drains. The tank is ventedthrough two vent float valves located in the forward and aft ends of the tank. A filler caplocated on each tip tank is used to service theentire airplane fuel system.

WING TANKSEach wing tank extends from the airplane cen-terline to the tip tank and holds 1,092 poundsof usable fuel (23 model) or 1,160 pounds (allother models). Areas which are not part of thewing fuel cell are the main landing gear wheelwell, the leading edge forward of spar one (wingheat area), and the trailing edge between sparsseven and eight (flap, spoiler, and aileron areas).

The 2.5° wing dihedral makes the inboard por-tions of the wing tanks the lowest areas. Twofuel pumps are located within each of theseareas and will remain submerged in fuel untilthe tanks are nearly empty.

NOTEThe 23 models without jet pumpshave two electric fuel pumps in eachwing tank. All other Learjets haveone electric standby pump and one jetpump in each wing tank.

Wing ribs and spars act as baffles to minimizefuel shifting. Flapper valves located in the wingribs allow unrestricted inboard flow of fuel andlimit outboard flow. Two pressure-relief valvesat the centerline rib equal equalize internalpressures between the two wing tanks (except23 models). The wing tanks begin to fill throughthe two tip tank flapper valves as tip tank fuelincreases beyond one-half full.

Two fuel probes in each wing tank provide in-formation to the fuel quantity indicating system.

FUSELAGE TANKThe fuselage tank consists of two bladder-type cells located in the cavity between frames22 and 24 on 23, 24, and 24B models and be-tween frames 22 and 25 on the 25 model. Thetank contains one electric fuel transfer pump,a float switch, a fuel probe, and a drain valve.The crossover interconnects the two cells toform one fuel storage tank. Fuselage tank fuelmust be transferred to the wing tanks for use.Fuel capacity on models 23, 24, and 24B is ap-proximately 840 pounds, and on model 25,approximately 1,306 pounds.

RAM-AIR VENT SYSTEMThe fuel vent system supplies ram-air pressureto the fuel tanks through two ram-air masts.One mast is located on each side of the aft fuse-lage. Air from the left mast flows through acheck valve to the forward wing tank sump(Figure 5A-4). Air from the right mast flowsto the forward wing tank sump and to the fuse-lage tank. From the forward sump, ram airflows through vent lines to the tip tanks andto the wing tanks through the wingtip inter-connects. Relief valves limit pressure to .8 psi.The vent system prevents fuel tank overpres-surization or vacuum formation.




Two vent drain valves are located at the fuelvent sumps. (Figure 5A-4). They must bedrained prior to every flight to ensure that thevent line is unobstructed.

FUEL INDICATINGSYSTEMS

FUEL QUANTITY INDICATINGSYSTEM/LOW FUEL WARNINGThe fuel quantity indicating system includesan indicator and tank selector switch located

on the fuel control panel (Figure 5A-5). Oneor two LOW FUEL warning lights illuminatewhenever either wing tank fuel level is lowon all airplanes except SNs 23-003 through23-014 (see Figures 5A-1 through 5A-3).

The six-position rotary selector switch en-ables the pilot to check the fuel quantity ineach of the five tanks and the airplane totalfuel quantity. The fuel quantity for the posi-tion selected is read on the fuel quantity in-dicator. The quantities printed beside eachselector switch position indicate usable fuelcapacities in pounds.




VACUUM-RELIEFVALVE

PRESSURE-RELIEFVALVES

CHECK VALVE

RAM-AIR MAST

FORWARD SUMPASSEMBLY

CHECK VALVE

AFT SUMP ASSEMBLY

OVERBOARD VENT LINE

Figure 5A-4. Fuel Vent System




CLOSE

JETPUMPS

OFF

STDBYPUMPS

0 0 0 0

0

1

2

3 4

5

6

FUELQUANTITY

LBS ✕ 1000

0

1

2

3 4

5

6

FUELQUANTITY

LBS ✕ 1000

L TIP1235

LOWFUEL

LOWFUEL

L WING1160

R WING1160

R TIP1195

TOTAL5590LBS

FUS840

L TIP1235

L WING1160

R WING1160

R TIP1195

TOTAL6055LBS

FUS1305

L RON

L

L R

RON

OFF FUEL JETTISON

OPEN

CROSSFLOWFULL

OFF

EMPTYXFER

FILL

FUS TANK

LDUMP

RDUMP

24 AND 24B MODELS

23 MODEL (LATE)

23 MODEL (EARLY)

25 MODEL

L TIP1196

L WING1092 R WING

1092

R TIP1196

TOTAL5388 LBS

FUS812

0 0 0 0

L. ENG

LH ENG RH ENG

R. ENG

STDBY STDBY

FULL

ON

OFFEMPTYTIPTIP

FUSFILLMAIN MAINFUS

CROSSFLOW

LOW FUEL

STDBY STDBYMAIN MAINCROSSFLOW

FUS

ON

OFF

EMPTY EMPTY

LOWFUEL

LOWFUEL

FULL

TIP TIP

EMPTY

Figure 5A-5. Fuel Control Panels

There are seven resistance (SNs 23-003 through024) or capacitance fuel quantity probes in-stalled. One fuel probe is located in each tiptank and in the fuselage tank. Each wing tankhas two probes wired in parallel. The inboardprobe in the left wing contains a density com-pensator which adjusts quantity readings forall switch selections for fuel density change.

Each probe uses an electrical resistance or ca-pacitance measuring system to sense the fuellevel. It then transmits an electrical signal tothe cockpit indicator where it is read as pounds✕ 1,000 on the gage.

NOTEThe total position of the selectorswitch should be used for flight plan-ning purposes due to tolerances in theindividual tank quantity readings.The quantity indicating system is setto indicate properly in level flight.

On all Learjets except SNs 23-003 through23-014, one or two LOW FUEL lights are ac-tuated by each wing tank fuel low-level floatswitch when fuel quantity in either tank is ap-proximately 450 pounds. On airplanes with-out a glareshield panel, left and right LOWFUEL lights (see Figure 5A-5) are located onthe fuel panel. On airplanes with a glareshieldwarning panel, a single LOW FUEL light (see“Annunciator Panel” section) will illuminatewhen either wing fuel quantity is low.

When the fuel quantity gage indi-cates 600 pounds or less remainingin either wing tank, prolonged noseupattitude of 10° or more may causefuel to be trapped in the aft area ofthe wing tank outboard of the wheelwell . Fuel starvation and engineflameout may occur. Reducing pitchattitude and thrust to the minimumrequired will prevent this situation.

Learjets SNs 23-015 through 23-089 use 115VAC for the fuel quantity indicating system.All other Learjets use 28 VDC.

FUEL FLOW INDICATING SYSTEMA fuel flow indicator provides a readout inpounds of fuel flow per hour (Figure 5A-6).Airplanes SNs 23-003 through 23-069 have asingle indicator with two needles labeled “L”and “R,” which uses 26 VAC. All other Learjetsuse a single 28-VDC-powered indicator withone needle and a selector switch labeled “L–R”adjacent to the indicator to select either engine.

The optional fuel counter (see Figure 5A-5)located on the fuel control panel provides afour-digit readout (in pounds of fuel consumedby both engines) and uses DC power. It shouldbe reset to zero using the reset button adjacentto the counter before starting the first engine.

WARNING




Figure 5A-6. Fuel Flow Indicator

FUEL DISTRIBUTION

GENERALEach engine is supplied with fuel from its re-spective wing fuel system; there is no cross-feed capability. Either the wing standby pumpsor the wing jet pumps supply fuel under pres-sure to the engine-driven pumps.

The wing jet pumps and standby pumps havecheck valves on the outboard side to preventreverse flow when they are inactive.

BOOST PUMPSThe electric fuel boost pumps are submergedcentrifugal pumps. Powered by 28 VDC, thepumps have a maximum output of 18 psi and arecontrolled by switches on the fuel control panel.

Seven electric fuel pumps are installed onmodel 23 airplanes which have not been mod-ified with jet pumps. A main and standbypump are located in each wing tank. Thesepumps are used to supply fuel to the enginesand for filling the fuselage tank. The pump ineach tip tank transfers the lower half of the tiptank fuel into the wing tank. The transfer

pump in the fuselage tank directs fuel for-ward to the wing tanks.

NOTEOn a i rp lanes wi th e l ec t r i c fue lpumps, the standby pumps are hotwired to the battery.

Airplanes with jet pumps have three electricfuel pumps. An electric standby pump in eachwing is used for engine start and as a backuppump for the engine supply as well as forcrossflow and filling of the fuselage tank. Thetransfer pump in the fuselage tank directsfuselage fuel forward to the wing tanks.

MOTIVE-FLOW FUEL AND JET PUMPSExcess high-pressure fuel from the engine-driven fuel pumps is the source of motive-flowfuel which causes the jet pumps to function.The fuel is routed through the motive-flowcontrol valves (one on each side) to the jetpumps, where it sprays through a small orificeinto a venturi. The low pressure created in theventuri draws fuel from the respective tank, re-sulting in a low-pressure, high-volume outputfrom the jet pump (Figure 5A-7).

Revision .01 5A-9FOR TRAINING PURPOSES ONLY



INPUTOUTPUT

WING TANKSTRUCTURE

FUEL SUPPLY HIGH PRESSURE LOW PRESSURE

LEGEND

Figure 5A-7. Jet Pump Schematic

Motive-flow pressure varies with engine rpm.Consequently, jet pump discharge pressurealso varies with engine rpm. At idle, dischargepressure is approximately 10 psi, while at full-power settings, discharge pressure is approx-imately 12 psi.

There are four jet pumps—one in each wingtank adjacent to the standby boost pump andone in each tip tank. The wing tank jet pumpsdraw fuel from the wing tanks and supply low-pressure fuel to the engine-driven, high-pressure fuel pumps. Wing jet pump output canbe supplemented by the wing standby pump toensure positive pressure to an engine. The tiptank jet pumps draw fuel from the tip tanks anddeliver it directly to the cavities where thestandby pumps and jet pumps are located.

Jet pumps require no electrical power andhave no moving parts. They are controlled bytwo jet pump switches which electrically openand close the motive-flow control valves. Theamber indicator lights next to the switches il-luminate when the motive-flow valves are intransit or are not in the position selected on theswitch. Each jet pump switch (and motive-flow control valve) controls both jet pumps(wing and tip) on that side.

NOTEShould operational requirementsmake it necessary to turn off a jetpump switch, the standby pump onthe same side must first be turned onto ensure against loss of pressure tothe engine-driven pump.

NOTEWith an engine shut down, both thewing and the tip tank jet pumps areinoperative on that side.

NOTEIf a jet pump indicator light remainsilluminated after a selection is made,the motive-flow valve is not in the po-sition selected by the switch.

FILTERS AND PRESSURE SWITCHESA fuel filter is installed in each engine feedline. Should the filters become clogged fromice or contaminants, the fuel is allowed to by-pass the filters. A differential pressure switchinstalled in each filter assembly signals thecockpit when one or both filters are bypass-ing fuel. One amber FUEL FILTER annunci-a to r i l l umina t e s when th i s occu r s ( s ee“Annunciator Panel” section).

MAIN FUEL SHUTOFF VALVESThe 28-VDC motorized fuel shutoff valvesare used to isolate fuel from the engines. Theyare two-position (open-closed). The valvesare controlled by the left and right FIREswitches on airplanes with glareshield warn-ing lights and by the red guarded FIREWALLSHUTOFF switches in airplanes withoutglareshield warning lights.

To close the valves on airplanes equipped withglareshield warning lights, raise the plasticguard and push on the FIRE light. Adjacent tothe FIRE light is a red shutoff valve warninglight hot wired to the battery bus, which in-dicates the position of the fuel shutoff valveswitch, not necessarily the valve position. Ifthe FIRE light has been depressed for theclosed position, the red shutoff valve lightwill be on. To reopen the fuel shutoff valve,depress the FIRE light again and the red shut-off valve light extinguishes.

NOTEOn all Learjets except the 23 models,when the fuel shutoff valves are closed,both fire extinguishers are armed.

LOW FUEL PRESSUREWARNING LIGHTSA low fuel pressure switch is located betweenthe fuel shutoff valve and the engine-drivenfuel pump in each engine feed l ine. Theswitches cause illumination of the appropri-ate red L or R FUEL PRES annunciator light




when fuel pressure drops below 0.25 psi. Thelight extinguishes when pressure increasesabove 1.0 psi. Illumination of a FUEL PRESannunciator light is an indication of loss of fuelpressure to the engine. The probable cause isfailure of the affected wing jet pump in air-planes equipped with jet pumps.

NOTEThe engine-driven pump is capableof suction-feeding enough fuel to sus-tain engine operation without eitherthe wing standby pump or jet pump.However, 25,000 feet is the highest al-titude at which continuous operationshould be attempted in this event; thelimiting factor is that possible dam-age to the engine-driven pump mayoccur due to reduced fuel flow.

PRESSURE-RELIEF VALVESA 75-psi relief valve is installed in each main fuelline on the engine side of the main shutoff valve.

The valves relieve pressure buildup caused bythermal expansion of trapped fuel when the en-gines are shut down by venting fuel overboard.

FUEL DRAIN VALVESDra in va lves a r e l oca t ed a t l ow po in t sthroughout the fuel system for draining con-densation or sediment. A small amount offuel should be drained from each valve dur-ing exterior preflight inspection. The valves,spring-loaded to the closed position, are lo-cated as follows: two on each tip tank, one forthe crossflow valve, one for each wing, onefor each line to the engines, one for each fuelfilter, and one for the fuselage tank (Figure5A-8). On model 23, filter drains are insidethe tail compartment.

A drain is located at each fuel vent sump. Thesevalves must be completely drained during theexterior preflight inspection to prevent possi-ble blockage of the fuselage ram-air vent line.




FLUSH SUMPDRAINS

ENGINE LINEDRAIN

FORWARDFUEL VENT SUMP DRAIN

FUEL FILTERDRAINS

FLUSH SUMPDRAINS

ENGINE LINEDRAIN

FUSELAGETANK DRAIN

REAR FUEL VENT SUMPDRAIN

WING SUMPDRAINS

CROSSFLOWDRAIN

Figure 5A-8. Fuel Drain Locations

FUEL TRANSFERSYSTEM

CROSSFLOW VALVEA 28-VDC motorized valve is provided to in-terconnect the left and right wing tanks (seeFigures 5A-1 through 5A-3). It is opened dur-ing fuel transfer and filling operations and forwing-to-wing fuel balancing. The valve iscontrolled by the CROSS FLOW switch andthe FUS FILL switch (23 models) or XFERFILL switch (24, 24B, and 25 models).

An amber light adjacent to the CROSS FLOWswitch illuminates when the valve is in tran-sit or is not in the position selected on allLearjets except the 23 models.

If wing fuel imbalance occurs on airplanesSNs 23-003 through 23-069, crossflow is ac-complished by opening the crossflow valve andallowing the fuel to balance by gravity. On allother Learjets, open the crossflow valve andturn on the standby pump in the heavy wingwhile ensuring that the standby pump in thelight wing is off. The transfer rate is approx-imately 50 pounds of fuel per minute.

NOTEOpening the crossflow valve to bal-ance the fuel should not be attemptedwhen a red FUEL PRES light is il-luminated unless it can be accom-plished below 25,000 feet. To do sodiverts pressure being delivered tothe affected engine pump by the op-erating standby pump on that side.Asymmetric power settings may beused to balance fuel, if necessary.

The above considerations do notapply to single-engine operations,and normal crossflow operations maybe performed.

FUSELAGE FUEL TRANSFERFuel in the fuselage tank must be transferred for-ward to the wing tanks to supply the engines.

To prevent overpressurization of the wing tanksand the tip tanks, the fuel quantity in the tiptanks must be decreased before fuel transfer.

On model 23 airplanes not incorporating jetpumps, when the fuselage transfer (FUS) switch(see Figure 5A-1) is moved to the ON position,the fuselage tank transfer valve and the cross-flow valve are opened, and the fuselage tankpump is energized to pump fuel equally intotwo wing tanks. The fuselage tank EMPTY light(see “Annunciator Panel” section) is illuminatedwhen fuel pressure in the transfer line dropsbelow the setting of a pressure switch in thetransfer line. When the EMPTY light is on, thefuselage transfer switch should be turned off toclose the valves and shut down the transfer pump.

On airplanes SNs 23-070 through 23-099, thestandby pumps switches (see Figure 5A-5)must be in the OFF position when transferringfuselage tank fuel.

On 24, 24B, 25, and 23 models with jet pumps,when the fuel transfer switch (see Figure 5A-3)is moved to XFER, the crossflow valve andfuselage transfer valves are opened and thefuselage tank pump is turned on. If operating,the standby pumps are automatically shutdown to allow fuel to be pumped to the wingtanks. When pressure drops in the transferline, the fuselage tank EMPTY light (see“Annunciator Panel” section) will illuminate.When the tank is empty, the transfer switchshould be moved to the OFF position to closethe valves and shut off the fuselage tank pump.

FLOAT, PRESSURE, ANDTEMPERATURE SWITCHES

Fuselage Tank Float SwitchA float switch in the fuselage tank prevents over-filling of the tank by terminating the fill operation.

On airplanes SNs 23-003 through 23-069,when the float switch senses a full fuselagetank, the refuel valve is closed, the left standbypump is shut down, and the FULL light (see“Annunciator Panel” section) on the fuel panelis illuminated. The crossflow valve is not af-fected by the float switch.




On airplanes SNs 23-070 through 23-099, 24,24B, and 25 models, when the float switch de-tects a full fuselage tank, the transfer and cross-flow valves are closed, both standby pumps areturned off, and the FULL light is illuminated.

Fuselage Tank Low-Pressure SwitchThe fuselage tank low-pressure switch, lo-cated in the fuselage tank transfer line, closesand t u rn s on t he EMPTY l i gh t ( s ee“Annunciator Panel” section) when pressuredrops below 2.75 psi. The switch does not au-tomatically close the transfer valve or shutdown the fuselage tank pump. The EMPTYlight indicates:

• The tank is empty.

• The fuselage transfer pump has failed.

Tip Tank Low-Pressure SwitchOn model 23 airplanes with electric pumps, alow-pressure switch in the output line fromeach tip tank fuel pump turns on the EMPTYwarning light (see Figure 5A-1) to indicate thatthe tank is empty or that the pump has failed.

Fuel Temperature Warning SwitchOn model 23 airplanes with a fuel temperaturewarning system installed, a thermal switch ineach wing tank sump turns on a HOT FUEL lightlocated on the fuel control panel. The light willbe on if fuel temperature increases to 82°F andwill go off when fuel temperature drops to75°F. The airplane shall not be flown at altitudesabove 25,000 feet if the HOT FUEL light is on.

FUSELAGE FUEL FILL OPERATION

Model 23 without Jet PumpsThe fuselage tank refuel valve on airplanes SNs23-003 through 23-0069 is open only when thefuselage tank is being filled. This valve is lo-cated in the refuel line between the left wing

tank and the fuselage tank. The fuselage refuel valve is controlled by the FUS FILLswitch (see Figure 5A-1).

Fuel may be pumped from the wings to thefuselage tank using the FUS FILL switch setto FILL. This turns on the left wing electricstandby boost pump (23-003 through 23-069)or the left and right wing electric standbyboost pumps (SNs 23-070 through 23-099). Inaddition, the CROSS FLOW switch must be setto OPEN with the battery switch on to allowfuel to be taken from the right wing.

Model 24, 24B, 25, and Model 23 with Jet PumpsOn these airplanes, the left and right wingelectric standby boost pumps fill the fuselagetank when the fuselage tank XFER–FILLswitch is set to FILL. In addition, the cross-flow valve and fuselage transfer valves openautomatically and the fuselage float switch isarmed. This allows fuel to be taken from bothwings to fill the fuselage tank.

If the tank is to be filled to capacity, the floatswitch actuation automatically:

• Deenergizes the wing standby pumps

• Closes the crossflow and transfer valves

• I l luminates the green FULL l ight , which will remain l ighted until theXFER–FILL switch is turned off

The filling process may be terminated at anypoint by turning the XFER–FILL switch off.

TIP TANK FUEL JETTISONSYSTEM (OPTIONAL)This system is an option on SNs 25-030 andsubsequent. Two 28-VDC shutoff valves areprovided for dumping tip tank fuel in case ofan emergency. Two guarded DUMP switches(fuel jettison) are installed. When the DUMPswitch(es) is placed to the DUMP or ON po-sition, the shutoff valves are opened and anamber light comes on to indicate that thevalves are open. Dumping may be started orstopped at any time.




FUEL SERVICINGGENERALFuel servicing includes those procedures neces-sary for refueling and adding anti-icing additive.

Fueling is accomplished through a filler inthe top of each tip tank. When gravity refuel-ing, fuel flows from the tip tanks into the wingtanks. The fuselage tank is filled by use ofthe wing electric standby boost pump(s).

The wide range of temperatures to which jetfuels are exposed in flight results in sub-stantial changes in the water solubility ofthe fuel and, consequently, in the amount offree water which must be coped with by theblending with anti-ice additives. The addi-tive functions as a freeze-point depressant.Another important characteristic of this jetfuel additive is its action as a biocidal agentwhich prevents fungal and bacterial growthin fuel systems and fuel tanks. Thus, whenused in proper amounts, it provides bothanti-icing and microbial growth protectionin a single additive.

Prolonged storage of the airplane will resultin a water buildup in the fuel which “leachesout” the additive. An indication of this is whenan excessive amount of water accumulates inthe fuel tank sumps. Refer to Section II,“Normal Operating Procedures,” of the AFMfor checking fuel additive.

SAFETY PRECAUTIONSThe airplane should be fueled only in areaswhich permit free movement of fire equipment.

Assure that the fuel truck is grounded to an ap-proved grounding source. Attach a groundfrom fuel truck to airplane nose gear uplatchroller and ground fuel nozzle to tip tank groundjack (Figure 5A-9).

APPROVED FUELS AND FUELDENSITY ADJUSTMENTS

Approved FuelsRefer to Section I of the AFM for a listing ofapproved fuels. Jet A, Jet A1, Jet B conform-ing to ASTM ES 2-74, “Emergency Standard Specification for Aviation Turbine Fuels,” maybe used as temporary fuels until further notice.

Fuel Density Adjustments

If optimum engine acceleration is desired, thedensity adjustment knob on the front of the fuelcontrol unit should be readjusted to the fueldensity setting for the fuel type being used.Refer to the AFM for recommended fuel con-trol adjustment knob positions (density setting)for each fuel type.




Figure 5A-9. Airplane Grounding Points

Engine stall margin may be reducedduring high-altitude operation if thefuel control density adjustment knobis positioned to a lower density set-ting than that recommended for typeof fuel being used.

Either of two types of fuel control units maybe installed on the engines. The fuel controldensity adjustment knobs are different on eachof these units and are designated Type A andType B (Figure 5A-10). To make an adjust-ment, push the knob in, rotate the knob to thedensity setting desired, and release pressure.If necessary, jiggle the knob until it locks.The adjustment knob must be seated in its po-sitioning detent for satisfactory operation.

AVIATION GASOLINEAviation gasoline (lowest octane available)may be used as an emergency fuel provided itsuse is limited to no more than 25 hours dur-ing one engine overhaul period. Keep an ac-curate record of engine operation time whenusing this fuel. Mixing aviation gasoline withjet fuel is allowed; all operation with avia-tion gasoline in excess of 50% by volume mustbe recorded in the Engine Log.

NOTECertain aviation gasoline and jetfue l ra t ios can produce grea te r

flammability hazards under partiallyfilled airplane tank or storage tankcond i t i ons encoun te r ed du r ing refueling or mixing. Follow estab-lished safety standards to eliminateany possible sparks.

ANTI-ICING ADDITIVEAnti-icing additive conforming to MIL-I-27686 must be added to all fuels for Learjetoperations. The additive concentration shallbe a minimum of .06% to a maximum of .15%by volume.

Assure that the additive is directed into theflowing fuel stream and that the additive flowis started after fuel flow starts and is stoppedbefore fuel flow stops (Figure 5A-11). Due tothe chemical composition of fuel additives,improper blending can cause deterioration offuel tank interior finishes and promote corro-sion. Do not use less than one container (20fluid ounces) of additive per 260 gallons of fuelor more than one container of additive per 104gallons of fuel. The blender (20 ounces) willdischarge completely in approximately fourminutes. Monitor the blender tube to assure thatadditive flow is maintained. Do not allow con-centrated additive to contact the coated inte-rior of the fuel tank or the airplane’s paintedsurfaces.

NOTEAnti-ice additives may be harmful ifinhaled or swallowed. Use adequateventilation. Avoid contact with skinand eyes. If sprayed into eyes, flushwith large amounts of water and con-tact a physician immediately.

NOTELack of anti-icing additive may causefuel filter icing and subsequent en-gine flameout.

CAUTION




75 7 2

69

86813

78

FUEL ADJUST—P

US

H&

TU

RN

JP-5&

6

JP-4

–1 +1

+2+

3+4

—6—5

—4

—3

—2

INDEXPIN

INDEXPIN

TYPE BTYPE A

Figure 5A-10. Fuel Density Adjustment

REFUELING

Gravity RefuelingRefueling is accomplished through the tiptank filler caps (Figure 5A-12). The fuel be-gins to flow by gravity into the wing tanks asthe tip tanks reach one-half full. The wingstandby pumps are used to fill the fuselagetank. (See Fuel Transfer System, this chapter.)

Do not attempt to completely fill onetank before adding fuel to the oppo-site tank because excessive lateralunbalance will occur. When fueling,

fill both tip tanks simultaneously oralternately add 125 gallons (472.3liters) of fuel to each tip tank until thedes i red amount i s obta ined. OnModel 23, the limits are 150 gallons(567.8 liters).

The fuselage tank may be filled after landingor while taxiing to the parking area to conservebattery power. If not accomplished in the abovemanner, a GPU should be used.

When the airplane is full of fuel, the airplane’sCG is near the aft limit, and the use of a tailstand should be considered.

CAUTION




MIL-I-

27686

FUEL ADDITIVE

TRIGGER

RING

HANDLE

BLENDER HOSE

FUEL NOZZLE

Figure 5A-11. Anti-ice Additive Blending Apparatus

JET FUEL• Maximum altitude with a HOT FUEL

light on is 25,000 feet on airplanes witha fuel temperature warning system.

• Maximum time with a low FUEL PRESSlight on is 10 hours between overhauls.

• On SNs 23-070 through 23-099, do nottransfer or crossflow fuel if below 5,000feet and engine rpm is above 95%.

With the LOW FUEL warning light illumi-nated, limit pitch attitude to the minimum required for go-around.

Maximum fuel in tip to begin transfer is 760pounds for model 25, and 800 pounds for mod-els 23, 24, and 24B.

With wing jet pumps (main fuel boost pumpson model 23) and standby pumps inoperative,the engine-driven fuel pumps suction-feedsufficient fuel to supply the engines at altitudesof 25,000 feet or below.

Do not energize fuselage transfer system whenwing and tip tank are full.

Do not crossflow with wing jet pump inoperative.

Do not transfer fuel if crossflow valve fails toopen.

Maximum tip tank fuel for landing is 800pounds in each tip for models 24, 24B, and 25,and 640 pounds for model 23.

A minimum of 600 pounds of fuel in eachwing tank is required for takeoff and inten-tional go-around.

If either wing tank is below 600 pounds, limitprolonged pitch attitude to 10° noseup.

Minimum fuel temperature for takeoff is:

• Jet B/JP-4 .................... –54°C (–65° F)

• Jet A/Jet A-1 ............... –29°C (–20° F)

• AVGAS ........................ –54°C (–65° F)




Figure 5A-12. Refueling Filler Cap

1. Approximately 23 gallons of fuel in the20 series models is trapped. This fuelweight:A. Must be added to the weight of the

fuel taken on board when servicingthe airplane

B. Is included in the airplane basicweight for airplanes certified in theUnited States

C. Must be accounted for in the fuse-lage tank for CG purposes

D. May be disregarded since it is lessthan 200 pounds

2. With the exception of the FUEL JETTI-SON lights, all other amber lights on thefuel control panel, when illuminatedsteady indicate that the respective:A. Valves are cycling or the pumps are

properly operating.B. Valves are in the correct position;

the pumps are inoperative.C. Switch position agrees with the

valve position or pump operation.D. Valve position disagrees with the

switch position.

3. The red LOW FUEL light illuminateswhen:A. 350 pounds total fuel remains.B. 250 to 350 pounds remain in either

wing, depending on the airplane serial number

C. Approximately 450 pounds totalfuel remains.

D. Approximately 450 pounds remain ineither wing.

4. The wing standby pumps are used forall of the following functions except:A. Engine startB. As a backup for the main jet pumpsC. Fuselage-to-wing fuel transferD. Wing-to-fuselage transfer of fuel

5. The crossflow valve opens:A. Only when the CROSS FLOW

switch is set to OPENB. Only when the CROSS FLOW

switch is set to OPEN or theXFER–FILL switch is set to XFER

C. Anytime electrical power is lostD. Whenever the CROSS FLOW, XFER–

FILL, or FUS FILL switches aremoved f rom the OFF o r CLOSE position

6. Steady illumination of an amber transfervalve light indicates:A. The valve has failed closed.B. The valve has failed open.C. The valve has stuck in an intermediate

position.D. The valve has failed to move to the

position commanded by theXFER–FILL switch.

E. Any of the above

7. Illumination of the red L or R FUELPRES light indicates:A. Fuel pressure to the respective en-

gine-driven fuel pump is low.B. Fuel pressure to the respective en-

gine is too high for safe operation.C. A fuel filter is bypassing.D. Fuel pressure to the respective engine

is optimum for engine start.




QUESTIONS

8. When the XFER–FILL switch is placedto the FILL position, the:A. Float switch is disabled.B. Wing standby pumps are disabled.C. Fuselage valve closes.D. Crossflow valve opens.

9. Motive-flow fuel for the jet pumps issupplied by the:A. Engine-driven fuel pumpsB. Wing standby pumpsC. Fuselage transfer pumpD. Motive-flow control unit

10. The amber FUEL FILTER light indicates:A. Low fuel pressure to the engine-

driven pump; the standby pumpsshould be turned on

B. That both fuel filters are being bypassed; the light does not illuminateif only one filter is bypassed

C. That one or both fuel filters are by-passing fuel

D. That only the secondary fuel filters arebeing bypassed

11. When using any aviation gasoline in thefuel mixture, operation is limited to thefollowing on airplanes with jet pumps:A. Do not take off with fuel tempera-

ture lower than –54°C (–65°F).B. Restrict flights to below 15,000 feet.C. Both JET PUMPS switches and both

wing STANDBY PUMPS switchesmust be on and the pumps must beoperating.

D. All the above.

12. The Learjet 20 series airplanes requireanti-icing additive:A. At all timesB. Only when temperatures of –37°C

and below are forecastC. Only for flights above 15,000 feetD. Only for flights above FL 290

13. On 23 models without jet pumps, tiptank fuel is transferred by:A. Inline boost pumpB. GravityC. Jet pumpsD. Submerged electric boost pumps

14. On 23 models without jet pumps, tank-to-engine fuel pressure is normally pro-vided by a:A. Jet pump in the main tankB. Submerged electric boost pumpC. Boost pump in the tank-to-engine

feed lineD. Boost pump driven by the engine

15. The crossflow valve permits:A. Pressurized crossfeed to the oppo-

site engineB. Fuel transfer between the tip tanksC. Pressurized flow from the fuselage

tank to either engineD. Gravity balance between wing tanks

16. Aviation gasoline (AVGAS) may be used:A. For not more than 25 hours between

engine overhaulB. In unlimited quantity continuouslyC. Continuously, if mixture ratio is

less than 50%D. If the total fuel does not contain Prist




This information normally contained in this chapter is not applicable to thisparticular airplane.

7-i

CHAPTER 7POWERPLANT

CONTENTS

Page

INTRODUCTION ................................................................................................................... 7-1

GENERAL............................................................................................................................... 7-1

Engine Parameters............................................................................................................ 7-2

ENGINE .................................................................................................................................. 7-2

General ............................................................................................................................. 7-2

Major Sections ................................................................................................................. 7-2

Engine Nacelles ............................................................................................................... 7-7

Operating Principles......................................................................................................... 7-8

ENGINE SYSTEMS ............................................................................................................... 7-8

General ............................................................................................................................. 7-8

Oil System........................................................................................................................ 7-8

Fuel System ................................................................................................................... 7-11

Ignition System.............................................................................................................. 7-13

Air Systems.................................................................................................................... 7-13

ENGINE POWER CONTROL ............................................................................................. 7-15

General........................................................................................................................... 7-15

Friction Control ............................................................................................................. 7-15

ENGINE INSTRUMENTATION.......................................................................................... 7-16

General........................................................................................................................... 7-16

Pressure Ratio ................................................................................................................ 7-16

Exhaust Temperature ..................................................................................................... 7-16




Engine RPM................................................................................................................... 7-17

Fuel Flow ....................................................................................................................... 7-17

Oil Pressure.................................................................................................................... 7-17

Oil Temperature ............................................................................................................. 7-17

ENGINE STARTING............................................................................................................ 7-17

ENGINE SYNCHRONIZING .............................................................................................. 7-17

Woodward Electronic Synchronizer (Optional) ............................................................ 7-17

Synchroscope (Optional) ............................................................................................... 7-18

THRUST REVERSERS (OPTIONAL) ................................................................................ 7-18

General........................................................................................................................... 7-18

Control and Indication ................................................................................................... 7-19

Automatic Stow ............................................................................................................. 7-19

Emergency Stow ............................................................................................................ 7-19

DIFFERENCES..................................................................................................................... 7-20

QUESTIONS......................................................................................................................... 7-21

7-ii FOR TRAINING PURPOSES ONLY



7-iii

ILLUSTRATIONS

Figure Title Page

7-1 CJ610 Engine (Typical) Major Sections and Gas Flow........................................... 7-3

7-2 Front Frame Section................................................................................................. 7-2

7-3 Compressor Section.................................................................................................. 7-4

7-4 Mainframe Section ................................................................................................... 7-5

7-5 Combustion Section ................................................................................................. 7-6

7-6 Turbine Section ........................................................................................................ 7-7

7-7 Oil Servicing Access ................................................................................................ 7-8

7-8 Center Instrument Panel ........................................................................................... 7-9

7-9 Oil System Schematic ............................................................................................ 7-10

7-10 Flow Divider and Nozzle ....................................................................................... 7-12

7-11 Fuel System Schematic .......................................................................................... 7-14

7-12 Center Switch Panel ............................................................................................... 7-13

7-13 Center Console....................................................................................................... 7-16

7-14 ENG SYNC Switch................................................................................................ 7-17

7-15 Thrust Reversers..................................................................................................... 7-18

7-16 Thrust Reverser Control, Test, and Indicating Panel (Typical) ............................. 7-19




INTRODUCTIONThis chapter describes the powerplant installed on Learjet 20 series airplanes. In addi-tion to the powerplant, the chapter describes such engine-related systems as oil, fuel,ignition, engine controls and instrumentation, engine synchronization, and the DeeHoward thrust reversers.

GENERALThe airplane is powered by two GeneralElectric, CJ610 single rotor, axial-flow turbo-jet engines. The engines are installed in pylon-mounted nacel les . The engine assemblyconsists of an eight-stage, axial-flow com-pressor, driven by a two-stage turbine, an an-nular combustion section, and a fixed-areaconcentric exhaust cone. Each engine is at-tached to the engine beam at three attachmentpoints. Eight thermocouples generate an av-erage value of exhaust gas temperature, indi-cated on the EGT gages in the cockpit. Anengine-driven starter-generator is installed on

each engine to provide engine starting andfurnish DC power to the airplane’s electricalsystem. Compressor discharge air (bleed air),through a mainframe port on each engine, isutilized for cabin pressurization and cabinheating, windshield defogging, wing anti-ice,engine front frame anti-icing, and, on models24 -297 and subsequen t and 25 -135 ,-181, and subsequent, for hydraulic reservoirpressurization. On 24-350, -352, and subse-quent and 25-227 and subsequent, bleed air isalso used for windshield/radome alcohol andcabin temperature/pressurization control.

#1 DCGEN

CHAPTER 7POWERPLANT




ENGINE PARAMETERS1. Model............................. General Electric

CJ610-1—2,850 pounds thrustGeneral Electric CJ610-4—2,850 pounds

thrustGeneral Electric CJ610-6—2,950 pounds

thrustGeneral Electric CJ610-8A—2,950

pounds thrust2. Type........................... Axial-flow turbojet

3. Compressor .................. Eight-stage, axialflow with variable

interstage bleed

4. Turbine .................................... Two-stage

5. Engine main bearings ..................... Three

6. Engine rotation ............. Clockwise whenviewed from rear looking

forward; at 100%, 16,500 rpm

7. Fuel............................... Consult the AFMfor a list of approved fuels.

8. Oil................................. Consult the AFMfor a list of approved oils.

9. Engine length........... Approximately 40.5inches from front frame

through exhaust cone flanges

10. Engine diameter(maximum) .................... Approximately

17.56 inches

11. Engine weight .................. Approximately389 pounds

ENGINEGENERALThe CJ610 series engine is a single-spoolturbojet designed as a power unit for small- tomedium-sized corporate airplanes.

MAJOR SECTIONSFor descriptive purposes, the engine (Figure7-1) is divided into seven major sections,as follows:

1. Front frame

2. Compressor

3. Mainframe

4. Combustion

5. Turbine

6. Exhaust

7. Accessory

Front FrameThe front frame (Figure 7-2) admits air to thecompressor, directing air to the first-stagerotor blades. The No. 1 bearing is mounted inthe front frame.

The front frame includes an inner and outershell joined by fifteen hollow struts. A mani-fold around the front frame allows eighth-stage bleed air to flow through the struts, thebullet nose, and the inlet guide vanes for anti-ice protection.




ANTI-ICINGMANIFOLD

SUMPCOVER

BULLETNOSE

ACTUATING RING

NO. 1 BEARING

CARBON SEAL

SHROUDAND SEAL

INLETGUIDEVANE

Figure 7-2. Front Frame Section

The No. 1 bearing housing is on the aft sideof the inner shell of the front frame. The bear-ing is a spherical, self-aligning roller bearingdesigned to absorb any misalignment of theforward compressor shaft. Three of the frontframe struts are oversized and contain the No.1 bearing service lines. The ten o’clock struthouses the bearing oil supply tube, the two o’-clock strut houses a sump vent tube, and thesix o’ clock strut houses an oil scavenge tube.

A variable inlet guide vane is mounted aft ofeach of the front frame struts. The outer endsof the movement of the actuator ring will causemovement of the inlet guide vanes to direct air-flow to the first-stage rotor blades at the properangle. The actuator ring and inlet guide vanes

are positioned by the variable geometry actu-ators, which utilize high-pressure fuel from theengine fuel control unit.

CompressorThe compressor (Figure 7-3) is an eight-stageaxial-flow unit consisting of rotors and stators.It functions to increase the pressure of theinlet air and direct the airflow rearward. Thestator casing assembly which surrounds thecompressor rotor assembly consists of a com-pressor casing upper half and lower half withseven stages of stator vanes. The compressorcasing is an annular chromalloy component,split and flanged along the horizontal cen-terline and reinforced externally by integral




EXHAUST

COMBUSTION AIR

COMPRESSOR AIR

AMBIENT AIR

LEGEND

FRONT FRAMESECTION

COMPRESSORSECTION

MAINFRAME

SECTION

COMBUSTIONSECTION

TURBINESECTION

EXHAUSTSECTION

ACCESSORIES SECTION

Figure 7-1. CJ610 Engine (Typical) Major Sections and Gas Flow

stiffening ribs. The upper and lower halves arebolted together along the horizontal centerline.The casing is bolted to the front frame at the for-ward end and to the mainframe at the aft end.

The compressor rotor assembly is made up ofan integrated front shaft-disc-spacer assembly,a drive shaft, spacers, and eight stages of rotordiscs and blades.

Stall and surge protection is provided for thecompressor by the variable geometry system.Variable inlet guide vanes operate in conjunc-tion with bleed valves at the third, fourth, andfifth compressor stages to provide a safe stall-surge margin during acceleration, deceleration,and steady-state operation.

MainframeThe mainframe (Figure 7-4) section is the mainstructural support of the engine. It consists ofa frame which houses the power takeoff hous-ing and drive assembly and provides a mountfor the twelve fuel flow divider fuel nozzles,the fuel manifolds, two eighth-stage air leak-age check valves, and the gearbox assembly.The eighth-stage stators and exit guide vanesare mounted on twelve segments dovetailedinto a track cut into the front of the mainframe.The No. 2 bearing is enclosed in a bearing sup-port which bolts to the mainframe.




CARBON SEAL RUNNERSPACER

NO. 2 BEARING INNER RACEFORWARD HALF

COMPRESSOR DRIVE SHAFT

AIR DUCT

8TH-STAGEAIR SEAL

ROTOR DISC

BLEED PORTS

FORWARD AIRSEAL

NO. 1 BEARINGINNER RACE

DISC SHAFT

CARBONSEAL

RUNNER

OIL SLINGER

BALANCEWASHER

ROTOR BLADE

STATOR CASINGSTATOR VANE

Figure 7-3. Compressor Section

The mainframe has inner and outer casingbands joined by six hollow struts. Each strutends in a pad on the outer casing band. Thesepads allow passage of oil and air for the lubeand vent system and provide a mount for thepoppet valves that control the eighth-stage

seal air leakage pressure level. A two-piecefuel manifold is mounted around the main-frame with a fuel drain valve connecting thesections at the six o’clock position.The en-gine is attached to the pylon by mounting padson the mainframe.




Figure 7-4. Mainframe Section

EXIT GUIDE VANES

P.T.O. HOUSING

SHAFTGEAR DRIVER

NO. 2 CARBON SEAL SUPPORT

NO. 2 CARBON SEAL

INSULATION BLANKET

FUEL NOZZLE PAD

NO. 2 BEARING

HEAT SHIELD

NO. 2 BEARING SUPPORT

EIGHTH-STAGE OUTER AIR SEAL

EIGHTH-STAGE INNER SEAL

VANE RETAINER

CombustionThe combustion section (Figure 7-5) is thatportion of the engine in which fuel is addedto the compressed air and ignited.

The section consists of the outer and innercasing, combustion liner, No. 3 bearing and itssupport, and the first-stage turbine nozzle.

The outer combustion casing is bolted to themainframe at the front and the turbine casingat the rear. Two igniter plugs are mounted inthe casing at the one and seven o’clock posi-tions. Fuel trapped in the combustion sectionduring engine shutdown drains through a com-bustion drain at the six o’clock position.

The inner combustion casing is one piece. At thefront, it is bolted to the inner casing of the main-frame, at the rear, to the No. 3 bearing support.

The combustion liner consists of a cowl section,a dome section, and an outer and inner shell.Air enters the combustion dome through swirlcups to support combustion and provide domecooling. Thimble holes direct air into the burn-ing area, while louvers provide a boundarylayer of cool air along the inner surface of theliner.

The No. 3 bearing and carbon seal are at theaft end of the inner combustion casing. The No.3 bearing support is part of the balance pistonchamber and is used to resist forward thrustof the engine shaft.




LINERIGNITER BOSS

SWIRL CUP

HAFT SHIELD

INSULATION BLANKET

FIRST-STAGE TURBINE NOZZLE

INNER CASING

OUTER LABYRINTH SEAL

INNER LABYRINTH SEAL

NO. 3 CARBON SEAL

NO. 3 BEARING

OUTER CASING

COMBUSTOR DRAIN BOSS

Figure 7-5. Combustion Section

The first-stage turbine nozzle has an outerand an inner band joined by hollow partitions.The vanes of the nozzle direct the gases fromthe combustion section onto the forward tur-bine rotor at the proper angle. Combustor cool-ing air is forced into the turbine nozzle vanes,forming a boundary layer to cool the vanes.

TurbineThe turbine section (Figure 7-6), consistingof the turbine stator casing assembly and atwo-stage turbine rotor assembly, utilizes com-bustion air from the combustion section todrive the turbine.

The turbine stator casing assembly consistsof several components mounted on an annu-lar outer casing that is split and flanged alongthe horizontal centerline. Shrouds for boththe first- and second-stage turbine wheelsand the second-stage turbine nozzles are in-stalled in the stator casing.

The major components of the turbine rotorassembly are two turbine wheel assemblies, aturbine interstage seal assembly, a torque ringassembly, and standard hardware. The first-stage turbine wheel is integral with an inter-nally splined shaft which supports the entireassembly on the engine drive shaft.

Baffles are inserted between turbine bladeshanks to prevent crossflow of gases and todampen vibration.

ExhaustThe exhaust section is formed by the outercasing and an exhaust cone. The exhaust sec-tions function to direct the combustion gasesfrom the engine.

Exhaust thermocouples and engine pressureratio (EPR) probes are mounted on the ex-haust cone outer casing.

Accessory SectionThe engine accessories are driven through ashaft gear driver splined to the aft end of thecompressor drive shaft. A bevel gear on thedriver shaft engages a gear on a radial driveshaft which transmits power to accessorygearbox. The accessory gearbox under theengine mainframe is the mount for the fuelpump and fuel control, the overspeed gover-nor, the oil pump and tachometer generator,the hydraulic pump, and the starter-generator.On CJ610-1 engines, the accessory gearbox ismounted under the front frame, receiving powerthrough a transfer gearbox and a transfer shaft.

ENGINE NACELLESEach engine nacelle assembly consists of foursections: a nose intake, a bottom section, a topsection, and a tailcone. Each section, inter-locked and secured with fasteners, is easily re-moved fo r eng ine ma in t enance and /o rinspection. Air scoops are installed in the na-celles for the starter-generator and enginecooling. An access plate, secured by fasteners,is installed in the top nacelle section for ac-cess to the engine oil filler and dipstick.




Figure 7-6. Turbine Section

SECOND-STAGETURBINENOZZLE

FIRST-STAGETURBINE WHEEL

OUTER CASING

OUTER LABYRINTH SEAL

INNERLABYRINTH

SEAL

SEALRUNNER

NO. 3BEARING

INNERRACE

BALANCEPISTON

CHAMBER

SECOND-STAGETURBINE WHEEL

OPERATING PRINCIPLESFigure 7-1 illustrates that the air entering thefront frame through the nacelle air inlet is di-rected to the compressor at the proper angleby the guide vanes. The air pressure is in-creased by a diffusion process as it passesthrough the eight-stage compressor.

The high-pressure air leaving the compressoris directed rearward through the mainframe tothe combustion section. A large portion of thisairflow is used for cooling and bearing sealpressurization. A precise volume of the airflowenters the combustion chamber where fuel isadded by the nozzles. The mixture is ignitedby the high-energy igniter plugs. The accel-eration and expanding gases flow rearwardtoward the turbine. The two-stage turbine ex-tracts sufficient energy from the gases to drivethe compressor and the accessories. The re-maining energy is directed through the ex-haus t to the a tmosphere to p rov ide thepropulsive force for the airplane.

ENGINE SYSTEMS

GENERALThe engine systems include the oil system, fuelsystem, ignition system, engine power control,engine instrumentation, engine starting, enginesynchronizing, and the thrust reverser system.

OIL SYSTEM

GeneralThe oil system utilizes a pressurized, closed-circuit, recirculating, dry sump system de-signed to furnish lubricating and cooling oil tothe necessary rotating components during en-gine operation. After circulating to those partsrequiring lubrication, it drains to the sumpswhere it is scavenged by individual elementswithin the pump and returned to the oil tank.

Oil TankThe oil is contained in an engine-mounted tankthat includes a dwell chamber for deaerating theoil, a filler port and dipstick, a vent relief valve,and overboard vent line. Tank capacity is fourquarts, and the dipstick is marked in pints.

An access (Figure 7-7) is provided on the na-celle for oil level checking and servicing.

Oil PumpThe oil pump contains a pressure element andfive scavenge elements and is attached to the ac-cessory gearbox. The tachometer generator is at-tached to and driven by the oil pump. Allscavenge oil from the pump is discharged intothe oil tank dwell chamber. Lube oil from the oiltank enters the pressure element through a pen-dulum-type swivel pickup tube. It is deliveredby the pressure element directly to the oil cooler,through the oil filter, and into the gearbox whereit is distributed throughout the system.




Figure 7-7. Oil Servicing Access

Oil CoolerA fully automatic oil cooler transfers the oilheat to cold fuel passing through the cooler.

The oil cooler is a liquid-to-liquid heat ex-changer mounted on the oil tank mounting flangeand against the right side of the oil tank. It con-sists of numerous longitudinal passages arrangedin a honeycomb pattern. Both fuel and oil flowsimultaneously through adjoining passages,and an exchange of heat occurs between hotengine oil and cool fuel. Oil enters and leavesthe cooler through ports located in the housingfor the pressure bypass valve.

The bypass valve provides protection against highpressure due to oil viscosity or cooler blockage.

Oil FilterThe oil filter assembly is mounted in the pumphousing. It is a full-flow, in-line filter with acorrosion-resistant screen element of corru-gated steel. The filter filters out oil contami-nants over 40 microns in size and is removablefor cleaning. A filter bypass valve is includedin the core of the filter element. If the pres-sure difference between oil entering the filterand oil leaving the filter exceeds 20–24 psi,the valve opens to permit a direct flow of oilthrough the unit without filtration.

IndicationOil pressure is sensed by an AC transmitter andsent to a dual-needle gage on the instrumentpanel (Figure 7-8).

Oil pressure is also sensed by a pressure switchthat will turn on the LO OIL PRES light (see“Annunciator Panel” section) if oil pressure isbelow 5 psi. The oil pressure gages must bechecked to determine the problem engine.

NOTELearjet Model 23 airplanes do nothave low oil pressure warning lights.

Oil temperature is sensed by a resistance bulbin the oil tank and transmitted to separate leftand right OIL TEMP gages located on the in-strument panel. The oil pressure gages require26 VAC. The OIL TEMP gages and the LO OILPRES lights require 28 VDC.

OperationFigure 7-9 illustrates operation of the engineoil system.




FOR TRAINING PURPOSES ONLYFOR TRAINING PURPOSES ONLY

OILPRESS

L R

PRESS

RATIO

3.0

2.5

1.5

1.0

2.0

PRESS

RATIO

3.0

2.5

1.5

1.0

2.0

EXH.TEMP

°C x 100

109

8

7

65

4 32

1

0

1020

PERCENTRPM

30

4050

607080

90

100

0 12

3456

78

910

20

PERCENTRPM

30

4050

6080

90

100

0 12

3456

78

9

2

3

1

100

150

P.S.I.

200

50

FUEL FLOW

Lbs/Hr x 10000 L

EXH.TEMP

°C x 100

109

8

7

65

4 32

1

0

R

Figure 7-8. Center Instrument Panel




0 00000 000

FILLER PORTAND DIPSTICK

OILTANK

BEARING

BEARING

BEARING

FRONTFRAME

OIL INLET

TURBINE ROTOR

FRONT FRAME SUMP VENT

EIGHTH-STAGE LEAKAGEAIR BLEED HOLE

OIL TEMPINDICATOR

DWELLCHAMBER

COMPRESSORROTOR

DRIVE SHAFT

ACCESSORYDRIVE SHAFT

NO. 3 BEARINGSCAVENGE

VENT RELIEF VALVE

OIL TEMPBULB

LUBE ANDSCAVENGE

PUMP

RELIEF VALVE

BYPASSVALVE

OIL FILTER

BYPASSVALVE

OILCOOLER

DRIVEEND

ACCESSORYGEARBOX

OIL VENT

LOW OIL PRESSURELIGHT

OIL PRESSURE SWITCH

OIL PRESSURE GAGE

LO OIL PRESS

ENGINE NACELLE BULKHEAD

OIL PRESSURE TRANSMITTER

GEARBOXSCAVENGE

FRONT FRAMESUMP SCAVENGER

6 5 4 3 2 1

VENT

SCAVENGE

PRESSURE

SUPPLY

LEGEND

OIL PRESS

PSI

200

150

10

50

BEARING

Figure 7-9. Oil System Schematic




FUEL SYSTEM

GeneralA hydromechanical engine fuel system providesfor automatic fuel metering, variable geometrysystem sequencing and operation, and engineoverspeed control. Automatic compensation isprovided for such variables as ambient pres-sure, temperature, and engine rpm throughout thecomplete operating range of the engine.

Fuel PumpThe fuel pump is a combination centrifugaland high-pressure positive displacement pump.The centrifugal pump supplies fuel under pos-itive pressure to the high-pressure pump and tothe pressurizing and drain valve. The high-pressure positive displacement pump supplieshigh-pressure fuel to the engine fuel controller,the overspeed governor servo, and the jet pumps.

Fuel Controller The fuel controller is mounted on and drivenby the fuel pump assembly and is divided intoa computing section and a metering section.

The computing section senses four parame-ters: compressor inlet temperature, compres-sor discharge pressure, engine rpm, and thrustlever position.

The metering section meters fuel to the com-bustion section to produce or maintain the de-sired thrust . The fuel control unit l imitsmaximum engine rpm to 101.2% while automat-ically maintaining the desired power settingas operating temperatures and pressures change.It regulates fuel flow to simultaneously posi-tion the variable inlet guide vanes and bleedvalves during transient and steady-state oper-ation, and it provides a positive fuel shutoff.

Variable Geometry SystemThe variable geometry system includes actu-ating, operating, and synchronizing linkage toposition the variable inlet guide vanes and theengine compressor bleed valves.

The function of the variable geometry systemis to prevent compressor stalls and surgingwhen operating in a low power range and dur-ing acceleration and deceleration.

The system includes bleed valves at the third,fourth, and fifth stages of the compressor.When these valves are moved toward open, avolume of compressor air is dumped, thus un-loading the compressor and maintaining asafe stall-surge margin between compressorinlet and outlet pressures.

The bleed valves operate together with vari-able guide vanes located in the front frame.The operating linkage is such that when thebleed valves are fully open, the variable guidevanes are at their fully closed position, re-stricting air mass flow to the compressor inlet.This is the position of the bleed valves andguide vanes at low power setting and/or whenthe engine is shut down.

A variable geometry servo in the fuel con-troller directs high-pressure (unmetered) fuelto the actuators for synchronous operation ofthe bleed valves and guide vanes. A feedbackcable supplies bleed valve and guide vane po-si t ion information to the fuel control lerthrough the variable geometry servo.

Overspeed GovernorThe overspeed governor is a hydromechani-cal, self-contained unit placed in series with themain fuel control and is driven by the accessorygearbox. Its purpose is to override the mainfuel control of the engine overspeeds. Under ab-normal operating conditions all fuel passesthrough the governor unhindered. However, ifan overspeed occurs (103.5 = .5% rpm), thegovernor cuts in and maintains a constant en-gine speed by metering the flow through it andbypassing the excess fuel to the main fuel pumpinlet. The overspeed governor has a ground testwhich limits engine rpm to 90.5 + 1% rpm untilplaced back to normal operation.




Fuel Pressurizing and Drain ValveThe pressurizing valve includes a slidingspring-loaded piston and a Teflon seat. A meansis provided to bring centrifugal pump pres-sure to the backside of the piston so that valveopening is a function of fuel pressure plusspring force. Simultaneously, fuel is dischargedto the left and right manifolds. The drain valveis a spring-loaded piston-type valve, which isspring-loaded to the open position and is closedby fuel pressure (60–100 psi).

Fuel NozzlesTwelve fuel nozzles provide the correct spraypattern of metered fuel to the combustion sec-tion for the entire operating range of the en-g ine . Each fue l nozz l e ha s two ma jo rsections—the flow divider and the nozzle(Figure 7-10). The flow divider housings areattached to pads on the mainframe which re-tain the nozzles in position in the combustionsection. This housing has a fuel inlet portwhich is connected to the fuel manifold, and

two outlet ports which supply fuel to the noz-zle section through two tubes—the primary andsecondary fuel nozzles tubes. The air-coolednozzles section contains a primary and second-ary orifice. The initial flow of fuel into the fuelnozzle passes through the divider housing andthe primary fuel tube to the primary nozzle ori-fice where the spray pattern is formed forcombustion.

As fuel pressure increases, fuel is allowed toflow through the secondary fuel tube and noz-zle orifice as well as the primary, forming an-other spray pattern to supplement the fuelrequirements. During engine operation, theprimary flow remains constant, and the sec-ondary spray flow increases to satisfy engineoperating requirements.

Fuel Drain CollectorA collector tank is installed to collect and re-tain the fuel drained from the fuel manifold andflow dividers through the drain section of thepressurizing and drain valve during engineshutdown. The canister volume is sufficient to

PISTONVALVE SPRING

SECONDARY TUBEPRIMARY TUBE

METERING BODY

FLOWDIVIDER

ASSEMBLY

AIR SHROUD ANDSECONDARY ORIFICE

NOZZLE ASSEMBLY

PRIMARY ORIFICE ANDPRIMARY AND SECONDARYDISTRIBUTOR

Figure 7-10. Flow Divider and Nozzle

retain the drainage from three engine shut-downs. Following this, the canister itself mustbe drained; otherwise, it will begin to dumpthe collected fuel overboard.

OperationFigure 7-11 illustrates operation of the fuelsystem and the associated variable geometryand engine overspeed systems.

IGNITION SYSTEMThe ignition system consists of an ignitionexciter and two igniters.

The igniter exciter is a solid-state unit with ex-tended duty capability. In operation, a 28-voltDC input is supplied to the ignition generatorfrom the airframe batteries or generators. Thecharging circuit converts the input voltage tointerrupted or pulsating DC and supplies it tothe primary winding of the transformer. Thetransformer steps up the 28-volt pulsating DCinput to 3,000 volts.

The igniter plugs are self-ionizing, shuntedgap-type igniters designed for relatively lowvoltage application.

ControlAutomatic ignition occurs during engine start-ing when the GEN switch (Figure 7-12) ismoved to START, and the associated thrustlever is moved from CUT OFF to IDLE.Ignition is terminated when the GEN switchis moved to OFF or GEN position.

Selective ignition provides for continuousoperation of the ignition system when thetwo-position AIR IGN switch on the centerswitch panel (Figure 7-12) is moved to theON position.

AIR SYSTEMSThe engine air systems consist of a primary airsystem and the engine bleed-air system.

The primary air system (see Figure 7-1) isthat portion of the engine air system whichis associated partly with combustion andpartly with the direct cooling. It is describedas follows:

The combustion air system is that portion ofthe primary air system which is directly asso-ciated with combustion and thrust. The vari-able inlet guide vanes in the front frameannulus guide the air into the compressor foroptimum airflow. The air passes through thecompressor and is diffused in the mainframe.It is then mixed with fuel in the combustionsection and ignited. The combustion gasesthen pass through the turbine where the greaterportion of available energy is used to drive theengine compressor. The remainder passesthrough the exhaust core and tailpipe to pro-vide thrust. The flow of primary cooling airaround the inner and outer surfaces of thecombustion liner cools the liner and the tur-bine area. The cooling air then combines withthe combustion airstream.

The pressurizing system provides air throughpassages in the engine components to pressur-ize the oil seals and the engine lube system.

Compressor discharge leakage air flows acrossthe compressor eighth-stage air seal and pres-surizes the No. 2 bearing carbon seal. Part of thisair circulates through the compressor rotor in-terior to pressurize the No. 1 bearing carbon seal.

Part of the combustion section cooling airflows into the balance piston chamber through




Figure 7-12. Center Switch Panel




Fig

ure

7-1

1.F

uel

Sys

tem

Sch

emat

ic

VA

RIA

BLE

GU

IDE

VA

NE

AC

TU

ATO

RS

FLO

W D

IVID

ER

/FU

EL

NO

ZZ

LES

(12

)

RH

MA

NIF

OLD

OIL

IN

TO E

PA C

OLL

EC

TOR

OIL

OU

T

SE

NS

E L

INE

FUEL

FIL

TER

BY

PAS

S

FR

OM

TA

NK

FILT

ER

FU

EL

SH

UTO

FF

VA

LVE

MO

TIV

E-F

LOW

SH

UTO

FF

VA

LVE

LH M

AN

IFO

LD

BO

OS

TE

LEM

GE

AR

ELE

MM

AIN

FU

EL

CO

NT

RO

LO

VE

RS

PE

ED

GO

VE

RN

OR

FLO

W-

ME

TE

R

FIL

TE

R

PR

ES

SU

RIZ

ING

AN

DD

RA

IN V

ALV

E

OIL

CO

OLE

R DR

AIN

FU

EL

ELE

CT

RIC

AL

HP

FU

EL

PR

ES

SU

RE

RE

GU

LAT

ED

PR

ES

SU

RE

SU

PP

LY F

UE

L

LE

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ND




holes in the aft flange of the inner combustioncasing and mating flanges. The air then flowsthrough the turbine wheel inner labyrinth sealto pressurize the No. 3 bearing carbon seal.

The sumps are pressurized by seal leakage airthat enters the sumps.

Vent lines and check valves between the en-gine sumps, gearboxes, and oil tank maintaina controlled positive pressure between thesecomponents.

The engine bleed-air system’s air is bled offto control airflow and to supply air for heat-ing, cooling, ventilation, and pressurizationpurposes in the engine and airframe. The bleedsystem functions as follows:

Part of the air bled from the compressor at theeighth stage is bled into the front frame foranti-icing. After the air heats the front framecasing, struts, variable vanes, and the bullet,it is discharged into the engine airstream.

Air is bled from the engine inlet duct into anaspirator hose. It then flows to a bellows in themain fuel control which senses the tempera-ture of this inlet air for fuel scheduling.

Part of the compressor eighth-stage dischargeair is bled off to pressurize the carbon oilseals. Part of the combustion cooling air flowsinto the turbine area through holes in the aftflange and mating flanges of the inner com-bustion casing where it cools the turbine noz-zle and opposes forward rotor thrust bypressurizing the balance piston chamber.

The air bled from the compressor at the third,fourth, and fifth stages is metered overboardas scheduled by the fuel controller to maintainair- flow stability within the compressor andto prevent compressor stalls.

Some of the compressor discharge air is bledoff through two main frame ports for pressur-ization and heating of the airplane cabin,windshield defog, wing heat, pressurizationjet pump, and, on 24-297 and subsequent and

25-135, -181 and subsequent, for hydraulicreservoir pressurization. On 24-350, -352 andsubsequent and 25-227 and subsequent, it isalso used for windshield-radome alcohol andcabin temperature and pressurization control.

The sump center vent system is designed tomaintain a controlled positive pressure in thefront bearing sump, main sump, gearcase, and oiltank. Maintaining a positive pressure makes thelube system sensitive to altitude changes.Controlled carbon seal leakage for the system isthe source of pressurizing air.

ENGINE POWER CONTROL

GENERALThe engine power control consists of a thrustlever located on the center console (Figure7-13). The thrust lever operates in a quadrantfrom the ful l af t , or CUT OFF, posi t ionthrough an IDLE position to a full forward,or maximum thrust, position. The thrust leverslot has a gate at the IDLE position to pre-vent the inadvertent selection of CUT OFF.The lever must be pulled left and moved aftto select CUT OFF.

Airplanes fitted with thrust reversers have re-verse levers (Figure 7-13) piggyback-mountedon the thrust levers and have a release at therear base of the thrust lever that must be raisedto allow the lever to be moved to the CUTOFF position.

FRICTION CONTROLA friction control twist knob located on the rightside of the center console (Figure 7-13) is used toset friction for both thrust levers. Forward orclockwise rotation increases friction.




ENGINE INSTRUMENTATION

GENERALThe engine instruments are mounted in twovertical rows on the center instrument panel(see Figure 7-8). From top to bottom, these are:

• Pressure ratio

• Exhaust temperature

• Engine rpm

• Fuel flow

• Oil pressure

• Oil temperature

PRESSURE RATIOEngine pressure ratio, or EPR, is an indicationof the ratio of engine inlet to exhaust gas pres-sure and provides an indication of the amountof thrust being produced by the engine.

The EPR system includes two probes in the ex-haust system, four transducers (two for eachengine) on 23 models and 24 models through24-180, and 25-003 through -024, and a probein the ram-air duct. On all other models, theinlet air pressure is supplied from the copilot’spitot tube, and there is only one transducer perengine. The inlet and exhaust pressure sig-nals are sent to the appropriate EPR trans-ducer where they are converted to electricalsignals, amplified, and sent to the EPR gageson the center instrument panel.

The EPR system on models 23-003 through 24-180 and 25-003 through 25-024 requires 26volts AC. All other models require 115 volts AC.

EXHAUST TEMPERATUREExhaust temperature is sensed by eight thermo-couples in the exhaust system and sent to a cock-pit gage calibrated in degrees Celsius. The sys-tem is self-generating and does not require anelectrical power supply.

Figure 7-13. Center Console

ENGINE RPMEngine rpm is supplied from an engine-driventach generator mounted on the rear of each en-gine-driven oil pump and driven by a com-mon shaft. Dial markings are based on percentof maximum allowable engine speed. The largedial is graduated in 2% increments from 0 to100%; the small dial is graduated in 1% incre-ments from 0 to 10%. Engine rpm of 100%equa l s approx ima te ly 16 ,500 rpm. Thetachometer system uses self-generated elec-trical power from the tach generator.

FUEL FLOWFuel flow is measured by a transmitter on themetered fuel line to the combustion section.Fuel flow is indicated on a single-needle gageand uses DC power. A two-position L–R switchlocated below the gage provides for selectionof fuel flow indication for the left or right en-gine. Models 24-354 and subsequent and 25-231and subsequent have dual-needle gages installed.

OIL PRESSUREThe oil pressure indicating system consists ofan oil pressure transmitter and a pressure switchinstalled on the engine nacelle bulkhead, anda dual needle oil pressure gage.

One red low oil pressure warning light (see“Annunciator Panel” section) is installed on theglareshield. If the oil pressure drops below 5 psi,the red low oil pressure light will illuminate. Thelight uses 28 VDC, and the gage uses 26 VAC.

NOTEThere is one light for both engines;therefore, if only one engine is op-erating, the red low oil pressure lightwill be illuminated.

OIL TEMPERATURETwo oil temperature indicators powered by 28VDC are installed in the instrument cluster on theswitch panel, and a probe is installed in the bot-tom of each oil tank.

ENGINE STARTINGMoving the starter-generator switch to the STARTposition applies power to close the respectivemotive flow control valve and energize the standbyfuel pump. With the motive flow control valveclosed, the starter relay is energized to apply powerto the starter. At 10% engine rpm, the thrust lever ismoved to IDLE, energizing the ignition and sup-plying fuel to the fuel nozzles. When the enginereaches an idle rpm of 48%, the starter-generatorswitch is moved to GEN. This deenergizes thestarter relay, opens the motive flow control valve,turns off the standby pump and ignition, and ener-gizes the generator relays. The generator fail lightextinguishes, and the generator supplies power tothe airplane electrical system.

ENGINE SYNCHRONIZING

WOODWARD ELECTRONICSYNCHRONIZER (OPTIONAL)The engine synchronizing system consists ofa magnetic pickup in each engine overspeedgovernor, a control box, throttle cable actu-ator, control switch (Figure 7-14), and indi-cator light (see “Annunciator Panel” section).




Figure 7-14. ENG SYNC Switch




The system should be off below 70% rpm.The indicator light will light if the system isinadvertently left on with the landing geardown. The control box senses pulse differ-ence from the pickups, then energizes thethrottle cable actuator on the right (slave) en-gine. The right engine rpm is matched to theleft (master) engine. The actuator travel islimited; hence, the engines must be nearlysynchronized manually before setting thecontrol switch to ON to obtain maximum±slave engine adjustment in synchroniza-tion. The system operates on 28 VDC sup-pl ied through a 7 .5-amp circui t breakerlabeled “ENG SYN” on the copilot’s circuit-breaker panel.

SYNCHROSCOPE (OPTIONAL)Manual engine synchronization can be moni-tored with the dual-engine synchroscope. Thesynchroscope is basically a synchronous motordriven by outputs from the tach generators andis otherwise independent of the airplane electri-cal system. A wheel marked by black and whitesections moves clockwise or counterclockwiseat a rate proportional to the difference in fre-quency of the tach generators. Clockwise rota-tion is for the right tach fast, and counterclock-wise is for the left tach fast. Synchronization isaccomplished by adjusting either engine thrustlever until the wheel stops rotation.

THRUST REVERSERS(OPTIONAL)

GENERALThe thrust reversers (Figure 7-15) are an ad-ditional deceleration system which may beused anytime the airplane is on the groundto produce shorter stopping distances. Theycannot be used to supersede runway lengthrequirements published in the performancechapter of the AFM.

Each engine is equipped with a target thrustreverser (T/R) consisting of upper and lowerclamshell doors (buckets), pivoted near theengine centerline. The reverser doors are hy-draulically actuated, and electrically con-trolled mechanical latches are provided foreach engine to secure the doors in the stowedposition. Hydraulic pressure for the system issupplied by the airplane hydraulic systemthrough a one-way check valve. The T/R sys-tem includes an accumulator. When fullycharged, the accumulator hydraulically pro-vides several cycles should the airplane hy-draulic system fail upstream of the one-waycheck valve. The accumulator is serviced to650 psi with dry air or nitrogen (hydraulicpressure zero).

Figure 7-15. Thrust Reversers




CONTROL AND INDICATIONThe thrust reverser control switches and an-nunciator lights are located in a subpanelmounted either above the glareshield lightpanel or in the center instrument panel (Figure7-16). Switches are provided to arm the sys-tems for normal operation, emergency stow-ing, and for the test functions. Annunciationlights are provided to indicate left and rightsystem armed, unsafe, and deployed. With thecontrol switches in NORMAL and ARMED,the deployment or stowing of the thrust re-versers and engine rpm in reverse is controlledby piggyback subthrottles (see Figure 7-13)mounted on the engine thrust levers. The sub-throttles can be raised only with the enginethrust levers in the IDLE position.

Raising the subthrottles to the IDLE detentposition deploys the reversers with no in-crease in engine rpm. A mechanical lockoutprevents engine acceleration in reverse untilthe appropriate doors are deployed. A groundadjustable stop in the thrust levers limits en-gine rpm in reverse to a maximum of 85%. Asystem of mechanical and electrical inter-locks is provided to prevent unwanted or in-advertent deployment. With the engine thrustlevers in forward thrust above idle, the T/Rsystem cannot be armed. Electrical inter-locks are provided through landing gear squatswitches to prevent in-flight deployment.

AUTOMATIC STOWAn automatic stow system is also incorpo-rated into the system. In flight, a stow com-mand is automatically introduced whenever theT/R is in any position except fully stowed.

EMERGENCY STOWAn emergency stow switch is mounted on thecontrol switch panel (Figure 7-16). Actuationof this switch bypasses all interlocks and allother signals with a direct stow command ifthe ARM switches are off. Proper functioningof the normal deploy/stow and emergencystow systems is accomplished during systemcheckout. If a malfunction is indicated, flightwith the system disabled and safety pins in-stalled is permitted.

NOTEWith the ARM switches on, the emer-gency stow function is inoperative.With switches to EMERG STOW, thearm function of the system is inoper-at ive . For l imita t ions , pref l ightchecks, normal and emergency proce-dures, consult the AFM.

Figure 7-16. Thrust Reverser Control, Test, and Indicating Panel (Typical)

DIFFERENCESThe two CJ610 engine configurations varygenerally in the gearbox and accessory area.

On the CJ610-1, power to drive the engine-mounted accessories is extracted from the en-gine through the power takeoff assembly. Adriver shaftgear is coupled to the aft end of thecompressor drive shaft by means of a splineand contains six holes for venting the mainsump. The forward end of the driver shaftgearis a bevel gear which engages with another bevelgear mounted in the power takeoff housing per-pendicular to the engine centerline. A radialshaft, which is splined to the driver shaftgear,transmits the power radially out of the enginethrough the six o’clock mainframe strut. The out-board end of this radial drive shaft is splined tothe transfer gearbox in the CJ610-1, and di-rectly to the accessory gearbox in the CJ610-4,-6 and -8 engines.

GE CJ610-1 and -4 engines are rated at 2,850pounds of thrust. CJ610-6 and -8A enginesare rated at 2,950 pounds.




1. The engines powering the 20 seriesLearjets are defined as:A. Single-spool, reverse flow turbojetsB. Twin-spool, straight flow turbojetsC. Single-spool, axial-flow turbojetsD. Twin-spool, centrifugal turbojets

2. The function of the variable geometrysystem is to:A. Prevent compressor stalls and surgesB. Offload the turbines at high altitudesC. Provide laminar airflow through the

compressorD. Limit combustion chamber pressure

3. The power source required for engineoil pressure gages is supplied by:A. A battery-operated power packB. The pilot’s 115-volt AC busC. Resistance bulbs at the oil pump outletD. The 26-VAC system

4. The primary thrust indicator for theCJ610 engine is:A. EGTB. Fuel flowC. Engine pressure ratio (EPR)D. Rpm

5. During start, ignition occurs when the:A. Rpm indicates 10%B. Thrust lever is moved from the CUT-

OFF positionC. Start GEN switch is moved to the

START positionD. Generator magnetic pickups sense 8%

rpm

6. If fuel flow, EGT, and EPR are normalfor the thrust setting but oil pressure andrpm are indicating zero, the probablecause is:A. A sheared oil pump shaftB. Loss of the electrical power sourceC. An open circuit breakerD. A faulty oil pressure switch




QUESTIONS

8-i

CHAPTER 8FIRE PROTECTION

CONTENTS

Page

INTRODUCTION ................................................................................................................... 8-1

GENERAL............................................................................................................................... 8-1

ENGINE FIRE DETECTION AND INDICATORS............................................................... 8-1

Sensing Elements and Control Units ............................................................................... 8-1

Fire Lights ........................................................................................................................ 8-2

Fire Detection System Test .............................................................................................. 8-2

Exterior Extinguisher Discharge Indicators ..................................................................... 8-3

ENGINE FIRE EXTINGUISHING ........................................................................................ 8-3

Extinguisher Containers................................................................................................... 8-3

Operation.......................................................................................................................... 8-3

QUESTIONS ........................................................................................................................... 8-6




8-iii

ILLUSTRATIONS

Figure Title Page

8-1 FIRE Warning Lights ............................................................................................... 8-2

8-2 FIRE DET Test Switch............................................................................................. 8-3

8-3 Fire Extinguisher Discharge Indicators .................................................................... 8-3

8-4 Engine Fire-Extinguishing System (Airplanes withGlareshield Warning Lights) .................................................................................... 8-4

8-5 Engine Fire-Extinguishing System (Airplanes withoutGlareshield Warning Lights) .................................................................................... 8-5




FIRE PULL

FIREWARN




INTRODUCTIONThe Learjet 20 series airplanes are equipped with engine fire detection and fire-extinguishingsystems as standard equipment (except model 23 airplanes, which have fire detection only).The systems include detection circuits which give visual warning in the cockpit and controlsto activate one or both fire extinguisher bottles. A test function for the fire detection systemis incorporated.

GENERALFire detection consists of separate and inde-pendent systems for the left and right engines,separate indicating (fire warning) systems,and a fire detection test system.

A two-bottle fire-extinguishing system is com-mon to both engines. The system is such thateither of two bottles of extinguishing agent canbe discharged to either engine, or both bottlescan be discharged to the same engine.

ENGINE FIREDETECTION ANDINDICATORSSENSING ELEMENTS ANDCONTROL UNITSWithin each engine cowling are two heat-sensing elements—one installed around theforward nacelle and one surrounding the en-gine tailcone. Each assembly is connected to

CHAPTER 8FIRE PROTECTION

a control unit which monitors the electrical re-sistance of the sensing elements. The sensingelements are made of Inconel metal tubingfilled with a pliable, heat-sensitive ceramic ma-terial which, in turn, encloses a conductorwire at its center. The electrical resistance ofthe ceramic material is relatively high at nor-mal temperatures. At high temperatures, how-ever, the electrical resistance decreases andallows increased current flow.

The control unit detects the increased currentflow and energizes the red FIRE warning lighton the glareshield (or on the readout panel onairplanes without glareshield warning lights)when temperature reaches 510ºF in the rearnacelle area, or 480ºF in the forward nacellearea. A system test switch is installed forwardof the throttle quadrant to check the continu-ity of the sensing elements and control units.The control units are located in the tailcone ex-cept for airplanes 23-003 through 24-129. Onthese airplanes, they are located in the nosecompartment.

DC electrical power for the system is suppliedthrough circuit breakers on the appropriatepilot’s and/or copilot’s circuit-breaker panels.

FIRE LIGHTSThe red FIRE warning lights are located ateach end of the glareshield (Figure 8-1) or onthe r eadou t pane l o f a i rp l anes w i thou tglareshield warning lights. In the event of anengine fire, the applicable light flashes untilthe overheat condition ceases to exist. On air-planes without glareshield lights, the FIRElight does not flash.

FIRE DETECTION SYSTEM TESTThe FIRE DET test switch on the forward endof the throttle quadrant (Figure 8-2) is used totest continuity of the sensing elements andcontrol units. On airplanes with glareshieldlights, positioning the switch to TEST causesthe FIRE lights to flash. On airplanes withoutglareshield lights, the FIRE lights remain on




Figure 8-1. FIRE Warning Lights

L FIRE R FIRE

AIRPLANES WITHOUT GLARESHIELD WARNING LIGHTS

AIRPLANES WITHGLARESHIELD WARNING LIGHTS

steady. Returning the switch to NORM extin-guishes the lights.

EXTERIOR EXTINGUISHERDISCHARGE INDICATORSTwo colored disc indicators are flush-mountedin the side of the fuselage below the left en-gine pylon (Figure 8-3). The red disc coversthe thermal discharge port. It will be rupturedif one or both thermal relief valves have re-leased bottle pressure. The yellow disc will beruptured if either bottle is discharged by de-pressing an illuminated ARMED light. The

integrity of the two discs is checked during theexternal preflight inspection.

ENGINE FIREEXTINGUISHING

EXTINGUISHER CONTAINERSTwo spherical extinguishing agent contain-ers are located in the tailcone area. Both con-tainers use common plumbing to both enginecowlings via shuttle valves, providing the air-plane with a two-shot system. The chemicalagent in the containers, monobromotrifluo-romethane (CF3Br), is not corrosive, and itsdischarge will not necessitate cleaning of theengine and cowl areas.

The containers are plumbed to each enginecowling to provide the airplane with a two-shotsystem. The containers will fully discharge inone to two seconds. A pressure gage mountedon each container indicates the charge of thecontainer. A thermal relief valve on each con-tainer is plumbed to a thermal (red) dischargeindicator to relieve bottle pressure when bot-tle temperature reaches 217ºF.

Access to the fire-extinguishing container isthrough the tailcone door.

OPERATIONWhen a FIRE light illuminates (or flashes), itindicates a fire or overheat condition in the re-spective engine cowling. The pilot should firstplace the thrust lever of the affected engine toCUT OFF, then proceed with the AFM proce-dures. A brief description of system operationis as follows:

Airplanes With GlareshieldWarning LightsLifting the guard and depressing the red FIREwarning light applies 28 VDC to close the hy-draulic and fuel shutoff valves, and on airplanesSNs 24-350 and subsequent and 25-227 andsubsequent, it closes the associated bleed-air




Figure 8-2. FIRE DET Test Switch

Figure 8-3. Fire Extinguisher DischargeIndicators

valve as well. A small, red shutoff valve light,commonly called a “pinhead light,” will illu-minate to indicate the FIRE warning light hasbeen depressed. Power is also applied to arm therespective amber ARMED lights (Figure 8-4).

Depressing an illuminated ARMED light mo-mentarily applies DC power to an explosivecartridge, which discharges the contents of abottle to the selected engine. Depressing theadjacent ARMED light fires the remainingbottle to the same engine. Depressing the FIREwarning light a second time disarms the sys-tem (ARMED lights go out), opens the hy-d rau l i c and fue l shu to ff va lve s , andextinguishes the pinhead light.

Airplanes Without GlareshieldWarning LightsOpening the red guard and actuating a FIRE-WALL SHUTOFF switch (LH or RH) selectsthe engine to which fire-extinguishing agentis to be applied. It also applies 28 VDC toclose the hydraulic and fuel shutoff valves(Figure 8-5) and arms the No. 1 and No. 2 dis-charge switches.

Depressing the No. 1 discharge switch firesa bottle to the selected engine. If a secondcharge of extinguishing agent is desired, de-pressing the No. 2 discharge switch fires ther e m a i n i n g b o t t l e t o t h e s a m e e n g i n e .Returning the guarded FIREWALL SHUT-OFF switch to the off position opens the hy-draulic and fuel shutoff valves and disarmsthe fire-extinguishing system.




SHUTOFF VALVE (PINHEAD) LIGHT

THERMALDISCHARGE INDICATOR

MANUALDISCHARGE INDICATOR

ARMED

FUEL SHUTOFF VALVE

HYDRAULICSHUTOFF VALVE

ARMED FIRE

LH

CONTAINER

RELIEF VALVE RELIEF VALVE

PRESSURE GAGEPRESSURE GAGE

RH NACELLELH NACELLE

ARMED

FUEL SHUTOFF VALVE*BLEED-

AIR VALVE*BLEED-

AIR VALVE


ARMEDFIRE

RH

CONTAINER

TWO-WAYCHECKVALVES

*SNS 24-350 AND SUBSEQUENT AND 25-227 AND SUBSEQUENT

ENGINE EXTINGUISHING

MANUAL DISCHARGE

THERMAL DISCHARGE

LEGEND

Figure 8-4. Engine Fire-Extinguishing System (Airplanes with Glareshield Warning Lights)




THERMALDISCHARGEINDICATOR

MANUALDISCHARGEINDICATOR

FUEL SHUTOFF VALVE

HYDRAULICSHUTOFF VALVE RH NACELLELH NACELLE

FUEL SHUTOFF VALVE


ENGINE EXTINGUISHING

MANUAL DISCHARGE

THERMAL DISCHARGE

LEGEND

LH RH

FIREWALL

SHUTOFF

FIREWALL

SHUTOFF

NO. 1 NO. 2

DISCHARGESWITCHES

LH

CONTAINER

RELIEF VALVE RELIEF VALVE

PRESSURE GAGEPRESSURE GAGE

RH

CONTAINER

TWO-WAYCHECKVALVES

Figure 8-5. Engine Fire-Extinguishing System (Airplanes withoutGlareshield Warning Lights)

1. Engine fire detection and/or overheat issensed by:A. ThermocouplesB. ThermostatsC. Continuous-wire sensing elementsD. Thermistors

2. The fire-extinguishing container chargeis indicated by:A. A red disc on the rear fuselageB. A gage on each containerC. Armed lightsD. A yellow disc on the rear fuselage

3. Engine fire extinguisher containers arelocated in:A. The nacellesB. The engine pylonsC. The tailconeD. The baggage compartment

4. If the fire persists after activating a fire bottle:A. A second fire bottle can be dis-

charged into the affected area.B. The second fire bottle can only be

used on an opposite-side fire.C. The first fire bottle can be dis-

charged a second time.D. No further activation of the system

is possible; both bottles dischargesimultaneously when eitherARMED button is pressed.

5. Airplanes with glareshield warninglights—When the left FIRE light isdepressed:A. It discharges one extinguisher into

the left nacelle.B. It closes the main fuel and hydraulic

shutoff valves for the left engineand arms the left ARMED lights.

C. It discharges one extinguisher andarms the second.

D. It ruptures the yellow dischargeindicator disc.

6. Airplanes without glareshield warninglights—When the left FIREWALLSHUTOFF switch is actuated:A. It closes the main fuel and hydraulic

shutoff valves for the left engineand arms the No. 1 and No. 2 dis-charge switches.

B. It discharges one extinguisher intothe left nacelle.

C. It discharges one extinguisher andarms the second.

D. It ruptures the yellow dischargeindicator disc.




QUESTIONS

9-i

CHAPTER 9PNEUMATICS

CONTENTS

Page

INTRODUCTION ................................................................................................................... 9-1

GENERAL............................................................................................................................... 9-1

DESCRIPTION AND OPERATION ...................................................................................... 9-2

SNs 23-003 through 24-349 and 25-003 through 25-226 ................................................ 9-2

SNs 24-350 and Subsequent and 25-227 and Subsequent ............................................... 9-2

QUESTIONS ........................................................................................................................... 9-6




9-iii

ILLUSTRATIONS

Figure Title Page

9-1 Pneumatic System—SNs 23-003 through 24-349 and 25-003 through 25-226 ...................................................................................... 9-3

9-2 Pneumatic System—SNs 24-350 and Subsequentand 25-227 and Subsequent ..................................................................................... 9-4

9-3 AIR BLEED Switches............................................................................................... 9-5




Bleed air from the engine compressor section isdelivered to a bleed-air manifold in theunpressurized aft compartment (tailcone). Theair from this bleed-air manifold is used forcabin pressurization and heating, wing anti-icing, windshield anti-icing, pressurizationcontrol (except early 23, 24, and 25 models),and hydraulic reservoir pressurization on SNs24-297 and subsequent, 25-135, and 25-181and subsequent.

SNs 24-350 and subsequent and 25-227 andsubsequent also use servo bleed air fromradome/windshield alcohol anti-icing, operationof the emergency pressurization valves, andcabin pressurization and temperature control.Control of bleed air is accomplished with the leftand right AIR BLEED switches located on thecopilot’s lower instrument panel. Visualindications of bleed-air duct overtemperaturesare given by illumination of warning lightson the glareshield annunciator panel (see“Annunciator Panel” section).

VALVE

L R

COBLEED AIR

515

20

AIR

CHAPTER 9PNEUMATICS




INTRODUCTIONThe Lear 20 series pneumatic system uses bleed air extracted from the engine compressorsection. Bleed air is used for cabin pressurization and heating and anti-icing systems on allairplanes, and servo air is used for hydraulic reservoir pressurization, pressurization control,temperature control, and radome/windshield alcohol anti-icing on later models. Compressor bleedair is also used for engine front frame anti-icing. There are two basic pneumatic systemconfigurations—one for SNs 23-003 through 24-349, and 25-003 through 25-226, and one forSNs 24-350 and subsequent, and 25-227 and subsequent. The later system incorporates a majordesign change including the installation of emergency pressurization valves.

GENERAL

DESCRIPTION ANDOPERATION

SNs 23-003 THROUGH 24-349AND 25-003 THROUGH 25-226

Bleed-Air Check ValvesA check valve is installed in the bleed-air ductingfrom each engine. Each check valve allowsairflow in one direction and blocks airflowapplied in the opposite direction. The checkvalves prevent loss of bleed air during single-engine operation (Figure 9-1).

NOTEOn SNs 23-003 through 23-029,unless modified by SK140, anadditional bleed-air shutoff valve isinstalled in the engine bleed-airducting between the engine and thebleed-air check valves discussedabove. The bleed-air shutoff valves areopened and closed by their respectiveAIR BLEED switch located on thecopilot instrument panel (see note onFigure 9-1).

Bleed-Air ManifoldThe bleed-air manifold serves as a collection pointfor engine bleed air from either or both engines.From the manifold, bleed air is distributed to theflow control valve for cabin pressurization andheating, pressurization jet pump (except early 23,24, and 25 models), windshield heat, wing heat,and hydraulic reservoir pressurization on SNs 24-297 and subsequent and 25-135, 25-181 andsubsequent (Figure 9-1).

Flow Control ValveThe flow control valve regulates the flow of bleedair to the cabin for normal pressurization andheating. On SNs 23-003 through 24-229 and 25-003 through 25-064, excluding 25-061, the flow

control valve is controlled by an ON–OFF AIRBLEED switch and a NORM–MAX CABINAIRFLOW switch. The AIR BLEED switchopens or closes the valve while the CABINAIRFLOW switch selects either the normal ormaximum airflow rate. On SNs 24-230 through24-349, 25-061 and 25-070, through 25-226, it iscontrolled by a three-position AIR BLEED switchlabeled “OFF,” “NORM,” and “MAX,” on thepressurization module. The OFF position closesthe valve completely. MAX position opens thevalve to full flow. NORM position allows senseline pressure from the venturi (locateddownstream) to modulate the flow control valveto maintain a constant airflow to the cabin.

SNs 24-350 AND SUBSEQUENTAND 25-227 AND SUBSEQUENT

LH and RH AIR BLEED SwitchesThe left and right bleed-air switches on thecopilot’s lower instrument panel control theirrespective bleed-air shutoff valves (see Figure 9-2).There are three-position ON–OFF–EMERswitches that draw DC power from the left or rightECS valve circuit breaker on the respective left orright main bus. In the ON position, the bleed-airshutoff valve is open. In the OFF position, thevalve is closed. In the EMER position, the bleed-air shutoff valve remains open, and electricalpower is removed from the respective emergencypressurization valve, causing it to route bleed airdirectly to the cabin.

LH and RH EmergencyPressurization ValvesThe emergency pressurization valves, one foreach engine (see Figure 9-2), are installed in theengine bleed-air line between the bleed-airshutoff valves and the bleed-air manifold in thetailcone. The valve has two positions—normaland emergency. In the normal position, bleed airis directed to the bleed-air manifold; in theemergency position, the bleed air is routeddirectly to the cabin. The valve is spring-loadedto the emergency position and requires electricalpower and servo air pressure to move it to the







LH ENGINEBLEED AIR

ENGINEFRONTFRAMEHEAT

WING ANTI-ICE

FLOWCONTROL VALVE

TO CABIN

BLEED-AIRSHUTOFFVALVE

ENGINEFRONTFRAMEHT

RH ENGINEBLEED AIR

RHENGINE

ON

OFF

AIRBLEEDHYDRAULIC

RESERVOIR PRESSURE ✽✽

✽

✽✽

BLEED-AIR SHUTOFF VALVES ARE INSTALLED ON 23-003 THROUGH 23-029 UNLESS MODIFIED BY SK140.

SN 24-297 THROUGH -349 AND 25-135 AND -181THROUGH -226

✽

(IF INSTALLED) PRESSURIZATIONJET PUMP

WINDSHIELD HEATANTI-ICE

BLEED-AIRSHUTOFF VALVE

ON

LH ENGINE

OFF

AIRBLEED

✽

REGULATED BLEED AIR

BLEED AIR

LEGEND

Figure 9-1. Pneumatic System—SNs 23-003 through 24-349 and 25-003 through 25-226







SERVOAIR

MANIFOLD

BLEED-AIRSHUTOFF

VALVE

PRESSJET PUMP

PRESSURIZATIONANEROID SWITCH9,500 FT (± 250 FT)

PRESSURIZATIONANEROID SWITCH9,500 FT (± 250 FT)

EMERGPRESSVALVE

LH AIR BLEED

EMER

ON

OFF

OVERTEMPERATURE SWITCH

LH ECS VALVE

TO CABIN

WINDSHIELD HEAT

WING HEAT

OFF

ON

EMER

OVERTEMPERATURE SWITCH

SE

RV

O P

RE

SS

RH ECS VALVE

TO CABIN

L CAB AIR

R CAB AIR

CONDITIONED AIR

RAM AIR

REGULATED SERVO AIR

BLEED AIR

LEGEND

H-VALVE

REGULATOR

RAM AIR OUT

HEATEXCHANGER

TEMPCONTROL

ALCOHOLANTI-ICE

FLOWCONTROL

VALVE

CABIN

RAM AIR IN

EMERGPRESSVALVE

BLEED-AIRMANIFOLD

HYDRESERVOIR

PRESS

PRESSUREAND HEAT

Figure 9-2. Pneumatic System—SNs 24-350 and Subsequent and 25-227 and Subsequent

normal position. The loss of either electricalpower or servo air automatically positions thevalve to the emergency position. Below aresituations where this occurs:

A. Both engines shut down (loss of servopressure).

B. Cabin altitude is 9,500 feet (aneroid switchopens circuit to the solenoid valve)

C. AIR BLEED switch to EMER (openscircuit to the solenoid valve)

D. Complete DC power failure (no poweravailable to solenoid valve)

E. Servo regulator failure (no servopressure available)

Bleed-Air Check Valves (6)Three bleed-air check valves (Figure 9-2) areinstalled in the bleed-air ducting from each engine. Each check valve allows airflow inone direction and blocks airflow in the oppositedirection. These check valves prevent loss ofbleed air during single-engine operation oractuation of the emergency pressurization valves.

Bleed-Air ManifoldThe bleed-air manifold serves as a collectionpoint for engine bleed air from either or bothengines. From the manifold, bleed air isdistributed to the flow control valve for cabinpressurization and heating, windshield anti-ice(defog) valve, wing anti-ice, and the hydraulicreservoir pressure regulator.

Flow Control ValveThe flow control valve regulates the flow ofbleed air to the cabin for normal pressurizationand heating. It is controlled by a two-positionCAB AIR switch on the pressurization module.The OFF position closes the valve completely.The ON position allows sense line pressure fromthe venturi (located downstream) to modulatethe flow control valve to maintain a constantairflow to the cabin. With loss of DC electricalpower, the valve fails to the normal ON position.

Cabin-Air Warning LightsThe red L and R CAB AIR warning lights on theglareshield annunciator panel illuminate whenan associated duct temperature sensor detectsexcessive temperatures.

Servo Air ManifoldBleed air is tapped prior to the bleed-air shutoffvalves and is routed through a check valve to theservo air manifold (Figure 9-2). From thismanifold, air is ducted directly to the alcohol anti-icing system and through two regulators. The airfrom one regulator is used for cabin temperaturecontrol. Air from the other regulator suppliesservo air pressure to operate the flow controlvalve, emergency pressurization valves, and thepressurization vacuum regulator (jet pump). SeeFigure 9-3 for AIR BLEED switch locations.




CABIN AIR FLOWINSTRUMENT PANEL

LOCATORAIR BLEED

EMERONOFF

ON

OFF

CABAIR

MAXNORMOFF

23-003 THROUGH 24-22925-003 THROUGH 25-064

24-230 THROUGH 24-34925-070 THROUGH 25-226

24-350 AND SUBSEQUENT25-227 AND SUBSEQUENT

NORM MAX

ON

OFF

AIR

BLEED AIR

BLEED

Figure 9-3. AIR BLEED Switches




QUESTIONS

1. Bleed air is used for:A. Pressurization/heatingB. Windshield defogC. Wing anti-iceD. All the above

2. With loss of DC electrical power, theflow control valve:A. Fails closedB. Fails to normal/onC. Remains in the last positionD. Moves to the fully open position

3. In the event of single-engine operation,the check valves in the bleed-air linefrom each engine:A. Must be electrically closedB. Allow negative airflowC. Prevent airflow into the inoperative

engineD. Shut off the pneumatic system

4. For normal cabin pressurization tooccur after takeoff, the pneumaticsystem bleed-air (AIR BLEED/CABAIR) switches:A. Must be inactiveB. Should be in emergencyC. Should be in defogD. Must be turned on

10-iFOR TRAINING PURPOSES ONLY



CHAPTER 10ICE AND RAIN PROTECTION

CONTENTS

Page

INTRODUCTION ................................................................................................................. 10-1

GENERAL ............................................................................................................................ 10-1

ICE DETECTION ................................................................................................................. 10-3

Windshield Ice Detection............................................................................................... 10-3

Wing Ice Detection ........................................................................................................ 10-3

ANTI-ICE SYSTEMS........................................................................................................... 10-3

Engine Anti-ice System (Nacelle Heat)......................................................................... 10-3

Windshield Anti-ice/Defog and Rain Removal System................................................. 10-5

Windshield/RADOME Alcohol Anti-ice System.......................................................... 10-8

Wing and Horizontal Stabilizer Anti-ice System ........................................................ 10-13

Pitot, Static, and Angle-of-Attack Vane Anti-ice System ........................................... 10-14

QUESTIONS....................................................................................................................... 10-15

10-iiiFOR TRAINING PURPOSES ONLY



ILLUSTRATIONS

Figure Title Page

10-1 Ice and Rain Protected Areas ............................................................................... 10-2

10-2 Ice Detection Lights ............................................................................................. 10-3

10-3 Recognition Light................................................................................................. 10-3

10-4 Engine Anti-ice System........................................................................................ 10-4

10-5 L and R NAC HEAT Switches ............................................................................. 10-5

10-6 Windshield Anti-ice/Defog Rain Removal System.............................................. 10-6

10-7 IN NORMAL/OUT DE-FOG Knob .................................................................... 10-5

10-8 WSHLD HEAT Switches ..................................................................................... 10-7

10-9 Windshield Heat Diagram.................................................................................... 10-8

10-10 Alcohol Reservoir ................................................................................................ 10-9

10-11 WSHLD & RADOME Switch ............................................................................. 10-9

10-12 Alcohol System—SNs 24-181 through 24-349 and 25-025 through 25-226 and Airplanes Modified for Known Icing Conditions ........................... 10-10

10-13 Alcohol System—SNs 24-350 and Subsequentand 25-227 and Subsequent................................................................................. 10-11

10-14 Wing and Horizontal Stabilizer Anti-ice System............................................... 10-12

10-15 Wing Temperature Indicator .............................................................................. 10-13

10-16 Pitot Tubes and AOA Vanes............................................................................... 10-14

10-17 L and R PITOT HEAT Switches ........................................................................ 10-14

INTRODUCTIONAnti-icing equipment on the Learjet 20 series is designed to prevent buildup of ice on:

• The engine nacelle lip, inlet guide vanes, and nose cone

• The windshield and radome

• The leading edge of the wings and horizontal stabilizer

• Pitot probes, static ports, AOA vanes, and shoulder static ports

The system is certificated for flight into known icing conditions on SNs 24-181 and sub-sequent, 25-025 and subsequent, and earlier SNs modified for flight into icing conditions.

Airplane anti-icing is accomplished throughthe use of electrically heated anti-ice systems,engine bleed-air anti-ice systems, and an al-cohol anti-ice system.

Electrically heated components include pitottubes, static ports, shoulder static ports, stallwarning vanes, nacelle inlet, and horizontalstabilizer.

CHAPTER 10ICE AND RAIN PROTECTION




GENERAL

Engine bleed air is used to heat the windshields,wing leading edges, and the engine front frame.

An alcohol system is installed for radome anti-icing and as a backup to the pilot’s windshieldbleed-air anti-icing.

NOTEAll anti-ice systems except enginefront frame heat require electricalpower for operation.

All anti-icing equipment must be turned on before icing conditions are encountered. Todelay until ice buildup is visually detected onaircraft surfaces constitutes an unacceptablehazard to safety of flight.

Icing conditions exist when visible moistureis present and the indicated ram-air tempera-ture (RAT) is +10º C or below. Takeoff intoicing conditions is permitted with all bleed-air anti-icing systems on. The RAT gage shouldbe checked frequently when flying in or en-tering areas of visible moisture.

During descents, the cabin altitude may in-crease unless sufficient engine rpm is main-ta ined to compensate for the addi t ionalbleed-air use. To ensure adequate bleed airfor anti-ice protection and cabin pressuriza-tion, maintain at least 80% engine rpm.

Figure 10-1 illustrates the ice-protected areas.




HORIZONTAL STABILIZERLEADING-EDGE ANTI-ICE

(ELECTRIC)

FRONT FRAMEANTI-ICING

(BLEED AIR)

WING LEADING EDGEANTI-ICE (BLEED AIR)

WINDSHIELD HEAT(BLEED AIR AND ALCOHOL)

NACELLE LIP ANTI-ICE(ELECTRIC)

RADOME ANTI-ICE(ALCOHOL)

Figure 10-1. Ice and Rain Protected Areas

ICE DETECTIONDuring daylight operation, ice accumulationcan be visually detected on the windshield, thewing leading edges, and tip tanks.

WINDSHIELD ICE DETECTIONDuring night operations, the windshield icedetection lights indicate ice or moisture for-mation on the windshield. Two probes, one onthe pilot’s side of the glareshield and one on thecopilot’s side, contain red lights which con-tinuously shine on the inside of the windshieldsurface (Figure 10-2). The ice detection lightsnormally shine through unseen. They reflect redspots approximately 1 1/2 inches in diameterif ice or moisture has formed on the windshield.

The ice detection light on the pilot’s side is in-side the anti-ice airstream; the light on thecopilot’s side is located outside the anti-iceairstream. For this reason, the copilot’s lightshould be monitored whenever windshieldheat or the alcohol anti-ice system is in oper-ation. The ice detection lights are illuminatedwhenever airplane electrical power is on. Thelights use DC power routed through the L andR ICE DET circuit breakers located on thepilot’s/copilot’s or baggage compartment circuit-breaker panels.

WING ICE DETECTIONDuring darkness, the recognition light (Figure10-3) can be used to confirm ice buildup on theright tip tank. Some airplanes are equipped witha second recognition light in the left tip tank.

The optional wing ice inspection/egress light,installed on the right side of the fuselage, allows ice buildup to be detected on the winginboard leading edge.

ANTI-ICE SYSTEMS

ENGINE ANTI-ICE SYSTEM(NACELLE HEAT)The engine anti-ice system (Figure 10-4) pro-vides anti-ice protection for the engine na-celle inlet lips and the inlet guide vanes andnose cone. The nacelle lips are heated elec-trically and the front frame pneumatically.




Figure 10-3. Recognition Light

WINDSHIELD(REF)

ICE DETECTORASSEMBLY

FORWARDGLARESHIELD

UP

FWD

Figure 10-2. Ice Detection Lights




Fig

ure

10-

4.E

ng

ine

An

ti-i

ce S

yste

m

R E

NG

ICE

NA

CE

LLE

HE

AT S

W28

VD

CON

OF

F

NA

CH

EAT

EIT

HE

R G

EN

ER

ATO

R

WA

RN

LIG

HT 28

VD

C

INLE

T H

TR

SQ

UAT

SW

RE

LAY

BO

X

Each engine anti-ice system is independentlycontrolled by the L and R NAC HEAT switches(Figure 10-5) located on the anti-ice control panel.

When a NAC HEAT switch is turned on, elec-tric power is supplied to heat the nacelle lip.This is done by energizing a nacelle heat relaywhich allows 28 VDC to flow from the gen-erator bus to the nacelle heating elements.Each nacelle requires 50 to 60 amperes.Simultaneously, power is removed from the en-gine anti-ice valve allowing bleed air to flowthrough the hollow inlet guide vanes, vari-able inlet guide vanes, and the engine nosecone. The bleed air is discharged into the inlet,where it reenters the engine. This heated air,being less dense, causes a decrease in the en-gine pressure ratio (EPR).

Engine Ice LightsThe amber L and R ENG ICE lights on theglareshield annunciator panel (see “AnnunciatorPanel” section) provide a visual indication ofinsufficient bleed-air pressure for adequatefront frame anti-ice protection. A pressureswitch on the front frame actuates and causesillumination of the lights whenever bleed-airpressure drops to less than 5 psi. On the ground,approximately 70% rpm is required to extin-guish the lights.

Inlet Heater LightA temperature-sensing switch is located in thenacelle on each engine to detect a possibleoverheat condition. A single amber INLETHTR light (see “Annunciator Panel” section) onthe glareshield illuminates when either nacellereaches 190 ±3ºF and extinguishes at 180ºF.

This light is connected through the squat switchrelay box subsequent to SNs 24-209 and 25-045so that it does not illuminate in flight. If theINLET O’HEAT light illuminates on the ground,the NAC HEAT switches must be turned off.

WINDSHIELD ANTI-ICE/DEFOGAND RAIN REMOVAL SYSTEM

GeneralThe windshield heat system utilizes bleed air fordefogging, rain removal, and anti-ice protection.

The windshield heat/defog system (Figure 10-6)can be controlled either automatically or manu-ally. The system is also used to supplement cock-pit heating through the pilot’s footwarmers andto provide an alternate bleed-air source for emergency pressurization.

An IN NORMAL/OUT DE-FOG knob locatedbelow the instrument panel to the left of thepedestal (Figure 10-7) controls a two-wayvalve within the ducting which directs bleedair to the windshield or cockpit footwarmers.




Figure 10-5. L and R NAC HEAT Switches

Figure 10-7. IN NORMAL/OUT DE-FOG Knob




LOW-LIMITTHERMOSTAT

HIGH-LIMITTHERMOSTAT

WSHLD OV'HT WSHLD OV'HT

FRAME 5(FWD PRESSUREBULKHEAD)

(AFT PRESSUREBULKHEAD)

WINDSHIELD

INNORMAL

OUTDE-FOG

OVERBOARDDRAIN

CHECK VALVECHECK VALVE

DEFOG PRESSURE REGULATOR VALVE

WSHLD HEAT

RHENGINE

LHENGINE

BLEED AIR

TO WING L EHEAT DUCT

DEFOG SHUTOFFVALVE

ON

OFF

AUTO

MANFOOTWARMERS

WINDSHIELD HEAT

BLEED AIR

TO CABINFLOW CONTROL

VALVE

BLEED-AIRMANIFOLD

Figure 10-6. Windshield Anti-ice/Defog Rain Removal System

When the knob is pushed in to NORMAL,bleed air is diverted to the pilot’s and copilot’sfootwarmers, increasing the amount of flowinto the cabin, which effectively becomes theemergency source for cabin pressurization.The knob is normally kept pushed in.

When the knob is pulled out to the DE-FOG po-sition, the bleed air is directed to the externalwindshield duct outlets for heating and anti-icing.

Two WSHLD HEAT switches (Figure 10-8) arelocated on the anti-icing panel. One is labeled“ON” and “OFF” and is spring-loaded to theneutral position; the other switch is labeled“AUTO” and “MAN.”

Bleed air from the manifold is routed throughtwo valves—the shutoff valve and the pressure-regulator valve. Figure 10-9 illustrates thewindshield heat system.

The shutoff valve is motor driven by either ofthe two switches on the anti-ice control paneland takes four to five seconds to cycle fully.Selecting AUTO opens the shutoff valve. Thegreen WSHLD HEAT light illuminates when-ever the shutoff valve is not fully closed. IfMAN is selected, the shutoff valve may beopened with the ON–OFF switch. Since thisswitch is spring-loaded to neutral, it must be

held in the ON position while the valve drivestoward the fully open position. The switchcan be released at any intermediate position,and the valve will stop moving. Since theWSHLD HEAT light illuminates when thevalve is in any position other than fully closed,there is no cockpit indication of intermediateshutoff valve position. Once the valve hasbeen opened, it can be closed only by holdingthe ON–OFF switch to OFF (with MAN se-lected) for at least four seconds.

The pressure-regulator valve is solenoid op-erated and is deenergized closed. It is energizedopen when airplane power is on. Its functionis to regulate the engine bleed air from the man-ifold to 16 psi.

Automatic OperationThe flow of bleed air to the windshields is con-trolled in the automatic mode by the high (250ºF)and low (215ºF) temperature thermoswitches installed in each windshield outlet nozzle.

On the ground, if the low-limit thermoswitchsenses 215ºF, it closes the shutoff valve, ex-tinguishes the green WSHLD HEAT light, andilluminates the red WSHLD OV’HT light. Ifthe low-limit switch fails or the shutoff valvefails to close fully, the temperature may risesufficiently to trigger the high-limit temper-ature switch, which causes the pressure-regulator valve to close. The red WSHLDOV’HT l ight i l luminates , and the greenWSHLD HEAT light remains illuminated be-cause the shutoff valve is not fully closed.

In flight, the windshield heat cycles off thelow-limit thermoswitches. When the temper-ature reaches 215ºF, the low-limit switchescomplete a circuit that drives the shutoff valveto the closed position, extinguishing the greenWSHLD HEAT light. When the switch cools,the shutoff valve opens, the WSHLD HEATlight illuminates, and the cycle repeats. If thehigh-temperature limit thermoswitches reach250ºF due to failure of the low-limit switches,the pressure-regula tor valve c loses , theWSHLD OV’HT lights illuminate, and theWSHLD HEAT light remains illuminated.




Figure 10-8. WSHLD HEAT Switches

Manual OperationSelecting MAN enables the spring-loadedON–OFF switch to control the shutoff valveand, therefore, the amount of bleed air suppliedto the windshields.

On the ground, in manual mode, a low-limitthermoswitch illuminates the WSHLD OV’HTlight, but does not close either the regulatorvalve or shutoff valve.

If the temperature activates the high-limitswitch, the pressure-regulator valve closes.

In flight, the low-limit switches are disabled.Overheat protection is provided by the high-limit switches which close the pressure-regulator valve, if activated.

WINDSHIELD/RADOMEALCOHOL ANTI-ICE SYSTEM

GeneralMethyl alcohol, from a reservoir (Figure 10-10) located in the left side of the nose com-partment, is provided to prevent ice formation




CONDITIONED AIR

BLEED-AIR DUCT

MANUAL DEFOG CONTROL VALVE

EXTERNAL DEFOG OUTLET

HIGH-LIMIT THERMOSWITCH

LOW-LIMIT THERMOSWITCH

EXTERNAL DRAIN

FOOTWARMER

INTERNAL DEFOG NOZZLE

DEFOG CONTROLKNOB

FWD

Figure 10-9. Windshield Heat Diagram

on the radome. If necessary, this system maybe used as a backup for the windshield anti-ice defog system for the pilot’s windshield.There are two different systems in use. The sys-tems are operated by DC power routed throughthe ALC PUMP circuit breaker.

SNs 24-181 through -349, and25-003 through -226, and Earlier Airplanes Modified for All-Weather OperationA DC motor-driven pump supplies filtered al-cohol from a 2 1/4 gallon reservoir to theradome only or to the radome and pilot’s wind-shield, depending on the position selected onthe WSHLD & RADOME switch (Figure 10-11) on the pilot’s anti-icing control panel.

Figure 10-12 illustrates the early system.

When the switch is positioned to RADOME,the pump is energized and alcohol is deliveredto the radome only. In this position, a fully ser-viced reservoir should dispense alcohol forapproximately 1 1/2 hours.

When the switch is positioned to WSHLD &RADOME, the pump is energized and alcoholis delivered to both surfaces. Flow to the wind-shield is dispensed through an orifice assem-bly integrated with the pilot’s defog outlet. Inthis position, a fully serviced reservoir shoulddispense alcohol for approximately 45 minutes.

A pressure switch installed in the radome supplyline actuates the amber ALC AI light when thereservoir is empty or if the pump fails. The lightextinguishes when the control switch is turned off.

SNs 24-350 and Subsequent,and 25-227 and SubsequentMethyl alcohol is stored in a 1.75-gallon reser-voir. When the cockpit control switch is po-s i t i oned t o WSHLD & RADOME or t oRADOME, circuits are completed to positiona three-way valve in the fluid supply line(Figure 10-13) This position also opens theshutoff valve and pressure regulator in theservo bleed-air supply line.

Servo bleed air trapped from the bleed-airmanifold passes through the shutoff valve andpressure regulator where it is regulated to 2.3psi. The air is then routed to pressurize the al-cohol reservoir.

The alcohol is forced through a filter to thethree-way valve which is positioned accord-ing to the selected switch position.

The pressure relief valve, set to 2.6 psi, relievesany overpressure in the reservoir should thepressure regulator fail. Residual pressure bleedsoff when the control switch is turned off.




Figure 10-11. WSHLD & RADOME Switch

SYSTEM RETURNLINE

INSTRUCTION PLACARDS

INSTRUCTIONPLACARDS

FILLER CAP

FILLER CAP

CHECK VALVE

VENT LINE

SUPPLY TANK

SUPPLY TANK

OVERBOARD VENT LINE

REGULATEDPRESSURELINE

AIRPLANES 25-061, 25-070 THROUGH 25-226

AIRPLANES 25-227 AND SUBSEQUENT, 24-350AND SUBSEQUENT

Figure 10-10. Alcohol Reservoir

A float switch in the reservoir illuminates theALC AI annunciator when the tank is empty.The light stays on even if the switch is off asa reminder to service the reservoir.

When the RADOME position is selected, afully serviced reservoir supplies only theradome with alcohol for approximately twohours and nine minutes. When positioned toWSHLD & RADOME, alcohol is also dis-

pensed to the pilot’s defog outlet via the three-way valve, and duration of the supply is re-duced to approximately 45 minutes.

NOTEThis system is still operational ifboth emergency pressurization valvesare in emergency (provided DCpower is available).




OFF

RADOME

WINDSHIELD/RADOME

ORIFICE ASSEMBLY

ANTI-ICEVALVE (N.C.)

LOW-PRESSURESWITCH

PRESSURERELIEF

OVERBOARDVENT

FILTERMOTOR-DRIVENPUMP

RADOME

ALC AI

PILOT’S EXTERNALDEFOG OUTLET

SUPPLY

PRESSURE

LEGEND

Figure 10-12. Alcohol System—SNs 24-181 through 24-349 and 25-025 through 25-226 andAirplanes Modified for Known Icing Conditions




FLOATSWITCH

SUPPLY TANK

BLEED AIRFROM RH ENGINE

CHECKVALVE

CHECKVALVE

PILOT’S WINDSHIELDOUTLET

THREE-WAYSHUTOFFVALVE

FILTER

SHUTOFFVALVE ANDPRESSURE

REGULATOR

OVERBOARDDRAIN

OVERBOARDDRAIN

BLEED AIRFROM LHENGINE

PRESSURE-RELIEFVALVE

ALC AI

RADOMEOUTLET

ELECTRICAL

BLEED AIR (UNREGULATED)

REGULATED PRESSURE

LEGEND

ALCOHOL PRESSURE

Figure 10-13. Alcohol System—SNs 24-350 and Subsequent and 25-227 and Subsequent




WING TEMP

WING TEMP


WING OV'HT

SCUPPER SCUPPER

STAB AND WINGHEAT

CONTROL VALVE

LH ENGINE RH ENGINE

TO DEFOG SYSTEM

TO AIR-CONDITIONING (PRESSURIZATION) SYSTEM

STAB HEAT

ANTI-ICE BLANKET

ON

OFF

RELAYBOX

SEQUENCETIMER

Figure 10-14. Wing and Horizontal Stabilizer Anti-ice System

WING AND HORIZONTALSTABILIZER ANTI-ICE SYSTEM

Wing HeatBleed air is used to prevent ice formation onthe wing leading edges. The bleed air is di-rected from the bleed-air manifold through asolenoid-operated pressure-regulator valve(control valve) (Figure 10-14) to the respec-tive leading-edge surfaces.

Controls and IndicationsThe STAB & WING HEAT switch (Figure 10-14) is located on the pilot’s anti-icing controlpanel. When the switch is turned to ON, thevalve is energized open. With the switch inOFF or if DC power fails, the valve fails to theclosed position.

With the valve open, bleed air is routed throughthe wing pressure-regulator valve where it isregulated to 16 psi and then routed on to tubesin the leading edges of the wing. After thebleed air has heated its respective leadingedge, it continues outboard where it is ventedoverboard through a scupper drain on the bot-tom, outer edge of the wing. The WINGOV’HT light (see “Annunciator Panel” section)on the glareshield illuminates when the rightwing leading-edge structure temperaturereaches 215ºF. The system must not be oper-ated when the light is illuminated. The ther-mocouple that operates the light is located atthe right wing fuselage root on the structure.

A wing temperature indicator (Figure 10-15)located on the pilot’s subpanel has three col-ored ranges on a round dial. From left to right,these arcs are red, green, and yellow.

Red arc—Wing temperature is below 35ºF.

Green arc—Wing temperature is above 35ºFand below 215ºF.

Yellow arc—Wing skin temperature is above215ºF. (When in the yellow arc, reduce rpm,and also monitor pressurization.)

NOTEOn SNs 25-260 and 24-358 and sub-sequent, the red segment has beenchanged to blue and the yellow seg-ment has been changed to red.

Stabilizer HeatThe horizontal stabilizer leading edge is pro-tected from ice accumulation by an anti-icesystem consisting of two electrical blanketscon t ro l l ed by a STAB & WING HEATON–OFF switch (Figure 10-14) located onthe pilot’s subpanel. With the switch in ON,electrical power is supplied from the battery-charging bus through a relay to the sequencetimer which controls the heating elements.On the ground, stabilizer heat is disabled bythe right squat switch.




Figure 10-15. Wing Temperature Indicator

With the airplane in flight and the STAB &WING HEAT switch in ON, intermittent DCelectrical power is distributed by the sequencetimer to individual heating elements in theblankets. Power is distributed in a forward-to-aft sequence of 15 seconds duration each. Thesingle leading-edge element is suppled withcontinuous electrical power. An amber STABHEAT light (see “Annunciator Panel” section)on the glareshield illuminates when power islost to the leading-edge wire. The airplaneammeters reflect a cyclic current drain of ap-proximately 60 amperes total in 15-secondcycles. This indicates proper operation of thesystem.

PITOT, STATIC, AND ANGLE-OF-ATTACK VANE ANTI-ICESYSTEM

Pitot and Angle-of-Attack VaneAnti-icingThe left and right pitot tubes and angle-of-at-tack (AOA) vanes (Figure 10-16) contain elec-trical heating elements.

When the L and R PITOT HEAT switches(Figure 10-17) located to the pilot’s subpanelare in ON, both pitot tubes and both stall vanesare heated. A condensate drain valve is installedfor each system. One is located adjacent to eachnose wheel well door. Some airplanes may havea light to indicate that a system has failed or thatthe PITOT HEAT switch is not turned on.

The static port heaters are wired through the re-spective PITOT HEAT circuit breakers but notthrough the PITOT HEAT switches. Therefore,static port heat operates when either batteryswitch or generator is turned on. The rear staticport on the right side is unheated.

There are two shoulder static ports located ontop of the nose in front of the windshield.These two static ports provide static pressureto the autopilot altitude controller. A con-densate drain valve is installed adjacent to theleft nosewheel gear door.

On aircraft with the Learjet Century III wing,the shoulder ports are heated whenever a bat-tery or generator is turned on and the PITOTHEAT circuit breakers are set.




Figure 10-16. Pitot Tubes and AOA Vanes

Figure 10-17. L and R PITOT HEAT Switches

1. Bleed air is used for anti-icing on:A. WindshieldsB. Engine front frameC. Wing leading edgesD. All the above

2. The L or R PITOT HEAT switches alsosupply heating element power for:A. The angle-of-attack vanesB. The shoulder static portsC. The instrument static portsD. The ice detect lights

3. The crew action required when the redWING OV’HT light illuminates is:A. No action is required, the system is

automatic.B. Position the STAB & WING HEAT

switch to STAB.C. Turn the STAB & WING HEAT

switch to OFF.D. Reduce power below 80% rpm.

4. Anti-icing equipment should be turned on:A. When in icing conditionsB. Before entering icing conditionsC. Before takeoffD. During climbout

5. With the loss of airplane electricalpower, anti-icing is lost for:A. All systemsB. Pitot tubes, static ports, and AOA

vanes onlyC. All systems except the nacelle inlet lipsD. All systems except the engine front

frame

6. When the NAC HEAT switches areturned on, the most probable cause forboth ENG ICE lights illuminating is:A. A DC power failureB. A short circuit to the front frame

anti-ice valveC. Less than 5 psi in the front frame

manifoldsD. Excessive temperature in the front

frame manifold




QUESTIONS

10A-iFOR TRAINING PURPOSES ONLY



CHAPTER 10AICE AND RAIN PROTECTION

CONTENTS

Page

INTRODUCTION .............................................................................................................. 10A-1

GENERAL.......................................................................................................................... 10A-1

ICE DETECTION .............................................................................................................. 10A-2

ANTI-ICE SYSTEMS........................................................................................................ 10A-3

Engine Anti-ice System (Nacelle Heat) ...................................................................... 10A-3

Windshield Anti-ice/Defog and Rain Removal System.............................................. 10A-5

Windshield Alcohol System........................................................................................ 10A-9

Wing Anti-ice System................................................................................................. 10A-9

Pitot, Static, and Angle-of-Attack Vane Anti-ice System......................................... 10A-11

QUESTIONS.................................................................................................................... 10A-12

10A-iiiFOR TRAINING PURPOSES ONLY



ILLUSTRATIONS

Figure Title Page

10A-1 Ice and Rain Protected Areas .............................................................................. 10A-2

10A-2 Ice Detection System........................................................................................... 10A-2

10A-3 Ice Detection Lights ............................................................................................ 10A-3

10A-4 Engine Anti-ice System....................................................................................... 10A-4

10A-5 Defog Push-Pull Knobs....................................................................................... 10A-5

10A-6 Defog System—Model 23................................................................................... 10A-6

10A-7 Defog System—SNs 24-100 through -180, and 25-003 through -024 ............... 10A-7

10A-8 Defog Knob ......................................................................................................... 10A-8

10A-9 Windshield Alcohol System................................................................................ 10A-9

10A-10 Wing Anti-icing System.................................................................................... 10A-10

10A-11 Pitot, Static, and AOA Anti-ice System............................................................ 10A-11

INTRODUCTIONAnti-icing equipment on Learjet SNs 23-003 through 24-180, and 25-002 through -024not certified for flight in known icing conditions is provided and approved for engineprotection only. The wing anti-ice system should be operated in concurrence with nacelle heat. Protection is provided on the following:

• Engine nacelle lip and engine front frame

• Windshield

• Wing leading edge (not certified for flight into known icing conditions)

The anti-icing equipment utilizes either electri-cally heated units or engine bleed air. Windshieldalcohol is used for deicing. Rain removal is ac-complished with bleed air by a service kit.

Engine bleed air is used to heat the windshields,the wing leading edge, and the engine frontframe. The engine front frame includes the guidevanes, variable guide vanes, and the nose cone.

Icing conditions exist if visible moisture is present and the ram-air temperature (RAT) is10ºC or lower. During daylight operation, iceaccumulation can be detected on the windshieldand on the wing leading edge, particularly wherethe wing and tip tank join and on the tip tanks.

Figure 10A-1 illustrates the ice and rain pro-tected areas.

CHAPTER 10AICE AND RAIN PROTECTION




GENERAL

ICE DETECTIONThe ice detection system (Figure 10A-2) in-cludes an ice detector and interpreter mountedin the left engine intake at the two o’clockposition as viewed from the front of the intake.A single detector and one red ICE DETECTlight on the warning panel comprise the warn-ing system. The system is protected with a10-ampere NAC ICE DETECT circuit breaker.

Pressure differential provides mechanical powerthat allows a three-phase cycling operation.The phases are neutral, sensing, and deicing.When the detector probe is subjected to ram airof 40 knots or more, a pressure differential is




FRONT FRAMEANTI-ICING

(BLEED AIR)

NACELLE CAP ANTI-ICE(ELECTRIC BLANKET)

WING LEADING-EDGEANTI-ICE (BLEED AIR)

WINDSHIELD ANTI-ICE(BLEED AIR AND ALCOHOL)

Figure 10A-1. Ice and Rain Protected Areas

LOCATION OFICE-DETECTIONPROBE

Figure 10A-2. Ice Detection System

LOCATION OFICE-DETECTIONPROBE

created which distends the detector diaphragm,actuating the diaphragm switch. This actionmoves the system to the sensing phase by en-ergizing the interpreter arming relay and di-recting current to the blade of the diaphragmswitch. When the probe perforations becomeblocked with ice, the resulting attenuated pres-sure differential allows the diaphragm to relaxand return to its normal position. Electricalpower illuminates the red ICE DETECT warn-ing light (see “Annunciator Panel” section) andheats the probe. The pressure differential isthen restored, and the system reverts to thesensing phase. It cycles between the sensing anddeicing phases until icing conditions abate.The system is completely automatic and can-not be adjusted.

For those airplanes without tip tank taxi lightsor the above system, windshield ice detectionlights (Figure 10A-3) may be installed.

The ice detection lights consist of two probesmounted on top of the glareshield. Inside eachprobe is a red light. Each probe receives 28 VDCfrom a separate ICE DETECT circuit breaker.

During night flight, the windshield ice de-tection lights (if installed) cause red areas,

approximately 1 1/2 inches in diameter, to appearon the windshield if particles of ice or moistureform. The light on the pilot’s side is located inthe defog airflow stream, and the light on the copi-lot’s side is outside the defog airflow stream. Ifthe windshield defog system is operating, thecopilot must monitor the light on his side for in-dication of ice or moisture formation. The wind-shield ice detection lights indicate moistureencounters when OAT is above freezing. At belowfreezing OAT, the lights indicate ice encounters.

ANTI-ICE SYSTEMS

ENGINE ANTI-ICE SYSTEM(NACELLE HEAT)Nacelle heat is controlled by a LH or RH NACHEAT switch (Figure 10A-4) on the mainswitch panel or pilot’s subpanel. Nacelle lipheating uses 28 VDC, and front frame heat isprovided by engine bleed air from the re-spective engine. The bleed air is used to heatthe hollow inlet struts, variable inlet guidevane, and the engine nose cone. After passingthrough the front frame, the bleed air is dis-charged into the engine inlet.




WINDSHIELD(REF)

ICE DETECTORASSEMBLY

FORWARDGLARESHIELD

UP

FWD

Figure 10A-3. Ice Detection Lights

Engine Anti-ice Control ValveThe engine anti-ice control valve is located onthe left forward side of the engine. The valveis spring-loaded to the open position. In orderto keep the valve in the closed position, 28VDC and bleed-air pressure are required. Ifelectrical power is lost, the spring in the valvecauses the valve to open. Therefore, during anelectrical power failure, the valve opens andallows bleed air to flow to the front frame andvanes. With the NAC HEAT switch in the OFFposition, 28 VDC is applied to the anti-icecontrol valve solenoid and directs air pres-sure to close the valve.

Engine Ice LightThe anti-ice pressure switch is located down-stream of the anti-ice control valve. Wheneverthe NAC HEAT switch is in ON and the pres-sure in the front frame drops below 5 psi, acircuit is completed that illuminates the amberENG ICE light (see “Annunciator Panel” section)

on the warning panel. Therefore, the ENGICE light only indicates insufficient bleed-airpressure to the front frame. Adequate pressureto the front frame requires approximately70% rpm on the ground and at least 80% rpmwhile airborne.

Electrical power for heating the nacelle lip isprovided by its respective generator. Loss ofa generator results in the loss of the associatednacelle heat. Each nacelle uses approximately50 amperes for heating.

A temperature-sensing switch is located inthe nacelle on each engine to detect a possi-ble overheat condition. A single amber INLETO’HEAT light on the warning panel illuminateswhen either nacelle temperature reaches 210ºF.The light extinguishes when nacelle temper-ature decreases to 190ºF. If the light illumi-nates in flight, it may be disregarded; however,for ground illumination, nacelle heat must beturned off.




R ENG ICE

ON

OFF

NACHEAT

INLET HTR

ICE DETECT

0 1020

3040

50608070

10090

ABOVE 70%

Figure 10A-4. Engine Anti-ice System

WINDSHIELD ANTI-ICE/DEFOGAND RAIN REMOVAL SYSTEM

Defog System—Model 23Windshield defogging is accomplished by di-recting engine bleed air over the outside of thewindshield or across the inside of the windshield,or by allowing conditioned cabin air to flow be-tween the dual panes of the windshield. The sys-tem may be used for emergency pressurizationas well as external and internal defogging.

External DefogIn most cases, there are three defog push-pullknobs located in the cockpit. One external andone internal control knob is located next to thepedestal on the left (Figure 10A-5), and one ex-ternal control control knob is on the right side.For outside airflow, pull both external controlknobs out, and push the internal knob in, thentoggle the DEFOG switch to the DEFOG po-sition for three to four seconds. This should pro-duce the desired airflow. To turn off the defogsystem, toggle the DEFOG switch to OFF po-sition for three to four seconds until the valveis closed, then place the external control knobsin and pull the internal control knob out.

On SNs 23-003 through -008, a thermal switchin the defog duct automatically closes thedefog shutoff valve (via the defog relay) whenthe temperature reaches 205ºF (Figure 10A-6). The defog relay then deenergizes whenthe temperature drops to 180ºF, at which timethe defog system can be turned back on.

On SNs 23-009 through -099, a thermal switchin the defog duct illuminates the red W/SO’HEAT light on the warning panel when tem-perature reaches 205ºF. the light extinguishesat 180ºF. The defog system should be turnedoff or airflow decreased when this light illu-minates in order to protect the windshieldfrom possible distortion.

The water accumulator, installed in the ex-ternal defog ducting forward of frame 5, pre-vents any accumulated water in the externalnozzles from entering the internal ducting.The accumulators are vented overboardthrough the nose compartment. A check valveis also installed between the internal and ex-ternal outlets to prevent a loss of cabin pres-surization through the external defog nozzles.

Internal DefogFor internal defog or emergency pressuriza-tion, the two external control knobs must bepushed in and the internal control knob pulledout. Toggle the spring-loaded DEFOG switchto the DEFOG position for three to four sec-onds until maximum airflow is attained.

Conditioned cabin air is routed through a clearplastic dehumidifier before passing between thewindow panes. This prevents moisture, dirt, orsmoke from collecting between the panes.

Defog System—SNs 24-100through -180, and 25-003through -024Windshield defogging is accomplished by di-recting engine bleed air over the outside of thewindshield and conditioned cabin air over theinside of the windshield. The defog system(Figure 10A-7) can be used for either exter-nal defogging or emergency pressurization.




Figure 10A-5. Defog Push-Pull Knobs




LH EXTERNALDEFOG OUTLET

RH EXTERNALDEFOG OUTLET

(OPTIONAL)

INTERNALOUTLETS

PLENUM CHAMBER PLENUM CHAMBER

CHECK VALVE

INTERNAL DEFOGCOPILOT’S

EXTERNAL DEFOG

WATERACCUMULATOR

WATERACCUMULATOR

PULLPULL

PULL

TO FOOTWARMER

DEFOG OFF

PILOT’SEXTERNAL

DEFOG

DEFOG RELAY

SHUTOFFVALVE

220°–180°F

W/S O’HEAT

SHUTOFFVALVE

SHUTOFFVALVE

CHECK VALVE

TOAIR-

CONDITIONINGSYSTEM

* SNs 23-003 THROUGH 23-008

RHENGINE

LHENGINE

Figure 10A-6. Defog System—Model 23




LHENGINE

BLEED AIRRHENGINE

DEFOG

BLEED-AIRMANIFOLD

TO WIND LEHEAT DUCT

CHECK VALVE

TO CABINFLOW CONTROL

VALVE

EMER CAB PRESSEXT DEFOG–ON

DEFOG OFF CHECKVALVE

CONDITIONED AIR FROM CABIN DISTRIBUTION SYSTEM

FOOTWARMERS

INNORMAL

OUTDEFOG

W/S O’HEATIN

PILOT AND COPILOTOUT

PILOT

PILOT’SEXTERNAL OUTLET

WINDSHIELD

INTERNAL OUTLET

COPILOT’SEXTERNAL OUTLET

CHECKVALVE

BLEED AIR

210°F

Figure 10A-7. Defog System—SNs 24-100 through -180, and 25-003 through -024

External DefogThere are two push-pull defog knobs locatedon the left side of the pedestal. The top knob(Figure 10A-8) is labeled “IN–NORMAL—OUT–DEFOG,” and the bottom is labeled“IN–PILOT & COPILOT—OUT–PILOT.” Forexternal use, pull the top knob out, then tog-g l e t he t h r ee -pos i t i on , sp r i ng - loadedDEFOG–OFF switch to the DEFOG positionfor three to four seconds until the desired air-flow is attained. Adjacent to the DEFOGswitch is an amber or blue light which illu-minates anytime the defog shutoff valve is notfully closed.

In the pilot’s defog nozzle is a thermal switchwhich causes the red W/S O’HEAT light to il-luminate when the temperature reaches 210ºF.

The light extinguishes when the temperaturedrops to 180ºF. When the W/S O’HEAT lightilluminates, the defog system should be turnedoff or the airflow decreased to prevent possi-ble distortion to the windshield. To turn off thedefog system, toggle the DEFOG switch toOFF until airflow has stopped, then push thetop knob in.

Internal DefogInternal windshield defogging is supplied bya continuous flow of conditioned cabin airand requires no management.

A defog nozzle drain located on the right sideof the fuselage under the right windshielddrains condensation from the lines.




Figure 10A-8. Defog Knob

WINDSHIELD ALCOHOLSYSTEM

SNs 24-100 through -180, and25-003 through -024The windshield alcohol system (Figure 10A-9)contains one gallon of methyl alcohol storedin a tank located in the left nose compartmentbetween frames 2 and 3.

Operation of the windshield alcohol system iscontrolled by the pilot. A variable rheostat or

a two-position switch may be used to controlthe pump motor speed. The maximum flowrate is 5 to 10 gph through the two externaldefog nozzles.

WING ANTI-ICE SYSTEMEighth-stage engine bleed air is provided toprevent the formation of ice on the wing lead-ing edges and is certified for use per FAR 25for flights below 22,000 feet. Bleed air ispassed through a tube with small holes that allows the heated air to be discharged against




DEFOG OUTLET

TUBE ASSEMBLY—RIGHT

BRACKET

DEFOG OUTLET

TUBE ASSEMBLY—LEFT

BRACKET

FILLER CAP

TANK ASSEMBLYVENT TUBE

PUMP

FILTER

TANK VENT

Figure 10A-9. Windshield Alcohol System

the leading edge (Figure 10A-10). The air isthen exhausted overboard through a scupperdrain on the bottom of the leading edge nearthe tip tanks.

The control switch is usually located on thepilot’s subpanel. The ON–OFF switch con-trols electrical power to the wing regulatorvalve located in the tail compartment.

The wing regulator valve is electrically con-trolled and pneumatically operated. These air-planes have one of two kinds of regulator

valves installed. The earlier kind is a Janitrolvalve which is electrically closed and fails tothe open position. The later kind is a Whittakervalve which is electrically energized open andfails to the closed position. Either valve reg-ulates pressure to the wing ducts at 8 to 20 psi.

A green WING HEAT light illuminates whenthe duct temperature reaches 215ºF. A redWING OVERHEAT light illuminates whenthe leading-edge structure temperature reaches215ºF. Do not operate the system with the redlight illuminated.




WING

WINGOVERHEAT

HEAT

WINGANTI-ICESWITCH

WINGON

OFFANTI-ICE

SCUPPER

SENSE LINEREGULATOR VALVE(WING LE ANTI-ICE)

DEFOG SHUTOFF VALVE

TO DEFOG SYSTEM

RH ENGINELH ENGINE

TO AIR-CONDITIONING

SYSTEM

BLEED-AIRMANIFOLD

Figure 10A-10. Wing Anti-icing System

PITOT, STATIC, AND ANGLE-OF-ATTACK VANE ANTI-ICESYSTEM

Pitot and Angle-of-Attack VaneAnti-icingEach pitot head is heated with a cockpit-controlled electrical heating element to pre-vent moisture from freezing and obstructingthe heads (Figure 10A-11). The PITOT HEATswitch for each pitot tube is located on theswitch panel. Two drain valves (except onSNs 23-003 through 24-137) for draining anycondensation prior to flight are located onthe lower skin adjacent to the nose wheelwell. The angle-of-attack vanes are heatedwhenever the PITOT HEAT switches areturned on.

Static Port HeatingThe five static ports on SNs 23-003 through24-137 are unheated unless SK 23/24-226 hasbeen completed. If SK 23/24-226 has beencompleted, the two forward ports on each sideof the nose are heated anytime the batteryswitch is turned on. The third and rearmoststatic port on the right side is not heated. Itdoes, however, have an alternate source insidethe nose compartment.

NOTEOn airplanes with a Century III wing,the shoulder static ports are heated.

NOTESNs 23-003 through -069 not modi-fied by SK 23/24-226 do not haveshoulder static ports.




LH SHOULDERSTATIC PORT

LH PITOTSTATIC PORT

LH FWDSTATIC

PORTAOA

VANE

LH AFTSTATIC PORT

PITOTHEAT

SWITCH

RH SHOULDERSTATIC PORT

RH PITOTSTATIC PORT

RH FWDSTATICPORT

AOAVANE

RH CENTERSTATIC PORT

RH AFTSTATIC PORT

PILOT'S SYSTEM

COPILOT'S SYSTEM

PITOT SYSTEM

LEGEND

Figure 10A-11. Pitot, Static, and AOA Anti-ice System

1. Bleed air is used for anti-icing on:A. Pitot tubes and static portsB. Nacelle lipsC. Wing leading edges and engine

front frameD. B and C

2. The L and R PITOT HEAT switches alsosupply heating element power for the:A. Angle-of-attack vanesB. Shoulder static portsC. Instrument static portsD. Ice detector probe

3. The crew action required when theWING OV’HT light illuminates is:A. No action is required; the system is

automatic.B. Position the WING HEAT switch to OFF.C. Reduce power.D. Turn the BLEED AIR switch to OFF

until the light extinguishes.

4. Anti-icing equipment should be turned on:A. During descentB. Before entering icing conditionsC. Before takeoffD. During climbout

5. With the loss of airplane electricalpower, anti-icing is lost on:A. All systemsB. Pitot, static, and AOA vanes onlyC. All systems except the nacelle inlet lipsD. All systems except engine front frame

6. The DEFOG switch controls airflow to:A. Windshield heatB. Wing leading-edge heatC. Front frame heatD. Nacelle lip heat




QUESTIONS

11-i

CHAPTER 11AIR CONDITIONING

CONTENTS

Page

INTRODUCTION ................................................................................................................. 11-1

GENERAL ............................................................................................................................ 11-1

ENGINE BLEED-AIR CONDITIONING AND DISTRIBUTION ..................................... 11-2

General........................................................................................................................... 11-2

Air-Conditioning System (SNs 24-350 and 25-227and Subsequent .............................................................................................................. 11-2

Air-Conditioning System (SNs 24-230 through -349 andSNs 25-003 through -226............................................................................................... 11-6

Air-Conditioning System (SNs 23-003 through 24-229) ............................................ 11-10

AUXILIARY AIR-CONDITIONING SYSTEMS.............................................................. 11-13

General......................................................................................................................... 11-13

Distribution System ..................................................................................................... 11-13

Auxiliary Cooling System ........................................................................................... 11-14

Auxiliary Heating System (Optional) (Not available on SNs 23-003 through 24-129) .................................................................................. 11-16

QUESTIONS....................................................................................................................... 11-19




11-iii

ILLUSTRATIONS

Figure Title Page

11-1 Conditioned Bleed-Air Distribution ...................................................................... 11-2

11-2 Air-Conditioning System—SNs 24-350 and 25-227 and Subsequent................... 11-3

11-3 CAB AIR Switch ................................................................................................... 11-4

11-4 H VALVE Indicator (Current Models)................................................................... 11-4

11-5 Cabin Climate Switch Panel .................................................................................. 11-6

11-6 Air-Conditioning System—SNs 24-230 through -349and 25-003 through -226........................................................................................ 11-7

11-7 AIR BLEED Switch............................................................................................... 11-8

11-8 Temperature Control Panel .................................................................................... 11-9

11-9 H VALVE Indicator ............................................................................................. 11-10

11-10 Air-Conditioning System—SNs 23-003 through 24-229 .................................... 11-11

11-11 AIR BLEED and NORM–MAX Switches .......................................................... 11-12

11-12 Evaporator and Blower Assembly—SNs 23-003 through 24-129 ...................... 11-13

11-13 Evaporator and Blower Assembly—SNs 24-130 through -203 and 25-003 through -039 ............................................................................................ 11-13

11-14 Evaporator and Blower Assembly—SNs 24-204 and Subsequent and25-040 and Subsequent........................................................................................ 11-13

11-15 Diverter Door Control.......................................................................................... 11-14

11-16 Fan Rheostat ........................................................................................................ 11-14

11-17 COOL SYS–OFF–FAN Switch ........................................................................... 11-14

11-18 Auxiliary Cooling System ................................................................................... 11-15

11-19 Auxiliary Heat System......................................................................................... 11-16

11-20 Auxiliary Heat Switch ......................................................................................... 11-17

11-21 Heater Element Installation ................................................................................. 11-18




INTRODUCTIONAir conditioning on the Lear 20 series is provided by regulated bleed air that is temperaturecontrolled and distributed throughout the cabin cockpit. Supplemental cooling and heat-ing are provided by a refrigeration system and an optional auxiliary heating system. Thesesupplemental systems have a separate distribution network through which the air is cir-culated by a cabin blower and cockpit fan.

GENERALPrimary heating and cooling are accomplishedby controlling the temperature of the bleed airentering the cabin. The distribution systemsvary slightly based on model and serial number.

Supplemental cooling is provided by a Freonrefrigeration system that may be used on theground and in the a i r up to a maximum

altitude of 18,000 feet.The auxiliary heatingsystem, if installed, may be used on the groundand in the air at any altitude.

A typical conditioned bleed-air distributionsystem is shown in Figure 11-1.

CHAPTER 11AIR CONDITIONING




ENGINE BLEED-AIRCONDITIONING ANDDISTRIBUTION

GENERALA detailed description of the air supply is pro-vided in Chapter 9 of this manual.

Engine bleed air is supplied to the bleed-airmanifold in the tail compartment and is thenrouted through the flow control valve to theheat exchanger. The air is then routed throughthe air distribution ducts to the cabin andcockpit. There are three systems divided asfollows: airplanes SNs 24-350 and 25-227and subsequent; SNs 24-230 through -340 andSNs 25-003 through 25-226; and SNs 23-003through 24-229.

These systems differ in the manner the bleedair is cooled, how the hot air bypass valve iscontrolled, and in the switches that control theflow control valve.

AIR-CONDITIONING SYSTEM(SNs 24-350 AND 25-227 ANDSUBSEQUENT)Figure 11-2 illustrates the air-conditioningsystem.

LH and RH Bleed-Air Shutoff ValvesTwo bleed-air shutoff valves are installed ineach duct to the bleed-air manifold in the tail-cone and are controlled by the respective LHand RH AIR BLEED switches on the controlpanel. In the OFF position, all bleed air isshut off from the respective engine to thebleed-air manifold. In the ON position, thevalve is open and permits bleed air to enter thebleed-air distribution system.

Bleed-Air ManifoldThe bleed-air manifold is essential ly aninsu la ted me ta l duc t which se rves a s acollection point for bleed air from either orboth engines.




Figure 11-1. Conditioned Bleed-Air Distribution

COPILOT’S AIROUTLET

WINDSHIELD HEAT DUCT

DIFFUSER (TYPICAL)

FOOTWARMERWINDSHIELD HEAT SYSTEM

PILOT’S AIROUTLETS

LOWER CABIN DOORAIR OUTLET

CHECK VALVE




Figure 11-2. Air-Conditioning System—SNs 24-350 and 25-227 and Subsequent

BLEED AIR

RAM AIR

CONDITIONED BLEED AIR

LEGEND

SERVO AIR

C O

1/4 3/41/2

•

• • •

•

FOOTWARMEROUTLET

INTERNAL DEFOGOUTLETS

CREWOUTLETS

CABINTEMP

SENSOR

TO SENSORBLOWER

MOTOR

CABIN AIRDIFFUSERS

(TYPICAL)

CHECKVALVE

BLEEDAIR

DUCT TEMPLIMITER

DUCT TEMPSENSOR

RAMAIR IN

LHBLEED-AIR

VALVE

SERVOBLEED AIR

RAM-AIRCHECK VALVE

VENTURI

HOT AIR BYPASS(H VALVE)

BLEED-AIRMANIFOLD EMERGENCY

PRESSURIZATIONVALVE

FLOWCONTROLVALVE

RAM-AIROUT

CHECKVALVE

H VALVEIND

RHBLEED-AIR

VALVE

COLD HOT

ON

OFF

CABAIR

COOL

COLD HOT

AUTO MAN

FAN

OFF CABIN

TEMP

H-VALVEH-VALVEH-VALVE

LHMAINBUS

AIRBL

LHESSBUS

RHECS

VALVE

RHMAINBUS

LHECS

VALVE

LHMAINBUS

AIR BLEEDEMER

ONOFF

HEATEXCHANGER

From the bleed-air manifold, bleed air isdistributed to the flow control valve for pres-surization and cabin heating.

Flow Control ValveThe flow control valve is a solenoid-operatedvalve which controls and regulates the flow ofbleed air into the cabin. The position of thevalve is determined by the CAB AIR switch(Figure 11-3).

Figure 11-3. CAB AIR Switch

When the CAB AIR switch is in OFF, the valveis energized and closes. When the switch isin ON, the valve is deenergized and opens. DCpower for its operation is drawn through theAIR BLEED circuit breaker on the left es-sential bus.

Other functions of the flow control valve andthe venturi located downstream of the flowcontrol valve are related to the pressurizationsystem. Those aspects of component operationare discussed in Chapter 12, “Pressurization.”

Hot Air Bypass Valve (H-Valve)The hot air bypass valve, more commonlyreferred to as the “H-valve,” is located in thebleed-air upstream of the heat exchanger. Its

function is to split the flow of bleed air,directing some to the heat exchanger for cool-ing and some to bypass the heat exchanger.The result is a mixture of the two airflows,thereby conditioning the bleed air before itenters the cabin area. The position of the H-valve is indicated on the H VALVE indicatorlocated on the copilot’s instrument panel(Figure 11-4).

Figure 11-4. H VALVE Indicator(Current Models)

The H-valve butterfly is positioned pneumat-ically by servo bleed air (see Chapter 9) fromthe climate control system. No electrical cir-cuits are involved except that the H VALVE in-dicator requires DC power. Approximatelyfive seconds is required for the valve to travelfrom fully opened to fully closed. The valveis spring-loaded to the full cold position anytime servo air pressure is not available.

Ram-Air Heat ExchangerThe heat exchanger is located inside the tail-cone. It consists of a bleed-air core surroundedby a ram-air plenum. Cool air enters the ram-air inlet in the dorsal fin and flows through theplenum, across the bleed-air core, thus cool-ing the bleed air. The ram air then exhausts intothe tail compartment.




DN

UP

ON

OFF

CABAIR

PRESSURIZATION

The cooled bleed air flowing out the heatexchanger core is ducted back to the bypassside of the H-valve where it mixes with hotbypassed bleed air. The resulting conditionedair is then directed into the cabin and cockpitdistribution system.

When the airplane is on the ground,do not perform extended engineoperation above idle with the CABAIR and AIR BLEED swi t chespositioned to ON. Since there is noram air for cooling of the bleed air,possible damage to the air-condi-tioning components could result.Damage might also occur to interiorcabin furnishings, as well as over-heating the tailcone area.

Ram-Air VentilationIn the event that the airplane is unpressurizedin flight, air for circulation and ventilation ofthe cabin and cockpit areas is provided by ramair, which is ducted into the conditioned bleed-air distribution system.

During normal operation, a one-way checkvalve in the connecting ram-air duct preventsloss of conditioned pressurization bleed airthough the ram-air plenum exhaust port.

Cabin and CockpitAir DistributionGeneralConditioned airflow distribution to the cabinand cockpit areas is essentially the same forall airplanes (see Figure 11-2). The condi-tioned air is routed from the tailcone into thecabin area through two ducts, one on eachside of the cabin. The left duct ends at theentry door, and the r ight duct continuesforward to the cockpit.

Cabin Air DistributionCabin air distribution is furnished by diffusersinstalled at intervals along the two ducts, andthey direct airflow toward the floor.

A one-way air distribution check valve islocated at the aft end of each cabin duct. Thesevalves are functionally related to the pres-surization system, as described in Chapter 12of this manual.

On SNs 24-350 and subsequent and 25-227 andsubsequent, distribution of air changes wheneither or both emergency pressurization valvesare positioned to emergency.

If only one emergency valve is positioned toemergency, all bleed air from that engine isrouted directly into only that side’s cabin dis-tribution duct, and temperature control of thatair is lost. However, bleed air from the oppo-site engine is still subject to the normal con-ditioning process. One-way check valves in thenormal distribution ducting prevent the emer-gency airflow from being lost through the nor-mal distribution system.

If both emergency valves are positioned toemergency, all bleed air from both engines isrouted directly into the respective left andright distribution ducts. Temperature controlis then sacrificed for pressurization.

Cockpit Air DistributionCockpit air distribution is provided by duct-ing connected to the forward end of the rightcabin duct. Outlets located on the sidewallpanels and adjacent to the outboard rudderpedals enable the pilots to control and directthe airflow as desired. A footwarmer diffuser,located below the instrument panel just for-ward of the center pedestal, directs continu-ous conditioned air along the center floor.Two piccolo tubes installed vertically on eachside of the windshield center support structuredirect a continuous flow of conditioned airacross the forward section of each pilot’s wind-shield for interior windshield defogging.

CAUTION




Temperature ControlTemperature control of the engine bleed airentering the cabin area is accomplished byvarying the position of the H-valve butterfly.As the valve opens, less bleed air is directedto the heat exchanger for cooling, while morebleed air is bypassed and mixed with the cooledair. Manual and automatic operation of theH-valve is achieved by controls on the CABINCLIMATE swi tch pane l , loca ted on thecopilot’s lower instrument panel (Figure 11-5).

Figure 11-5. Cabin Climate Switch Panel

On SNs 24-350 and subsequent and 25-227 andsubsequent, the H-valve is positioned pneu-matically by servo bleed air (see Chapter 9), andno e lec t r ica l c i rcu i t s a re involved . TheAUTO–MAN knob is actually a servo bleed-airselector valve. The COLD–HOT knob is a nee-dle valve that controls the servo air pressure ap-plied to the H-valve butterfly (spring-loaded tothe full cold position). Other system compo-nents include a temperature sensor located inthe upper forward cabin, a duct temperaturesensor, and a duct temperature limiter locatedin the air duct downstream of the H-valve (seeFigure 11-2). The control system consists of aninterconnected servo bleed-air network.

With the AUTO–MAN knob in the MANposition, the selector valve isolates the con-trol system from the influences of the cabintemperature sensor and the duct temperaturesensor. Servo air pressure is routed directlythrough the needle valve (controlled by the

COLD–HOT knob) to the H-valve butterfly.Changing the COLD–HOT knob positionsimply changes the servo air pressure on theH-valve butterfly. The H VALVE indicator(see Figure 11-4) displays the relative positionof the H-valve which is the only componentin the system that requires DC electrical power.DC power is provided through the H VAL INDcircuit breaker on the left main bus.

With the AUTO–MAN knob (selector valve) inthe AUTO position, the servo pressure controlnetwork samples the needle valve setting(COLD–HOT knob position), the cabin temper-ature sensor (existing cabin temperature), andthe duct temperature sensor (actual temperatureof the bleed air inside the duct). Servo air pres-sures are modulated by the control system,which causes the H-valve butterfly to modulate,thereby keeping the cabin temperature constant.

Whether the system is being operated manu-ally or automatically, the duct temperaturelimiter signals the control unit if the ducttemperature increases to approximately 350°F.The control unit’s response is to drive theH-valve to the full cold position, and direct allbleed air through the heat exchanger.

AIR-CONDITIONING SYSTEM(SNs 24-230 THROUGH -349 ANDSNs 25-003 THROUGH -226Figure 11-6 illustrates the air-conditioningsystem.

Bleed-Air Check ValvesA bleed-air check valve is installed in thepylon bleed-air duct for each engine. Eachcheck valve is a flapper check valve allowingairflow in one direction and closes with air-flow in the opposite direction. The purpose ofthe check valves is to prevent loss of bleed airin the event of single-engine operation.

Bleed-Air ManifoldThe bleed-air manifold is a metal duct whichserves as a collection point for bleed air fromeither or both engines. From the bleed-air




AUTO MAN

COOL

COLD HOT

FAN

OFF

CABIN

TEMP




Figure 11-6. Air-Conditioning System—SNs 24-230 through -349 and 25-003 through -226

EMER CAB PRESSEXT DEFOG ON

CABIN AIR FLOWMAN COOLHOT

HOTHOT

DECR

RATE

1

2

34 5 6

789

10

INCRCABIN CLIMATECONTROL

HOT

COLD

COOLSYS

MAX

OFF

NORM

FAN

MAN

SYS

FAN

MANMAN

COLDAUTO AUTO

PRESSREGULATOR MAX

NORM

MAN

AUTOUP

OFF

DN

OFF

UP

DN

AUTO

MAN

OFF

DUMP

SAFETYVALVE

AUTO

DEFOG

OFF

AIR

BLEED

OFF

OFF

PRESSURIZATION

CAB

TEMP

CAB

TEMP CABIN X 1000 FT

ALTITUDE X 1000

*SKINTEMP

SENSOR

CABIN TEMPSENSOR

CHECKVALVE

TO WING L EANTI-ICE SYSTEM

LH ENGINE RH ENGINE

BYPASS VALVE

RAM-AIR OUT

HEAT EXCHANGER

FLOW CONTROL VALVEBLEED-AIRCHECK VALVE


DUCT TEMPSENSOR

RAM-AIRCHECK VALVE

RAM-AIRIN

CHECKVALVE

TEMPERATURECONTROL

UNIT

FOOTWARMERS ANDEXTERNAL DEFOG(BLEED AIR)

FOOTWARMERS ANDEXTERNAL DEFOG(BLEED AIR)

CABIN OUTLETS(CONDITIONED AIR)

INTERNAL DEFOG(CONDITIONED AIR)

DEFOGPRESSUREREGULATOR

DEFOG SHUTOFFVALVE

BLEED AIRCHECK VALVE

HV POS IND

RAM AIR

BLEED AIR

CONDITIONED AIR

LEGEND

*AIRPLANE SNs 24-255 AND SUBSEQUENT, 25-090, 25-095 AND SUBSEQUENT

**AIRPLANE SNs 25-003 THROUGH 25-063

CABIN CLIMATE CONTROLCABIN CLIMATE CONTROLCABIN CLIMATE CONTROL

manifold, bleed air is distributed to the flowcontrol valve for pressurization and cabinheating, to the pressurization jet pump, to thewindshield heat (defog) valve, and to the winganti-ice valve.

Flow Control ValveThe flow control valve regulates the flow ofbleed air to the cabin for normal pressuriza-tion and heating. It is controlled by a three-position AIR BLEED switch labeled “OFF,”“NORM,” and “MAX” on the pressurizationmodule (Figure 11-7). The OFF position closesthe valve completely. MAX position opensthe valve to full flow. NORM position allowssense line pressure from the venturi (locateddownstream) to modulate the airflow and main-tain a constant airflow into the cabin.

Figure 11-7. AIR BLEED Switch

Hot Air Bypass Valve (H-Valve)The bypass valve is connected between thebleed-air and precooled ducts adjacent to theheat exchanger and is operated by 28 VDC.The bypass valve either routes bleed airthrough the heat exchanger for cooling or by-passes the heat exchanger for cabin heating.By varying the combination of heated andcooled air, cabin temperature may be main-tained at a comfortable level. This valverequires 35 to 55 seconds to travel through its

entire range (full heat to full cold) and is con-trolled by the cabin heat system in both theautomatic and manual modes. A potentiome-ter and the temperature control circuit providea balancing signal. The cabin temperaturecontrol indicator (H VALVE indicator) on thecopilot’s subpanel indicates the position of heH-valve and is operated by 28 VDC.

Ram-Air Heat ExchangerThe heat exchanger is located inside he tail-cone. It consists of a bleed-air core surroundedby a ram-air plenum. Cool air enters the room-air inlet in the dorsal fin and flows through theplenum, across the bleed-air core, thus cool-ing the bleed air. The ram air then exhausts intothe tail compartment.

The cooled bleed air flowing out of the heatexchanger core is ducted back to the bypassside of the H-valve where it mixes with hot by-passed bleed air. The resulting conditionedair is then directed into the cabin and cockpitdistribution system.

When the airplane is on the ground,do not perform extended engineoperation above idle with the switchpositioned to NORM. Since there isno ram air for cooling of the bleedair, poss ible damage to the a i r-condi t ioning components couldresult. Damage might also occur tointerior cabin furnishings, as well asoverheating the tailcone area.

Ram-Air VentilationIn the event that the airplane is unpressurizedin flight, air for circulation and ventilation ofthe cabin and cockpit areas is provided by ramair, which is ducted into the conditioned bleed-air distribution system.

During normal operation, a one-way checkvalve in the connecting ram-air duct preventsloss of conditioned pressurization bleed airthrough the ram-air plenum exhaust port.

CAUTION




Cabin and CockpitAir DistributionConditioned airflow distribution to the cabinand cockpit areas is essentially the same forall airplanes (see Figure 11-1). The condi-tioned air is routed from the tailcone into thecabin area through two ducts, one on eachside of the cabin. The left duct ends at theentry door, and the right duct continues for-ward to the cockpit. A footwarmer diffuser, lo-cated below the instrument panel just forwardof the center pedestal, directs continuous con-ditioned air along the center floor. Two pic-colo tubes, installed vertically on each side ofthe windshield center support structure, directa continuous flow of conditioned air across theforward section of each pilot’s windshield forinterior windshield defogging.

Temperature ControlCabin temperature control is accomplished byconditioning the engine bleed air, which isalso used for cabin pressurization. The cabinheating system can be controlled both manu-ally and automatically. In both modes, thecabin temperature is regulated by the bypass(H) valve which controls the amount of enginebleed air that is allowed to pass through theair-to-air heat exchanger. The bypass valve isregulated either manually or automaticallythrough the cabin temperature knob locatedon the climate control. The engine bleed airfollows a route from the engine(s) to the bleed-air manifold, flow control valve, venturi, to thebypass valve, and then either through the heatexchanger to the cabin air distribution ducts ordirectly to the cabin air distribution ducts(bypassing the heat exchanger) depending onthe position of the H-valve. Rapid cabin tem-perature fluctuations are prevented by a ducttemperature sensor in the supply duct and ahigh-temperature limit switch prevents theduct temperature from exceeding safe limits.

There are usually two 28 VDC circuit break-ers—one for automatic and one for manualcabin temperature control. A cabin tempera-ture control knob (Figure 11-8) is used toselect MAN (in detent full CCW); the cabin

temperature is controlled by using the spring-loaded HOT–COLD switch to operate thebypass valve.

Figure 11-8. Temperature Control Panel

Auto-Manual SelectorThis rotary switch allows selection of eitherautomatic or manual control of cabin tempera-ture. In the AUTO range, the rotary controlfunctions as a thermostat and allows the pilotto set cabin temperature as desired. In theautomatic mode, the H-valve changes positionin response to signals from the cabin tempera-ture sensor, the duct temperature sensor, and anoutside skin temperature sensor to maintain aconstant cabin temperature. With the selectorswitch in MAN (full CCW), the H-valve may bemanually repositioned using the toggle switchat the upper left of the climate control panel.

Auto Mode OperationWith 28 VDC available through the AUTOCABIN HEAT circuit breaker, turn the cabintemperature control knob to approximatelythe 10 o’clock position (a comfortable set-ting for cruise). Automatic cabin heat is nowcontrolled automatically using the cabintemperature sensor located behind the copilot’sseat, the skin temperature sensor located on the




fuselage skin beside the copilot seat, and a ducttemperature sensor located in the tailcone.These sensors determine what the H-valve po-sition should be to produce a cabin tempera-ture equal to that selection on the cabintemperature control knob. If automatic cabinheat is not functioning properly, manual cabinheat should be used.

Manual Mode OperationWith 28 VDC available through the MANCABIN HEAT circuit breaker, turn the cabintemperature control knob full counterclock-wise into the detent, then toggle the spring-loaded HOT–COLD switch for the desiredcabin temperature.

NOTEH-valve movement may be observedby reference to the H VALVE posi-tion indicator (optional on some air-craft). See Figure 11-9.

Figure 11-9. H VALVE Indicator

Duct Temperature SensorA duct temperature sensor is installed in thecabin bleed-air duct between the H-valve andthe cabin. In automatic temperature controlmode, it monitors the temperature of the bleedair entering the cabin and provides H-valverepositioning signals. The signals compen-sate for changes in bleed-air temperature dueto changes in engine power setting. The ducttemperature sensor does not function in themanual temperature control mode.

Duct Temperature LimiterA duct temperature limiter, adjacent to theduct temperature sensor, functions in both au-tomatic and manual temperature control modesand limits the maximum temperature of thecabin bleed-air duct. Excessive duct temper-atures provide a signal voltage that causes theH-valve to drive toward the cold position.

Cabin Temperature SensorA cabin temperature sensor, located behind thecopilot’s seat, monitors cabin temperature andprovides signals to reposition the H-valve.The cabin temperature sensor is functionalonly in the automatic mode of cabin temper-ature control.

Skin Temperature SensorOn SNs 24-255 through -349 and 25-090through -226, a skin temperature sensor ismounted on the a i rplane skin below thecopilot’s position. It provides signals to repo-sition the H-valve in response to atmospherictemperature changes. The skin temperaturesensor is functional only in the automaticmode of cabin temperature control.

AIR-CONDITIONING SYSTEM(SNs 23-003 THROUGH 24-229)

Bleed-Air Shutoff Valves(SNs 23-003 through 23-029 notModified by SK-140)Figure 11-10 illustrates the air-conditioningsystem.

The bleed-air shutoff valves control the flow ofengine bleed air through the flow control valveand through the heat exchanger. These valvesare controlled by the two AIR BLEED switches.

Bleed-Air Check ValvesTwo bleed-air check valves are installed inthe engine bleed-air outlet lines to preventloss of compressor bleed air during single-engine operation.







Figure 11-10. Air-Conditioning System—SNs 23-003 through 24-229

AIRBLEED

OPEN ON

OFF

OPEN

AIRBLEED

CLOSED

MAN

MAX

MANHOT

COLD

NORM

HOT

CLOSED

AUTO

AIR

BLEED

TEMPERATURECONTROL

UNIT

FOOTWARMER ANDEXTERNAL DEFOG

(BLEED AIR)


INTERNAL DEFOG(CONDITIONED AIR)

CABINTEMP

SENSOR


CABINCHECKVALVE

RAM-AIRCHECKVALVE

CABINCHECKVALVE


HEATEXCHANGERAND DAMPER

FLOWCONTROLSOLENOIDFLOW

CONTROLVALVE

RAMOUT

VENTURI

CHECKVALVE

DEFOGSHUTOFF

VALVE

DORSALRAM

INLET

CHECKVALVE

RHENGINE

*SNs 23-003 THROUGH 23-029 EXCEPT AIRPLANES MODIFIED BY SK23-140**SNs 24-181 AND SUBSEQUENT

LHENGINE


**DUCT TEMPERATURE SENSOR

CONDITIONED AIR

RAM AIR

BLEED AIR

LEGEND

Flow Control ValveThe flow control valve admits air from thebleed-air manifold and routes it to the heat ex-changer for temperature control. The twoswitches which control the valve are labeled“AIR BLEED” and “NORM–MAX” (Figure11-11). On models 23, 24, and 24B, the AIRBLEED switch, when in ON, allows bleed airto flow through the flow control valve, throughthe heat exchanger to the venturi, and into thecabin through the distribution ducts. The ven-turi senses differential pressure which is sentback to the flow control valve and modulatesthe airflow in order to compensate for enginepower changes. The NORM–MAX airflowswitch, when positioned to MAX, opens theflow control valve full open, allowing maxi-mum airflow into the cabin.

Figure 11-11. AIR BLEED and NORM–MAX Switches

Damper ValveThe damper valve is located just above the heatexchanger in the ram-air inlet. The purpose ofthe damper valve is to regulate the amount ofram air passing through the ram-air scoopsacross the heat exchanger and overboard, therebycontrolling the temperature of the bleed air forcabin heating. The damper valve is operated bythe AUTO or MAN cabin heat control in thecockpit. This valve requires approximately 27seconds to travel from fully closed to fully open.

Heat ExchangerThe heat exchanger is a high-effectiveness,two-pass crosscounterflow plate fin unit. High-pressure bleed air is ducted into the heatexchanger and routed through the core in cross-counterflow directions. Ram air is routed overthe core channels and overboard just aft ofthe tailcone access door. This results in a sub-stantial reduction of bleed-air temperature.

Ram-Air VentilationIn the event that the aircraft is unpressurizedin flight, air for circulation and ventilation ofthe cabin and cockpit areas is provided by ramair ducted into the conditioned bleed-airdistribution system.

During the normal operation, a one-way checkvalve in the connecting ram-air duct preventsloss of conditioned pressurization bleed airthrough the ram-air plenum exhaust port.

Cabin and CockpitAir DistributionConditioned airflow distribution to the cabinand cockpit areas is essentially the same for allaircraft (see Figure 11-1). The conditioned airis routed from the tailcone into the cabin areathrough two ducts, one on each side of thecabin. The left duct ends at the entry door, andthe right duct continues forward to the cockpit.A footwarmer diffuser, located below theinstrument panel just forward of the centerpedestal, directs continuous conditioned airalong the center floor. Two piccolo tubes(SNs 24-165 and subsequent and all model25s), installed vertically on each side of thewindshield center support structure, direct acontinuous flow of conditioned air across theforward section of each pilot’s windshield forinterior windshield defogging.

Temperature ControlCabin temperature control is accomplished byconditioning the engine bleed air, which is alsoused for cabin pressurization. The cabin heat-ing system can be controlled both manually




AIRBLEED

OPEN ON

OFF

OPEN

AIRBLEED

CLOSED

MAN

MAX

MANHOT

COLD

AUTONORM

HOT

CLOSED

AIR

BLEED

and automatically. In both modes, the desiredtemperature is maintained by varying theamount of external ram air that is allowed toflow through the air-to-air heat exchanger bythe use of a damper valve located in the ram-air inlet duct. In the automatic mode, the dampervalve is controlled by the amount of signaldifference between the temperature selectorand the cabin temperature sensing elements.Rapid temperature fluctuations are preventedby a duct temperature sensor in the supply duct,and a high-temperature limit switch prevents theduct air temperature from exceeding safe limits.In the manual mode, the position of the dampervalve is controlled directly by the pilot usingthe HOT–COLD switch.

AUXILIARY AIR-CONDITIONINGSYSTEMS

GENERALSupplemental cooling and heating are pro-vided by a Freon refrigeration system and anoptional electric heating system. A cabinblower distributes the conditioned air through-out the interior. Beginning with SNs 24-204 and25-040, a cockpit cooling fan is provided forincreased air circulation in the cockpit area. Thecabin blower and cockpit fan may be usedsimply to recirculate air within the cabin andcockpit areas or, by using the auxiliary cooleror heater, to cool or to heat the recirculated air.A ground power unit or an engine-drivengenerator must be used to operate either thecooling or heating systems on the ground.

DISTRIBUTION SYSTEMThe heart of the distribution system is the evap-orator and blower assembly, installed in the cab-inet behind the pilot’s seat on SNs 23-003 through24-129 (Figure 11-12). On SNs 24-130 through-203 and 25-003 through -039, the evaporatorwith a single duct and the blower are installedon the ceiling above the baggage compartment(Figure 11-13). On SNs 24-204 and subsequent

and 25-040 and subsequent, a dual duct andblower are installed with a cockpit cooling fanmounted between the ducts (Figure 11-14).




EVAPORATORDUCT

DRAIN LINE

EVAPORATOR BLOWERASSEMBLY

COCKPITBLOWER

EVAPORATORDUCT

MANUAL DAMPERCONTROL

Figure 11-12. Evaporator and BlowerAssembly—SNs 23-003through 24-129

Figure 11-13. Evaporator and BlowerAssembly—SNs 24-130through -203 and 25-003through -039

Figure 11-14. Evaporator and BlowerAssembly—SNs 24-204and Subsequent and25-040 and Subsequent

Cabin Blower DistributionThe cabin bower assembly consists of a squirrel-cage blower(s) driven by a DC motor. Theblowers draw air from the cabin through theevaporator and discharge through adjustablelouvered vents behind the pilot on SNs 23-003through 24-129, and through the single or dualducts located on the ceiling above the baggagecompartment in later airplanes. When installed,the optional heating elements are locatedwithin these ducts.

On airplanes with dual ducts, two variable-position diverter doors are installed, one on thebottom of each duct. They are manually con-trolled by the CLOSE–OPEN knob adjacent tothe louvered grille (Figure 11-15). When rotatedin the OPEN direction, both diverter doors areraised up into the path of the blower airflow,causing some of the flow to be diverted back intothe baggage area. This results in a reduction ofthe airflow into the cabin. When the knob isrotated to its extreme in the CLOSE direction,the doors are flushed with the bottom of theducting, maximizing airflow into the cabin.

Cockpit Fan DistributionBetween the two ducts fed by the cabin blowersis another duct which encloses the axial cock-pit fan. The cockpit fan is a 28-VDC fan con-trolled by a rheostat located on the copilot’ssidewall (Figure 11-16). This fan draws air fromthe baggage compartment area. These ducts rundirectly to the overhead outlets in the cockpit.

The cabin blower is controlled by the FANposition on the COOL SYS–OFF–FAN switch

on the climate control panel (Figure 11-17). Thecockpit cooling fan is controlled by a rheostaton the copilot’s side panel.

Figure 11-16. Fan Rheostat

Figure 11-17. COOL SYS–OFF–FAN Switch

AUXILIARY COOL SYSTEMThe Freon refrigeration system is installed toprovide supplemental cooling for ground andin-flight operations and can also be used for dehumidification. See Figure 11-18 for thesystem diagram.




Figure 11-15. Diverter Door Control

COCKPITOUTLETS

TOCABIN

BLOWER

COCKPITFAN

CABINBLOWERASSEMBLY

EVAPORATOR

AIRFLOW

EXPANSION VALVE DEHYDRATOR HIGH-PRESSURELIQUID

CONDENSERCONDENSER

HIGH-PRESSUREVAPOR

LOW-PRESSUREVAPOR COMPRESSOR

PRESSURESWITCH

MOTOR

COCKPIT

AIR

COOL

FAN

OFFOFFMIN

MAXMAXOFFOFF

CONDENSER

OFF

MAXOFF

NOTE: SNs 25-343 AND SUBSEQUENT ARE EQUIPPED WITH A ROTARY COMPRESSOR

Figure 11-18. Auxiliary Cooling System

System components include the compressormotor, compressor, condenser, receiver-dehydrator, evaporator and blower, and thecontrol switch.

The compressor motor has a 3 3/4-hp ratingat 7,000 rpm. The motor requires no externalcooling and receives 28-VDC power througha 150-ampere current limiter in the tailcone todrive the compressor via a V-belt unit.

The compressor is a two-cylinder unit with apressure switch plumbed to the compressor dis-charge port. If the discharge pressure reachesapproximately 335 psi, the switch contacts openand deenergize the compressor motor. Whenthe pressure drops to 205 psi, the switch con-tacts again close, reenergizing the compressormotor. Operation of this pressure switch doesnot affect the fan operation. The refrigerantcondenser is a plate and fin unit. High-pressure,high-temperature vapor enters the condenserfrom the compressor. This vapor is then cooledby the air passing over the condenser surface andchanges to liquid. Heat from the condenser is

removed by a fan mounted on the compressormotor shaft. The receiver-dehydrator removesthe small traces of moisture that may remain inthe system. A sight glass for observing refrig-erant flow is installed in the tip of the receiver-dehydrator. If the sight glass is generally clearand performance is satisfactory, occasionalbubbles do not indicate refrigerant shortage.

The evaporator is installed to cool, dry, and,through filtration, clean the air in the cabinsection. Refrigerant enters the evaporator fromthe expansion valve as a low-pressure mixtureof liquid and vapor. The liquid vaporizes at thislow pressure, absorbing large quantities ofheat from the air that passes through theevaporator fins, thus cooling the air beingrecirculated by the cabin blower. As heat istransferred through the walls of the evapora-tor from warm air passing over them, moisturein the air condenses and is drained overboard.

The Freon system is manually controlled withthe COOL SYS–OFF–FAN switch (Figure11-17) located on the climate control panel.




TO COCKPIT

TOCABIN

AUXILIARYHEATERELEMENTS

BLOWER

BATTERYCHARGE

BUS

150 A

AUX HEATERCONTROL BOX

LHMAINBUS

AUXCAB HT

AUX HT

OFF

HI

LO

OperationThe Freon system may be operated on theground or in flight below 18,000 feet (15,000feet on Lear 23 airplanes). When the controlswitch is placed in the COOL SYS position,28 VDC is applied to the compressor motor andthe cabin blower through control relays. Theserelays allow motor operation only if an airplanegenerator or a GPU is supplying electricalpower and neither START–GEN switch is inthe START position.

The control switch should be in the OFF po-sition during engine start to prevent over-loading th 275-ampere battery charging buscurrent limiters when the START–GEN switchis placed to GEN. It is recommended to waituntil the generator load is below 150 amperesper generator before turning the Freon systemswitch to the COOL position.

NOTEFor maximum effectiveness on theground, the AIR BLEED or CAB AIRswitch should be turned off when theengines are running.

Manually operated diverter doors installed in theair ducts from the evaporator may be positionedto direct some of the cool air into the baggagecompartment.

When the control switch is placed in the FANposition, 28 VDC is applied to the cabin fanwhich circulates only the cabin air. The fanmay be used at any altitude.

AUXILIARY HEATING SYSTEM(OPTIONAL) (NOT AVAILABLEON SNs 23-003 THROUGH 24-129)The auxiliary cabin heat system (Figure 11-19)may be used in flight or on the ground and mustbe powered by either a GPU or an engine gen-erator. The system consists of two heater circuitswith a three-position switch (Figure 11-20) onthe copilot’s subpanel. Each heater circuit con-sists of two heater elements wired in parallelusing the cabin fan for air circulation. Theheater elements are located in the evaporatorassembly above the baggage compartment(Figure 11-21).




Figure 11-19. Auxiliary Heat System

OperationElectrical power for system operation must besupplied by either a GPU or an engine-drivengenerator. The system is operated by the AUXHEAT switch located on the copilot’s lowerright instrument panel (Figure 11-20). Providedthe diverter doors are fully closed (safety switchactuated), selecting HIGH on the switch powersboth heating coils in each element. SelectingLOW on the switch powers only one coil perelement. The system is powered from the batterycharging bus through the 150-ampere air-conditioning current limiter and a control relayfed from the AUX CAB HEAT circuit breaker.

When either the LOW or HIGH selection ismade, the cabin blower motor is energized, buton SNs 24-271 and 25-126 and subsequent, fanspeed is initially limited to 1/10 the normalspeed until one of the thermoswitches sensesthe high limit. At that time, the blower motorbegins to run continuously at its maximumspeed. Power to the heating elements thencycles off and on between the low and high

l imi ts as sensed by the respect ive ther-moswitches, maintaining a constant airflowtemperature.

Should either thermal switch fail to shut offpower to the affected element at the high limit,the thermal fuse melts through and discon-nects the affected element, precluding elementoverheat.

Circulation of air is described in Cabin BlowerDistribution. When the auxiliary heater is inoperation, the heater elements are disconnectedby the safety switch if, for any reason, thediverter doors are opened. In this case, the cabinblower soon begins to blow progressively coolerair until the condition is recognized.

The cockpit fan can be used concurrently withthe auxiliary heater system, but the air it cir-culates does not pass across the heater coils.Therefore, warm air from the auxiliary heateris not available.

When the auxiliary heater is beingpowered by a GPU only, it is possiblein some conditions for the aircraftbatteries to be depleted if GPU fail-ure occurs.

Power to the heaters is NOT auto-matically disconnected when theSTART–GEN switch is positionedto START. Normal operating proce-dures require that the AUX HEATswitch be in the OFF position priorto engine start to preclude possibleelectrical system damage.

CAUTION

CAUTION




Figure 11-20. Auxiliary Heat Switch




Figure 11-21. Heater Element Installation

AIR FLOW

AIR FLOW

AIR FLOW

EVAPORATOR ANDBLOWER ASSEMBLY

EVAPORATOR ANDBLOWER ASSEMBLYTHERMAL

FUSEHEATING

ELEMENTS

THERMALFUSE

HEATERCOILS

DIVERTER DOORCONTROL CABLE

CABIN BLOWERDUCT

HEATERELEMENTS

THERMOSWITCH(HIGH LIMIT — 150°F)(LOW LIMIT — 125°F)

HEATERSAFETY SWITCH

DIVERTER DOORBELLCRANK SWITCH

ROLLER DIVERTER DOOR

THERMOSWITCH

AIRPLANES WITH SINGLE EVAPORATOR DUCT–SNs 25-003 THROUGH 25-039 ANDSNs 24-130 THROUGH 24-203

EVAPORATOR ANDBLOWER ASSEMBLY

AUXILIARY CABIN HEATERSAFETY SWITCH

AIRPLANES WITH DUAL EVAPORATOR DUCTS—SNs 24-204 AND SUBSEQUENT ANDSNs 25-040 AND SUBSEQUENT

QUESTIONS




1. The diverter doors must be fully closed to:A. Operate the cockpit fanB. Operate the Freon systemC. Operate the auxiliary heating systemD. Prevent loss of bleed-air conditioning

2. The valve used to control cabin temper-ature is the:A. Flow control valveB. Damper valve (SNs 23-003 through

24-229)C. Bypass H-valveD. Outflow valve

3. When the aircraft is unpressurized on the ground, air circulation isprovided by:A. Ram airB. The cockpit fan and the cabin

blowerC. Bleed-air systemD. Auxiliary defog system

4. The primary air conditioning in flightis provided by:A. Engine bleed airB. The heat pumpC. The auxiliary heaterD. The Freon refrigeration system

5. When using the auxiliary heater, theheated air blows out through the:A. Conditioned air outletsB. Louvered grille above the divan seatC. Overhead cockpit air outletsD. Overhead passenger outlets

6. The Freon system should not be used above:A. 5,000 feetB. 8,000 feetC. 10,000 feetD. 18,000 feet (15,000 feet, model 23)

7. The Freon system automaticallydisengages:A. During engine startB. Upon touchdownC. When unpressurizedD. If the main door is opened

8. When the Freon system is operating,it cools:A. Ram airB. Cabin airC. Outside airD. Bleed air

9. When operating the Freon system onthe ground with engines running, theswitch(es) that should be in OFF formaximum cooling effectiveness is the:A. START-GENB. CABIN BLOWERC. CAB AIR or AIR BLEEDD. COCKPIT AIR

10. In order to operate the auxiliary heateror Freon system:A. The engines cannot be running.B. The CAB AIR switch must be off.C. Either a GPU or an engine-driven

generator is required.D. The airplane must be on the ground.

12-i

CHAPTER 12PRESSURIZATION

CONTENTS

Page

INTRODUCTION ................................................................................................................. 12-1

GENERAL ............................................................................................................................ 12-1

SYSTEM DESCRIPTION .................................................................................................... 12-3

Components ................................................................................................................... 12-3

Indicators ....................................................................................................................... 12-8

Normal Operation .......................................................................................................... 12-9

Flight Operation............................................................................................................. 12-9

EMERGENCY PRESSURIZATION.................................................................................. 12-10

QUESTIONS....................................................................................................................... 12-13




12-iii

ILLUSTRATIONS

Figure Title Page

12-1 Pressurization Control System Diagram—SNs 24-350and Subsequent, and 25-227 and Subsequent ........................................................ 12-2

12-2 Pressurization Control Panel .................................................................................. 12-3

12-3 Pressurization Control System Diagram—SNs 24-230 through -349and 25-070 through -226........................................................................................ 12-4

12-4 UP–DN Lever......................................................................................................... 12-5

12-5 HORN SILENCE Switch....................................................................................... 12-6

12-6 AIR BLEED Switches............................................................................................ 12-8

12-7 Pressurization Indicators ........................................................................................ 12-8

12-8 IN NORMAL/OUT DE-FOG Knob .................................................................... 12-10

TABLES

Table Title Page

12-1 Automatic Protection and Warning Features—SNs 24-230through -349, and 25-067 through -226............................................................... 12-11

12-2 Automatic Protection and Warning Features—SNs 24-350and Subsequent, and 25-227 and Subsequent ...................................................... 12-12




INTRODUCTIONThe Learjet 20 series airplanes are equipped with a pressurization system that maintains a spec-ified level of cabin pressure consistent with built-in limits. Cabin pressure is regulated by con-trolling the outflow of conditioned bleed air supplied from the engines. During normal operation,the system functions automatically to provide crew and passenger comfort within the operationalenvelope of the airplane. This chapter explains how the pressurization system is regulated.

The rate of air exhausted from the cabin iscontrolled by an outflow valve which is pneu-matically operated to maintain the proper dif-ferential between cabin and ambient pressures.Inward and outward relief for both negative andexcess positive differential conditions are in-corporated to protect the aircraft structure.Thepressurization control system is completelypneumatic during normal in-flight automaticoperation. Pneumatic pressure is provided by

a vacuum jet pump. Control pressure (vacuum)is applied to the outflow valve through thepressurization control module. The pressur-ization controller provides for both automaticand manual operation. Electrically actuatedsolenoid valves and switches are incorporatedfor ground and manual operation.

CHAPTER 12PRESSURIZATION




GENERAL




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During climbs and descents, the controllerregulates the outflow discharge rate. This ratecontrol is necessary to maintain a cabin changerate that is comfortable regardless of the air-craft rate of climb or descent.

SYSTEM DESCRIPTION

COMPONENTSThe pressurization control system (Figures12-1 and 12-3) incorporates a cabin outflowvalve, vacuum jet pump and regulator assem-bly, pressurization control module, and a cabinsafety valve.

Cabin Outflow ValveThe outflow valve is located on the forwardpressure bulkhead in front of the copilot’s po-sition. Excessive cabin air pressure is meteredinto the unpressurized nose section through theoutflow valve as necessary to maintain thedesired cabin pressure.

Vacuum Jet Pump or VacuumRegulator AssemblyThe pneumatic pressure source for control ofthe outflow valve is established by a vacuumjet pump and regulator assembly in the tailconesection of the airplane. Engine bleed air isrouted through a venturi (jet pump) whichgenerates a negative pressure, while a regulatorensures that the negative pressure maintainsa constant differential pressure with respect tocabin pressure. This negative pressure (vac-uum) is furnished to the pressurization con-trol module which uses it to control the outflowvalve.

Pressurization Control ModuleThe pressurization control module is locatedon the copilot’s lower instrument panel. Thecontrols on the front of the module are locatedon the pressurization control panel. Figure12-2 shows typical aircraft pressurizationcontrol panel configurations.

AUTO–MAN SwitchPressurization control is normally accom-plished in the automatic mode. With theAUTO–MAN switch in the AUTO position, thecabin controller automatically adjusts thepneumatic pressure sent to the outflow valve,thereby regulating cabin pressure. If there isa malfunction in the cabin controller, the au-tomatic pneumatic circuit can be isolated fromthe outflow valve by selecting MAN position.The outflow valve is then manually controlledwith the manual cabin altitude control valve.




Figure 12-2. Pressurization Control Panel




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2426

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6

CABIN CONTROLLERIn the automatic mode of operation, the CABINCONTROLLER regulates cabin pressure inrelation to the altitude that is set on the alti-tude selector knob. Rotating the knob on theface of the CABIN CONTROLLER eitherturns a dial or aligns a window to indicate twoscales with a fixed index between them. Theouter scale represents cabin altitude, and theinner scale represents aircraft altitude.

Rate ControlA RATE knob is located to the lower left of theCABIN CONTROLLER to control the rate atwhich the cabin climbs and descends. TheRATE control knob allows variable controlw i th in t he app rox ima te l im i t s o f 175feet/minute and 2,500 feet/minute. In auto-matic mode, the CABIN CONTROLLERmaintains the desired rate of climb or descentuntil the selected altitude is attained.

Manual Cabin Altitude Control ValveA lever labeled UP–DN (Figure 12-4) can beused to pneumatically control the outflow valve.Because of the red knob on the end of the lever,it is frequently referred to as the “cherry picker.”

The lever is spring-loaded to the center posi-tion and is guarded on later airplanes to pre-vent inadvertent activation.

The lever can be used to increase or decreasecabin altitude in either the automatic or man-ual mode. However, if it is used in the automaticmode, the CABIN CONTROLLER also attemptsto control the outflow valve, and, as soon as thelever is released, the CABIN CONTROLLERreturns cabin pressure to the original value.

Differential Pressure-ReliefValve (Primary)The primary differential pressure-relief valvefunctions in association with the CABIN CON-TROLLER. Its purpose is to prevent exceed-ing normal differential limits.

On SNs 24-230 through -349, and 25-061 and-070 through -226, the relief valve is set for 8.9psi. For later SNs, the valve is set for 9.4 psi.

The primary differential pressure-relief valveis not functional in the manual control mode.During a rapid aircraft climb, with a low set-ting on the RATE knob, it is possible to reachthe differential pressure-relief setting priorto attaining the selected aircraft altitude.

Cabin Altitude Limiter(Outflow Valve)A cabin altitude limiter is installed on SNs 24-218, -350 and subsequent, 25-227 and subse-quent, and earlier SNs incorporating AMK78-5. The limiter functions to restrict the lossof cabin pressure due to a malfunctioning con-troller or inadvertent operation of the primarydifferential pressure-relief valve.

If cabin altitude reaches 11,000 ±1,000 feet onSNs prior to 24-350 and 25-227, or 11,500±1,500 feet on later SNs, the altitude limitermodulates the outflow valve to restrict thecabin altitude to the listed level.




Figure 12-4. UP–DN Lever

Controller Solenoid ValvesThree solenoid-operated valves are installed inthe controller and are used to control the rout-ing of pneumatic control pressure to the out-flow valve. All three valves are energized on theground, causing the outflow valve to open,thereby depressurizing the cabin.

One of the valves is used in flight to effect man-ual control of the outflow valve and is referredto as the “manual mode solenoid valve.”

For normal automatic in-flight operation, allthree solenoid valves are deenergized.

Aneroid SwitchesTwo aneroid switches are installed in the pres-surization system—one for operation of themanual solenoid and another for operation ofthe horn.

Manual Aneroid SwitchThe pressurization aneroid is located inside thepressurization module.

On SNs 24-230 through -349 and 25-061through -226, if cabin altitude increases to10,000 feet or above, the aneroid switch com-pletes a power circuit to the normally openmanual control solenoid valve. The solenoidvalve is energized closed, isolating all auto-matic pneumatic circuits from the outflowvalve. The outflow valve, when isolated, main-tains the last attained position. When cabin al-titude decreases to 8,000 feet or below, theaneroid resets and deenergizes the solenoidvalve open provided the AUTO–MAN switchis in AUTO.

On SNs 24-350 and subsequent, and 25-227and subsequent, the operation is the same asearly SNs except that the aneroid switch ac-tuates at 8,750 ±250 feet and resets at 7,200feet. Also, when actuated, it causes the amber

CAB ALT annunciator (see “AnnunciatorPanel” section) to illuminate. When the aneroidresets, the annunciator extinguishes.

Should the above cabin altitudes be reachedor exceeded, the “cherry picker” is the onlyway to control the outflow valve.

Cabin Altitude WarningHorn Aneroid SwitchThe cabin altitude warning horn aneroid switchactivates the cabin altitude warning horn when-ever the cabin altitude exceeds approximately10,000 feet. A spring-loaded HORN SILENCEswitch (Figure 12-5) on the test panel forwardof the thrust levers may be used to silence thehorn for periods of 30 to 50 seconds. The horn

continues to reactivate after each use of theHORN SILENCE switch until the aneroid re-sets at approximately 8,000 feet cabin altitude.

The CAB ALT TEST switch on the pedestaltest panel is used to test the warning horn. Ifthe CAB ALT warning light is installed, itdoes not illuminate during this test.




Figure 12-5. HORN SILENCE Switch

Cabin Safety ValveA cabin safety valve is installed in the aftpressure bulkhead. Its purpose is to relieve acabin overpressure or a negative pressure dif-ferential caused by a malfunction in the nor-mal control system. In flight, it normallyremains fully closed unless the secondary dif-ferential pressure-relief valve opens it to re-lieve an overpressure condition. In the case ofa negative differential pressure condition, am-bient pressure opens the safety valve, allow-ing an inward flow to equalize the cabin andambient pressures.

Operation of the safety valve is automatic inflight; there is no crew control. On SNs 24-350and subsequent, and 25-227 and subsequent,a fourth solenoid valve is installed in the pneu-matic control circuit to allow control of thesafety valve on the ground with the enginerunning and the AIR BLEED switches in ON.The solenoid valve is powered open when theCAB AIR switch is turned to OFF and opensthe safety valve. The valve closes 10 secondsafter the CAB AIR switch is turned ON andcloses the safety valve. The solenoid is deen-ergized in flight regardless of CAB AIR switchposition.

On airplanes prior to SNs 24-350 and 25-227,the safety valve does not open on the ground.

Differential Pressure-ReliefValve (Secondary)The secondary pressure-relief valve functionsin association with the safety valve. Should theprimary pressure-relief valve not operate prop-erly, the secondary pressure-relief valve opensthe safety valve to limit cabin pressure. Thesafety valve limits differential pressure to 9.2psi on SNs 24-230 through -349, and 25-067through -226. On subsequent SNs, the pressureis limited to 9.7 psi.

Cabin Altitude Limiter(Secondary)The cabin altitude limiter for the cabin safetyvalve serves the same purpose as the cabin al-titude limiter for the outflow valve. If the sec-ondary differential pressure-rel ief valvemalfunctions, causing the safety valve to open,the cabin altitude climbs. Should cabin altitudeclimb to 11,000 ±1,000 feet on SNs 24-349 and25-226 and prior, or 11,500 ±1,500 feet onlater airplanes, the cabin altitude limiter in-troduces cabin air pressure into the safetyvalve. This causes the valve to modulate andmaintain cabin altitude at the limiter value.

CAB AIR Switch (SNs 24-350and Subsequent, and 25-227and Subsequent)The CAB AIR switch (Figure 12-6) controlsthe flow control valve. Switch positions are la-beled “ON” and “OFF.” With the switch inON, bleed air flows through the flow controlvalve and into the air distribution system. Inthe OFF position, the flow control valve closes,blocking airflow into the cabin. The switch op-erates in conjunction with the three-position(ON–OFF–EMER) AIR BLEED switches. TheAIR BLEED switches control bleed-air flowfrom the engines to the bleed-air manifold.

AIR BLEED Switch (SNs 24-349and Earlier, and 25-226 andEarlier)The three-position AIR BLEED switch (Figure12-6) is labeled “OFF–NORM–MAX.” TheOFF position closes the flow control valve. TheMAX position opens the valve completely.The NORM position permits the flow controlvalve to open and modulate to maintain a con-stant airflow regardless of altitude or enginepower setting.




INDICATORS

Cabin Altitude and DifferentialPressure IndicatorCabin altitude and differential pressure areindicated on a single indicator incorporatingtwo scales and two pointers (Figure 12-7).

The cabin altitude is indicated by a singlepointer and circular scale on the outer edge withCABIN ALT markings from 0 to 50,000 feet.

The cabin differential pressure is indicatedby a circular scale on the inner portion of theindicator and a single pointer. The scale rep-resents differential from 0 to 10 psi. It has agreen band from 0 to 8.9 psi on SNs prior to24-350 and 25-227, and from 0 to 9.4 psi onlater SNs. It has an amber band from 8.9 to 9.2psi on early aircraft and 9.4 to 9.7 psi on cur-rent aircraft. A red band appears above 9.2 psion early aircraft and above 9.7 on current air-craft. Cabin altitude should always be equalto or less than the aircraft altitude.

Cabin Vertical Speed IndicatorThe cabin vertical speed indicator is positionedto the right of the cabin altimeter. It providesan indication of cabin climb or descent ratesof between 0 and 6,000 feet per minute.




Figure 12-7. Pressurization Indicators

MAXNORMOFF

AIRBLEED

INSTRUMENT PANELLOCATOR

ON

AIR BLEEDEMER

ONOFF

OFF

24-230 THROUGH 24-34925-070 THROUGH 25-226


CABAIR

Figure 12-6. AIR BLEED Switches

NORMAL OPERATION

Before TakeoffDuring ground operation, the CAB AIR orAIR BLEED switch is normally not turned toON until just prior to takeoff unless enginebleed air is desired for cabin heating.

When accomplishing the Before StartingEngines checklist from the approved AFM,the crew normally sets the AUTO–MAN switchto AUTO, the expected cruise altitude on theACFT (inner) scale of the CABIN CON-TROLLER dial, and the RATE knob to ap-proximately the nine o’clock position.

When the CAB AIR or AIR BLEED switch ispositioned on prior to takeoff, the flow con-trol valve is opened, allowing engine bleed airto enter the cabin. On SNs 24-350 and subse-quent, and 25-227 and subsequent, there is adelay of approximately 10 seconds before thesafety valve closes.

FLIGHT OPERATION

AutomaticAt liftoff, the squat switch relay box deener-gizes all pneumatic solenoids and pressur-ization begins. The cabin altitude begins toclimb at a rate based on the RATE knob set-ting. It should be adjusted, as necessary, tomaintain a comfortable cabin altitude climbrate of approximately 600 feet per minute. Asthe aircraft climbs to cruise altitude, theCABIN CONTROLLER automatically adjuststhe outflow valve to desired cabin climb rateuntil the altitude set on the controller dial isreached. As the airplane continues to climb,the differential pressure increases while thecabin altitude remains constant until the air-craft arrives at the selected aircraft altitude.If it is observed that the DIFF PRESS indica-tor is riding on the yellow/red line, a slightlyhigher cabin altitude should be selected. Adjustthe CABIN CONTROLLER as necessary whenchanging cruise altitude.

Monitor cabin altitude and differential pres-sure throughout the flight.

ManualIf the CABIN CONTROLLER is not function-ing properly, follow the Manual Mode Operationprocedures in Section 2 of the approved AFM.

Manual mode operation is established when theAUTO–MAN switch is placed to MAN. Thiscloses the manual mode solenoid valve, block-ing the automatic pneumatic circuit.

The UP–DN lever (cherry picker) then controlsthe outflow valve directly by using the staticair source or existing cabin pressure to causethe cabin to climb or descend.

The manual control valve can be very sensi-tive, and even small, momentary displace-ments of the lever generates significant cabinclimb or descent rates.

In manual mode, the cabin altitude must bemonitored much more closely than in auto-matic mode. The outflow valve position mustbe adjusted frequently during climbs and de-scents and when making power adjustments.

DescentDuring descent for landing, destination fieldelevation should be set on the CABIN scale ofthe CABIN CONTROLLER dial. The airplanerate of descent should be controlled so thecabin descent rate is comfortable (approxi-mately 600 feet per minute).

LandingAs the aircraft descends and reaches the pre-selected cabin altitude, the outflow valve mod-ulates toward the open position. The cabinshould be unpressurized at landing. At touch-down, the squat switch relay box actuatesthree pneumatic solenoid valves in the con-troller, causing the outflow valve to open com-pletely and ensure cabin depressurization. Inaddition, when the CAB AIR or AIR BLEEDSWITCH is placed to OFF, the flow control




valve closes, and on SNs 24-350 and subse-quent, and 25-227 and subsequent, the safetyvalve simultaneously opens.

EMERGENCYPRESSURIZATIONAn emergency source of pressurization bleedair is provided to increase the flow of air intothe cabin in the event of a pressure leak.

SNs 24-230 through -349, and25-067 through -226Emergency pressurization is provided by useof the windshield ant i - ice/defog system(Chapter 10). This is accomplished by pushingthe IN NORMAL/OUT DE-FOG knob (Figure12-8) in, then positioning the WSHLD HEATAUTO–MAN switch to AUTO. This causes thedefog valve to fully open and also illuminatesthe WSHLD HEAT light. These actions intro-duce air directly into the cabin area through the

pilot’s footwarmer and bypass possible leaksin the conditioned bleed-air distribution system.To isolate bleed-air leaks downstream of theflow control valve, it is necessary to place theAIR BLEED switch to OFF.

SNs 24-350 and Subsequent,and 25-227 and SubsequentEmergency pressurization is accomplished byrouting air directly into the cabin from eitheror both engines through the emergency pres-surization valves. This air completely by-passes the entire manifold and conditionedbleed-air distribution system (see Chapter 9).

The valves are spring-loaded to the emergencyposition and require both servo bleed-air pres-sure and DC power to position them to normal.Cockpit control of the valves is provided bythe three-position (OFF–ON–EMER) AIRBLEED switches (on some airplanes, theswi t ches a r e l abe l ed “BLEED-AIR”) .Automatic functioning occurs as a result of ex-cessive cabin altitudes or DC power failure.

With the AIR BLEED switches in ON, a solenoidon each emergency valve is energized, causingservo bleed-air pressure to move the valve to thenormal position.

Positioning either AIR BLEED switch to EMERdeenergizes the respective solenoid, blockingthe servo bleed-air pressure, and the valve repo-sitions to emergency by spring pressure.

The emergency pressurization valves are alsocontrolled by two cabin aneroid switches (onefor each valve). The aneroids are set to oper-ate at 9,500 feet ±250 feet cabin altitude.Should cabin depressurization occur for anyreason, the aneroid switches deenergize theirrespective valve solenoids, causing the valvesto position to emergency. The aneroids resetwhen cabin altitude decreases to approxi-mately 8,300 feet.

Finally, the same sequence of events occurs as aresult of DC power failure. On unmodified air-planes, the L and R ECS VAL circuit breakers, lo-cated on the left and right main buses, provide DCpower to hold the respective emergency valves inthe normal position.




Figure 12-8. IN NORMAL/OUT DE-FOG Knob

With an ECS VAL circuit breaker open, theemergency pressurization valve will positionto emergency and the bleed-air shutoff valvewill fail open. In this case, positioning theBLEED AIR switch to OFF will not stop air-flow into the cabin since DC electrical poweris required to close the bleed-air shutoff valve.

On airplanes modified by AMK 90-3, theemergency pressurization valves are poweredby the L and R EMER PRESS circuit break-ers on the left and right main DC buses. Onthese airplanes, the bleed-air shutoff valves arepowered by separate circuit breakers labeledL and R BLEED AIR, also located on the leftand right main DC buses. With an EMERPRESS circuit breaker open, the emergencypressurization valve will position to emer-gency and the bleed-air shutoff valve will re-main open. In this case, positioning the BLEEDAIR switch to OFF will stop airflow into thecabin since DC elecrical power, from theBLEED AIR circuit breaker, will be availableto close the bleed-air shutoff valve.

If the emergency valves are positioned to emer-gency for any reason, they may be reset by cy-cling the AIR BLEED switches to OFF, thenback to ON if power is available, and the cabinaltitude is at or below 7,200 feet.

During the first engine start, the valves auto-matically shift from emergency to normal asservo air pressure from the engine becomesavailable. A slight rush of air into the cabin isnormal during start.

The pressurization system automatic protectionand warning features relative to airplane serialnumbers are shown in Tables 12-1 and 12-2.




CABIN ALTITUDE

10,000 ±250 feet

11,000 ±1,000 feet

14,000 ±750 feet

PROTECTION AND WARNING

• Pressurization aneroid automatically switches the system to manual control.

• Cabin altitude warning horn sounds—initiate emergency descent.

• Cabin altitude limiters actuate.

• Passenger oxygen masks are deployed. (See Chapter 17.)• Cabin overhead lights illuminate.*

*24-255 and subsequent, 25-090,-094 and subsequent and earlier airplanes modified by AAK 73-6.

Table 12-1. AUTOMATIC PROTECTION AND WARNING FEATURES—SNs 24-230 THROUGH -349, AND 25-067 THROUGH -226




CABIN ALTITUDE

8,750 ±250 feet

11,500 ±1,500 feet

14,000 ±750 feet

PROTECTION AND WARNING

• Pressurization aneroid automatically switches to manual control.

• CABIN ALT caution light illuminates.

9,500 ±250 feet • Emergency pressurization valves are activated by aneroid switches, directing engine bleed air directly into the cabin.

• Cabin altitude limiters actuate.

10,100 ±250 feet • Cabin altitude warning horn sounds—initiate emergency descent.

• Passenger oxygen masks are deployed. (See Chapter 17.)

• Cabin overhead lights illuminate.*

*24-255 and subsequent, 25-090, -094 and subsequent, and earlier airplanes modified by AAK 73-6.

Table 12-2. AUTOMATIC PROTECTION AND WARNING FEATURES—SNs 24-350 ANDSUBSEQUENT, AND 25-227 AND SUBSEQUENT

1. To regulate cabin pressure, the CABINCONTROLLER modulates the:A. Cabin safety valveB. Flow control valveC. Outflow valveD. Primary differential pressure-relief

valve

2. If installed, illumination of the CABALT light indicates:A. Cabin altitude is at or above 8,750

±250 feet and the pressurizationcontrol system is in manual mode.

B. Cabin altitude is at or above 8,750±250 feet and the pressurizationcontrol system is in either auto-matic or manual mode.

C. Cabin altitude is at or above 8,750±250 feet and the emergency pres-surization mode is activated.

D. The CAB AIR switch is in the OFF position.

3. In the event of airplane electrical failure:A. Cabin pressurization must be con-

trolled manually with UP–DN knob.B. Cabin pressurization is controlled

in the automatic mode only.C. The emergency pressurization

valves (if installed) automaticallyactuate to provide cabin pressure.

D. Both B and C

4. The cabin altitude warning horn soundswhen the cabin altitude reaches:A. 8,750 ±250 feetB. 9,500 ±250 feetC. 10,100 ±250 feetD. 11,500 ±1,500 feet

5. To dump residual cabin pressure ontouchdown:A. The outflow valve opens automati-

cally.B. The cabin safety valve opens auto-

matically.C. The flow control valve closes auto-

matically.D. The bleed-air shutoff valves close

automatically.

6. On airplanes without the emergencypressurization valves, if DC power fails:A. The windshield anti-ice/defog system

can be used if pressurization fails.B. There is no emergency pressuriza-

tion capability.C. The flow control valve fails closed.D. The bleed-air shutoff and regulator

valves fail closed.

7. On your airplane, if DC power fails:A. Pressurization control reverts to

manual control.B. Pressurization control reverts to

automatic control.C. Cabin pressure is not controlled.D. The cabin slowly depressurizes.




QUESTIONS

12A-i

CHAPTER 12APRESSURIZATION

CONTENTS

Page

INTRODUCTION .............................................................................................................. 12A-1

GENERAL.......................................................................................................................... 12A-1

SYSTEM DESCRIPTION ................................................................................................. 12A-1

Components ................................................................................................................ 12A-1

Indicators .................................................................................................................. 12A-10

Operation .................................................................................................................. 12A-11

EMERGENCY PRESSURIZATION............................................................................... 12A-12

QUESTIONS.................................................................................................................... 12A-13




12A-iii

ILLUSTRATIONS

Figure Title Page

12A-1 Pressurization System—SNs 23-003 through 24-139......................................... 12A-2

12A-2 Pressurization System—SNs 24-140 through -169 and25-003 through -010............................................................................................ 12A-3

12A-3 Pressurization System—SNs 24-170 through 24B-184, -186, -187,and -189, and 25-011 through -026 and 25-028.................................................. 12A-4

12A-4 Pressurization System—SNs 24B-185, -188 and -190through -209, and 25-027 and -029 through -045 ............................................... 12A-5

12A-5 Pressurization System—SNs 24B-210 through -229and 25-046 through -063, except 25-061 ............................................................ 12A-6

12A-6 Pressurization Control Module............................................................................ 12A-7

12A-7 Outflow Valve (Typical) ...................................................................................... 12A-8

12A-8 Cabin Safety Dump Switch ................................................................................. 12A-9

12A-9 Cabin Safety Valve (Typical) .............................................................................. 12A-9

12A-10 Cabin Altitude and Differential Pressure Indicator........................................... 12A-10

12A-11 Cabin Vertical Speed Indicator ......................................................................... 12A-10




INTRODUCTIONThe Learjet models 23, 24, 24B, and 25 are equipped with a pressurization system thatmaintains cabin pressure within comfortable limits. At maximum cabin pressure dif-ferential, the system maintains sea level pressure between 21,000 and 26,000 feet andan 8,000-foot cabin altitude at 45,000 feet. Cabin pressure is maintained by controllingthe amount of air being exhausted through an outflow valve. Provision is made to pre-vent over- or underpressurization.

GENERALCabin pressurization is provided by bleed airtaken from the eighth stage of the engine com-pressor. The air enters the cabin through the nor-mal air distribution ducts. The level of cabinpressure is controlled by regulating the dis-charge of air overboard through the outflowvalve. Control may be either manual or auto-matic. For automatic pressurization control,115 VAC is required.

SYSTEM DESCRIPTIONCOMPONENTSThe pressurization system (Figures 12A-1through 12A-5) consists of an outflow valve onthe forward pressure bulkhead, a cabin safetyvalve on the aft pressure bulkhead, a pressur-ization jet pump (on some airplanes), a pres-surization control module on the copilot’sinstrument panel, and a mechanical pressure re-lief valve (in some airplanes) located on the for-ward bulkhead between the pilot’s rudder pedals.

CHAPTER 12APRESSURIZATION







CABIN ALTCONTROL

NORMAL OUTFLOW INTONOSE COMPARTMENTCABIN PRESSURE

RELIEF VALVE

CABIN AIR EXHAUSTCONTROL VALVE

* CABINUNDERPRESS

SWITCH

CABINPRESS

GEAR CB

CABINDUMP

CABINDUMP

GEAR CONTROLSWITCH

CABINSAFETYVALVE

ANEROIDSWITCH

LH MAIN GEARSQUAT SWITCH

SAFETY SWITCH

SAFETYSWITCH

CABINSAFETYVALVE

ANEROIDSWITCH

115 VAC

28 VDC

UP

DOWNDOWNDOWN

GND

DUMP

FILTER

UPALT

DNALT

NORMALSTATIC SOURCE

ALTERNATESTATIC SOURCE

CABINALTITUDEEMERGENCYCONTROLVALVE

AUTOMANUAL

CABINPRESSSWITCH

LH MAIN GEARSQUAT SWITCHAIRBORNE

GNDAUTODUMP

SNs 24-100 THROUGH 24-103

BELOW8,000 FT

ABOVEABOVE10,000 FT10,000 FT

ABOVE10,000 FT

AIR

BELOW8,000 FT

ABOVE10,000 FT

*SNs 23-003 THROUGH 24-138 MODIFIED BY SK 294

NORMALSTATIC SOURCE

DUMP INTOTAILCONE

CABIN SAFETYVALVE

STATIC PRESSURE

CABIN PRESSURE

CONTROL PRESSURE

LEGEND

AUTO

Figure 12A-1. Pressurization System — SNs 23-003 through 24-139




Figure 12A-2. Pressurization System — SNs 24-140 through -169 and 25-003 through -010

CABIN ALTCONTROL

NORMAL OUTFLOW INTONOSE COMPARTMENTCABIN PRESSURE

RELIEF VALVE


CABINUNDERPRESS

SWITCH

CABINPRESS

115 VAC

*FILTER

UPALT

DNALT

NORMALSTATIC SOURCE


CABINALTITUDEEMERGENCYCONTROLVALVE

AUTOMANUAL

AIRBORNELH MAIN

SQUAT SWON GROUND

DUMP

CABINPRESSSWITCH

NORMALSTATIC SOURCE

DUMP INTOTAILCONE

CABIN SAFETYVALVE

STATIC PRESSURE

CABIN PRESSURE

CONTROL PRESSURE

LEGEND

CABINDUMP

28 VDC

*ON SNs 24-150THROUGH 24-169, THEFILTER IS REMOVEDAND THE CABIN AIR EX-HAUST CONTROL VALVEIS PLUMBED TO THECONDITIONED AIR DIS-TRIBUTION DUCT.




CABIN ALTCONTROL

NORMAL OUTFLOW INTONOSE COMPARTMENTCABIN PRESSURE-

RELIEF VALVE


115 VAC

TO CABIN AIR DISTRIBUTIONDUCT (CONDITIONED)

UPALT

DNALT

NORMALSTATIC SOURCE


CABIN ALTITUDEEMERGENCY SWITCH

AUTOMANUAL

UNDER

CABINPRESSSWITCH

CABINUNDERPRESSSWITCH

CABIN PRESS

CABINSAFETYVALVEANEROIDSWITCH

NORMALSTATIC SOURCE

DUMP INTOTAILCONE

CABIN SAFETYVALVE

STATIC PRESSURE

CABIN PRESSURE

CONTROL PRESSURE

LEGEND

FILTER

BELOW8,000 FT

ABOVE10,000 FT

CABINDUMP

SAFETYSWITCH

L MAIN GEARSQUAT SWITCH

AIRBORNE

GNDNORM

DUMP28VDC

INCRALT

Figure 12A-3. Pressurization System — SNs 24-170 through 24B-184, -186, -187,and -189, and 25-011 through -026 and 25-028




Figure 12A-4. Pressurization System — SNs 24B-185, -188 and -190 through -209, and25-027 and -029 through -045

CABIN ALTCONTROL

NORMAL OUTFLOW INTONOSE COMPARTMENTCABIN PRESSURE-

RELIEF VALVE*


RHSQUAT

SWITCH

AIR DISTRIBUTIONDUCT

VACUUMREGULATOR

LHSQUATSWITCH

SECONDARYSOLENOID VALVE

SECONDARYSOLENOID VALVE

FILTER

115 VAC

CHECKVALVE

UPALT

DNALT

NORMALSTATIC SOURCE

ALTERNATESOURCE

CABIN ALTITUDEMANUAL CONTROLVALVE

AUTOMANUAL

UNDER

CABINPRESSSWITCH


CABIN PRESS

DUMP INTOTAILCONE

PRESSURIZATIONJET PUMP

VENT INTOTAILCONE

28 VDCAIR BLEED

AIR

AIR

GND

GND

TIMEDELAYRELAY

TIMEDELAYRELAY

28VDC

DN

GEAR CONTROLSWITCH

AIRCRAFT 25-040 THROUGH25-045 ONLY

UP

FILTER

BLEED AIR

PRESSUREREGULATOR

CABIN SAFETYFILTER

NORMALSTATIC SOURCE

*NOT INSTALLED ON 24B-190 AND SUBSEQUENT, AND 25-030 AND SUBSEQUENT

CONDITIONED AIR

REGULATEDPRESSURE

REGULATEDVACUUM

BLEED AIR

STATIC PRESSURE

CONTROLPRESSURE

LEGEND




CABIN ALTCONTROL

NORMAL OUTFLOW INTONOSE COMPARTMENT

SQUATSWITCHRELAYBOX

NORMALSTATICSOURCE

115 VAC

UPALT

DNALT

NORMALSTATIC SOURCE

AUTOMANUAL

UNDER

CABINPRESSSWITCH


CABINPRESS GND

AIR

CABIN AIREXHAUSTCONTROL

VALVE

CABINALTITUDE

EMERGENCYSWITCH

FRAME 5

REGULATED PRESSURE

REGULATED VACUUM

TIMEDELAY RELAY

GEAR CONTROLSWITCH

CONDITIONEDAIR

DUCT

UP

DN

28VDC

REGVAC

REGPRESS

PRESSURIZATIONMANIFOLD

PRESSURIZATIONJET PUMP

FRAME 22

TO BLEED-AIRMANIFOLD

VENTINTO

TAILCONE

CABIN SAFETYVALVE

C BA

DUMP INTOTAILCONE

REGULATED PRESSURE

BLEED AIR

REGULATED VACUUM

LEGEND

CONDITIONED AIR

STATIC PRESSURE

CONTROL PRESSURE

Figure 12A-5. Pressurization System — SNs 24B-210 through -229and 25-046 through -063, except 25-061

Cabin Outflow ValveThe outflow valve is located on the forwardpressure bulkhead and maintains the desiredcabin pressure by regulating cabin air exhaustflow. The outflow valve position is controlledby two solenoid-operated valves which re-ceive signals from the cabin altitude controller.These solenoid-operated valves allow eitherpositive or negative air pressure to enter a ref-erence chamber to open or close the outflowvalve. The cabin altitude manual control valve,commonly referred to as the “cherry picker,”provides manual control of air routed to the ref-erence chamber. On SNs 24-140 through 24-169, 24B-185, -188 and -190 through -229, and25-003 through 25-010, and 25-027 and 25-029through -063, except 25-061, the outflow valveis opened on the ground by a squat switch ora squat switch relay box.

Pressurization Jet PumpOn airplanes SNs 24-185, -188, -190, andsubsequent, and 25-027, -029, and subse-quent, a pressurization jet pump is mountedon the aft side of frame 25 and provides reg-ulated vacuum and pressure to the outflowvalve. Starting with SNs 24-210 and 25-046,the pressure regulator and the vacuum regu-lator are combined into a single unit calledthe “pressurization manifold.”

Pressurization ControlModule TypicalGeneralThe pressurization control module (Figure12A-6) is mounted in the copilot’s instru-ment panel. The control contains the modeswitch (used to select manual or automaticoperation), an AIR BLEED switch(es), aCABIN AIR FLOW switch (which selectsmaximum airflow), a safety switch (on someairplanes), a cabin altimeter and differentialpressure indicator, a cabin vertical speedindicator, and a cabin altitude manual con-trol valve. See Figure 12A-6 for a typicalpressurization module.

Mode Selector SwitchThis switch, labeled “AUTO–MAN,” selectseither the automatic or manual mode of oper-ation. In AUTO position, a combined signalfrom the cabin altitude selector and the ratesensor are amplified and sent to either the out-flow valve open solenoid (higher cabin alti-tude) or the close solenoid ( lower cabinaltitude). In the MAN position, the signal fromthe amplifier is interrupted, and control of theoutflow valve is provided by the cabin altitudemanual control valve.

Cabin Altitude ControllerThe cabin altitude controller provides meansfor automatically controlling the aircraft cabinpressure during climb, descent, and level flight.The automatic cabin altitude controller modeis activated when AUTO is selected on themode selector switch. In AUTO, the cabin al-titude controller automatically maintains thecabin altitude as the airplane changes altitude.The manual mode deenergizes the cabin alti-tude controller, and the cabin altitude is con-trolled by the cabin altitude manual controlvalve. The system allows comfortable climband descent rates inside the cabin independentof the climb and descent rates of the airplane.Cabin pressure is automatically relieved by




Figure 12A-6. Pressurization ControlModule

the outflow valve when the differential be-tween atmospheric pressure and cabin pressurereaches a preset value. The preset value formodel 23 airplanes is 8.3 psid and for model24, 24B, and 25 airplanes, 8.77 psid.

RATE ControlA knob on the left side of the cabin altitudecontroller dial normally rests in a detent po-sition and provides a 600–700 foot-per-minuteclimb and descent rate. Pulling out the knoband rotating it left or right decreases or in-creases the rate of climb or descent.

Cabin Altitude ManualControl ValveOn SNs 23-003 through -039, a three-posi-tion, spring-loaded cabin altitude switch canbe used to position the outflow valve if the au-tomatic pressurization control system fails.To use the switch, the mode switch, on thecontroller, must first be positioned to MAN.Then, holding the cabin altitude switch to UP

or INCREASE will open the outflow valveand the cabin will climb. Holding the switch toDOWN or DECREASE will close the outflowvalve and the cabin will descend. On most air-planes, this switch has been replaced by a man-ual cabin altitude control valve, such as thoseinstalled on later airplanes and described below.

On SNs 23-030 through 24-229, and 25-003 andsubsequent, and as a retrofit on earlier air-planes, a manual cabin altitude control valve,commonly called the “cherry picker,” is pro-vided to control the outflow valve if the auto-matic pressurization control system fails. Thisthree-position valve, spring-loaded to the cen-ter, is labeled “UP–DN” or “INCREASE–DE-CREASE.” When the valve is held to UP orINCREASE, the outflow valve will open andthe cabin will climb. When the valve is held toDN or DECREASE, the outflow valve willclose and the cabin will descend. The valve isoperational with the mode switch in AUTO orMAN. However, if the mode switch is in AUTO,when the cherry picker is released, the outflowvalve will revert to automatic control.




NORMAL OUTFLOW INTONOSE COMPARTMENT

CABIN ALTITUDEPRESSURE LIMITER

CABIN PRESSURERELIEF VALVE


UPALT

DNALT

NORMALSTATIC SOURCE


CABIN ALTITUDEEMERGENCY SWITCH

CONTROL PRESSURE

CABIN PRESSURE

LEGEND

STATIC PRESSURE

Figure 12A-7. Outflow Valve (Typical)

Differential Pressure Relief(Outflow Valve)A pilot valve in the outflow valve (Figure12A-7) prevents overpressurization by ad-mitting ambient air into the reference cham-ber to open the outflow valve. This valve is setto maintain 8.3 psid on SNs 23-003 through23-099 and 8.77 psid on SNs 24-100 through-299 and 25-003 through 25-063.

Cabin Altitude LimiterA cabin altitude limiter may be installed on theoutflow and cabin safety valve. This device ad-mits cabin air to the reference chamber toclose the valve and keep the cabin altitudefrom exceeding 11,000 ±1,000 feet.

Cabin Pressure Relief ValveSNs 23-003 through 24-189 and 25-003through 25-029 have a cabin pressure reliefvalve installed. The cabin pressure relief valveis a mechanical spring-loaded valve located be-tween the pilot’s rudder pedals on frame 5. Thevalve starts to open to relieve excessive cabinpressure differential at 9.3 psid and is fullyopen at 10 psid.

Cabin Safety (Dump) SwitchOn SNs 23-003 through 24-184, -186, -187,and -189, and 25-003 through -026 and -028,a guarded safety or dump switch (Figure 12A-8) installed on the pressurization module per-mits depressurizing the cabin by opening thecabin safety valve. On some airplanes, thisswitch is deactivated by an aneroid switch

above 10,000 feet and does not reset until theairplane descends below 8,000 feet.

Cabin Under Pressure SwitchAbove 10,000 feet cabin altitude, an aneroidswitch directs all signals from the controlleramplifier to the outflow valve down solenoid.

Cabin Altitude Warning Horn

The horn is activated by an aneroid switchforward of the instrument panel above 10,000feet cabin altitude. A silence switch, locatedon the pedestal forward of the thrust levers,may be used to silence the horn for periods of30 to 50 seconds. The horn may be tested usingthe cabin altitude test switch located next tothe silence switch.




Figure 12A-9. Cabin Safety Valve (Typical)

NORMAL STATIC SOURCE

CABIN ALTITUDEPRESSURE LIMITER

DUMP INTO TAILCONE

CABIN SAFETYVALVE

CABIN PRESSURE

STATIC PRESSURE

LEGEND

Figure 12A-8. Cabin Safety Dump Switch

Cabin Safety ValveA cabin safety valve (Figure 12A-9) is in-stalled in the aft pressure bulkhead to relievea cabin overpressure differential. In flight,the valve normally remains fully closed unlessopened by the pilot valve or activated by thecabin safety (dump) switch. It serves to limitmaximum cabin pressure in case the outflowvalve does not operate properly. On SNs 23-003 through 24-139 and 24-170 through 24B-184,-186,-187, and -189, and 25-011 through-026 and 25-028, the safety valve is opened onthe ground by a squat switch.

Differential Pressure Relief(Cabin Safety Valve)A pilot valve in the cabin safety valve preventsoverpressurization by admitting ambient airinto the reference chamber and increasing cabinaltitude. It is set to maintain 8.7 psid on model23 and 9.12 psid on models 24, 24B, and 25.

CABIN AIR FLOW and AIRBLEED SwitchesAIR BLEED switches on SNs 23-029 and ear-lier allow engine bleed air to flow to the bleed-air manifold and then to the flow control valve.Airplanes with a single AIR BLEED switchcause the flow control valve to move to theclosed position when turned off. On all air-planes, the CABIN AIR FLOW switch deter-mines whe the r a i r f l ow i s r egu l a t ed t oapproximately 180 cfm or whether the flowcontrol valve moves to the fully open position.

INDICATORS

Cabin Altitude and DifferentialPressure IndicatorCabin altitude and differential pressure are in-dicated on a single indicator (Figure 12A-10)incorporating two scales and two pointers onSNs 24-156 and 25-010 and subsequent, and onairplanes with glareshield warning lights.




Figure 12A-11. Cabin Vertical SpeedIndicator

Figure 12A-10. Cabin Altitude and Differ-ential Pressure Indicator

The cabin altitude is indicated by a single pointerand a circular scale on the outer edge withCABIN ALT markings from 0 to 50,000 feet.

The cabin differential pressure is indicatedby a circular scale on the inner portion of theindicator and a single pointer. The scale rep-resents differential pressure from 0 to 10 psi.Earlier airplanes do not have a cabin differ-ential pressure scale, and the cabin altimeteronly indicates to 45,000 feet.

Cabin Vertical Speed IndicatorThe cabin vertical speed indicator (Figure12A-11) indicates a cabin climb or descentfrom 0 to 2,000 feet per minute.

OPERATION

Normal System OperationBefore TakeoffDuring ground operation, the AIR BLEEDswitches are not turned on while the Freonsystem is being used. If the bleed air is turnedon, then engine power must be limited to 70%rpm or less because possible damage to thebleed-air conditioning system may result.Before takeoff, the desired flight altitude andrate of climb should be selected, the AIRBLEED switches turned on, and normal cabinairflow selected.

Flight OperationAutomaticAt liftoff, the left squat switch or the squatswitch relay box on SNs 24-210 and subse-quent and 25-046 and subsequent (except -061), permits either the outflow or cabinsafety valve, as appropriate, to close and pres-surization to begin. The cabin altitude beginsto climb at the rate selected on the controller.On airplanes SNs 24-185, -188, -190, and sub-sequent, and 25-027, -029, and subsequent, atimer activates a secondary down solenoid for23 to 50 seconds after takeoff. This allowsthe cabin to pressurize sooner by closing theoutflow valve more quickly.

ManualIn case of a malfunction of the cabin altitudecontroller or airplane electrical power fail-ure, cabin altitude can be manually controlledby selecting MAN position. The MAN posi-tion removes the cabin altitude controller, ratesensor, and amplifier from the system. Withthe cabin altitude emergency valve (cherrypicker) in the UP ALT position, the referencechamber pressure bleeds to static.

This action opens the outflow path and allowsthe cabin to depressurize. Setting the cherrypicker to DN ALT position allows cabin pres-sure to be admitted to the reference chamber.

The closed outflow path under the main di-aphragm allows the cabin to pressurize. Inmanual mode, the cabin altitude must be mon-itored much more closely than in automaticmode, and the outflow valve position must beadjusted frequently during climbs, descents,and when making power adjustments. On SNs23-003 through 23-029, with 28 VDC notavailable, cabin altitude cannot be adjusted.

DescentWhen descending for a landing, the field baro-metric pressure can be set in the cabin con-troller subdial by pulling out the right knob androtating it to the desired setting. The knobshould return to its normal spring-loaded po-sition with the field elevation set on the outerscale. The barometric pressure on the subdialcan be left at 29.92 unless a radical baromet-ric pressure exists at destination, at whichpoint the subdial should be set to that value.

If the aircraft descends rapidly, it is possiblefor the cabin pressure to become lower thanthe outside pressure. If a negative differentialdevelops, the outflow and/or cabin safety valveopens to relieve the negative pressure.

Should the selected altitude require a cabin dif-ferential pressure greater than the maximumdifferential, the outflow valve overrides thecontroller and maintains the differential withinprescribed limits.




LandingAs the airplane descends and reaches the pre-selected cabin altitude, the outflow valve mod-ulates toward the open position. The cabinshould be unpressurized at landing.

SNs 23-003 through 24-139, -170through -184, -186, -187, and -189, and25-011 through -026 and -028

On landing, the left squat switch opens thesafety valve which dumps any remaining cabinpressure upon touchdown .

SNs 24-140 through -169, -185, -188,-190 through -209, and 25-003 through-010, -027, and -029 through -045

The left squat switch opens the outflow valveand dumps any remaining cabin pressure upontouchdown.

SNs 24-210 through -229 and 25-046through -063, except -061

Either the left or right squat switch controls thesquat switch relay box. When the airplane lands,the squat switch relay box opens the outflowvalve and dumps any remaining cabin pressureinto the nose.

EMERGENCYPRESSURIZATIONStarting with SN 23-030, on those airplanesmodified by SK 23-140 and all subsequentairplanes, an emergency source of pressur-ization bleed air is provided to increase theflow of air into the cabin in the event of a leakor other malfunction.

On SN 23-030 and subsequent, emergencypressurization is provided by use of the wind-shield defog system (see Chapter 10). This isaccomplished by directing bleed air from thebleed-air manifold through the defog shutoffvalve to the cockpit. By closing the externaldefog knob, opening the internal defog knob,and holding the defog switch to DEFOG forapproximately five seconds, emergency pres-surization is accomplished.

On model 23 airplanes converted to model 24and SNs 24-100 and all subsequent airplanes,bleed air from the bleed-air manifold is routedthrough the defog shutoff valve to the cock-pit by holding either the defog switch toDEFOG or the WSHLD HEAT switch toAUTO and pushing the IN NORMAL/OUTDE-FOG knob to IN.

On all airplanes, the AIR BLEED switch isturned off to isolate any leaks between theflow control valve and the check valves in theaft pressurization bulkhead.







1. To regulate cabin pressure, the cabincontroller modulates the:A. Cabin safety valveB. Flow control valveC. Outflow valveD. Primary differential pressure relief

valve

2. The cabin altitude warning horn soundswhen cabin altitude reaches:A. 8,750 feetB. 9,500 feetC. 10,000 feetD. 11,000 feet

3. To dump residual cabin pressure ontouchdown:A. The outflow valve opens automati-

cally.B. The cabin safety valve opens auto-

matically.C. The flow control valve closes auto-

matically.D. The bleed-air shutoff and regulator

valves close automatically.

4. On all airplanes except SNs 23-003through 23-030, if DC power fails:A. Pressurization control reverts to

manual control.B. Pressurization control reverts to

automatic control.C. Cabin pressure is not controlled.D. The cabin slowly depressurizes.

5. The Flight Manual recommends use ofthe maximum airflow position:A. In case of anti-ice failure due to

bleed-air lossB. In case of unsatisfactory pressuriza-

tion in manual modeC. When required during heavy smoke

and fume eliminationD. All the above

6. On airplanes SNs 23-002 through 24-189, and 25-003 through -029, the man-ual pressure relief valve is set to beginopening at:A. 8.77 psidB. 8.3 psidC. 9.12 psidD. 9.3 psid

QUESTIONS

13-i

CHAPTER 13HYDRAULIC POWER SYSTEMS

CONTENTS

Page

INTRODUCTION ................................................................................................................. 13-1

GENERAL ............................................................................................................................ 13-1

HYDRAULIC SYSTEM OPERATION................................................................................ 13-2

HYDRAULIC SUBSYSTEMS............................................................................................. 13-6

QUESTIONS......................................................................................................................... 13-7




13-iii

ILLUSTRATIONS

Figure Title Page

13-1 Hydraulic System Schematic (SNs 24-296 and Prior, and 25-180 and Prior, Except 25-135) ...................................................................................... 13-3

13-2 Hydraulic System Schematic (SNs 24-297 and Subsequent, and25-135, -181, and Subsequent)............................................................................... 13-4

13-3 System Controls and Indicators (Typical) .............................................................. 13-5




INTRODUCTIONTwo engine-driven pumps normally provide hydraulic pressure for operation of thelanding gear, wheel brakes, flaps, spoilers, and thrust reversers. An electrically drivenauxiliary pump, incorporated for use in the event of system failure, is normally used onlyon the ground for operation of the brakes, spoilers, and flaps when both engines are shutdown. It can pressurize all hydraulic subsystems, but should not be used to extend thelanding gear in the event of hydraulic system failure.

GENERALA pressurized reservoir ensures a positive sup-ply of hydraulic fluid to both engine-drivenpumps and to the auxiliary pump. The reser-voir is pressurized to 10 psi by cabin air drawnin by an aspirator in the return line, or to 20psi by regulated engine bleed-air pressure be-ginning with SNs 24-297 and 25-181.

The engine-driven pumps are either constant-volume or variable-volume. The pumps aresupplied from lines connected to the side of thereservoir that reserve approximately .4 gallonfor use by the auxiliary pump that is suppliedby a line connected to the bottom of the reser-voir. Total reservoir capacity is 1.9 gallons.

CHAPTER 13HYDRAULIC POWER

SYSTEMS




Hydraulic shutoff valves installed at the reser-voir in each engine-driven pump supply linecan be closed from the cockpit in the event offire or when maintenance is to be performed.

An accumulator precharged with dry air ornitrogen dampens pressure surges and helpsmaintain system pressure. An indicator on thecopilot’s instrument panel displays systempressure. An optional amber annunciator lightwarns of low pressure.

There are two filters in the system—one in thepressure line, and one in the return line.Systems with reservoirs pressurized by an as-pirator incorporate an air filter.

A pressure regulator on airplanes SNs 24-296 andprior, and 25-180 and prior maintains systempressure at 1,500 psi. On subsequent airplanes,pressure is regulated by the engine-driven pumps.A system relief valve set to relieve at 1,700 psiprevents system damage by relieving excessivepressure into the return line.

The system accumulator and reservoir arelocated in the tai lcone. Accumulator airprecharge, indicated by a gage on the accu-mulator, should be 850 psi when hydraulicpressure is zero. A second, bladder-type ac-cumulator precharged to 600 psi and connectedto the hydraulic system through a one-waycheck valve, is located in the tail compartmenton airplanes with thrust reversers.

HYDRAULIC SYSTEMOPERATIONThe engine-driven pumps are supplied withfluid from the reservoir through hydraulicshutoff valves (Figures 13-1 and 13-2). Thevalves are DC motor-driven and controlled by the guarded FIRE switches on airplaneswith glareshie ld warning l ights , and byguarded FIRE WALL SHUTOFF switches onall other airplanes.

After starting the first engine, the HYDRAULICPRESSURE indicator should be checked toverify engine-driven pump operation.

When the second engine is started, there is nochange in pressure indication, but output isdoubled. There is no positive indication thatthe second pump is operating properly; there-fore, after landing, shut down the enginestarted first and actuate a hydraulic subsystem.If pressure drops slightly then returns to nor-mal, the second pump is functioning properly.

If an engine-driven pump fails in flight, theother engine-driven pump is capable of meet-ing system demands.

The electrically driven auxiliary hydraulicpump, controlled by a switch on the instrumentpanel and a pressure switch (Figures 13-1 and13-2), draws fluid from the bottom of thereservoir. With the HYD PUMP switch in theon (HYD PUMP) position, the auxiliary pumpautomatically cycles between 1,200 and 1,400psi (approximately). It may be used to pres-surize the system in the event of a main hy-draul ic sys tem fa i lure , or for opera t ingsubsystems when the engines are not running.Output is .5 gpm; therefore, a slower-than-normal flap extension should be anticipatedsubsequent to a hydraulic system failure.

Loss of fluid due to a system leak is the mostprobable cause of complete loss of hydraulicpressure. If installed, the amber LO HYDPRESS light will illuminate as pressure de-creases below 1,200 psi. Do not operate theauxiliary pump until emergency landing gearextension procedures have been accomplished,as directed by the approved AFM. Otherwise,the auxiliary pump may discharge the .4 gal-lon of reserve fluid through the same leak.Circuit protection is provided by a circuitbreaker on SNs 23-003 through 24-129 and bya current limiter on all subsequent airplanes.







LANDING GEARDOOR

SELECTORVALVE**

LANDING GEARSELECTOR

VALVE

BRAKESYSTEM

SPOILERSYSTEM

FLAPSYSTEM

THRUSTREVERSERS

RELIEFVALVE(10 PSI)

CONSTANT-DISPLACEMENTENGINE-DRIVEN

PUMP (2)

LEGEND

PRESSURE

SUPPLY

RETURN

AIR (NITROGEN)

PRESSURIZED AIR

ELECTRICAL

FILTER

PRESSUREINDICATOR

CHARGEVALVE

ACCUMULATOR

EFFECTIVITY24D-285 AND SUBSEQUENT25B-154 AND SUBSEQUENT

ACCUMULATOR

OVERBOARD

VACUUMRELIEF

GROUNDSERVICE

AUXILIARYPUMP

PRESSURESWITCH

HYDPUMP

1,700PSI

FIRE

FIRE

AIR FROMPRESSURIZED

CABINJET PUMP(VENTURI)

FIREWALL

SHUTOFF*

28 VDC

28 VDC

PRESSUREREGULATOR

* AIRPLANES WITHOUT GLARESHIELD LIGHTS** NOT INSTALLED ON SNs 23-003 THROUGH -079

Figure 13-1. Hydraulic System Schematic (SNs 24-296 and Prior,and 25-180 and Prior, Except 25-135)




LANDING GEARDOOR

SELECTORVALVE

LANDING GEARSELECTOR

VALVE

BRAKESYSTEM

SPOILERSYSTEM

FLAPSYSTEM

THRUSTREVERSERS

BLEED AIR

FILTER

REGULATOR

RELIEFVALVE(20 PSI)

VARIABLE-VOLUMEENGINE-DRIVEN

PUMP (2)

LEGEND

PRESSURE

SUPPLY

RETURN

AIR (NITROGEN)

BLEED AIR

REGULATED BLEED AIR

ELECTRICAL

FILTER

ACCUMULATOR(850-PSI AIR)

OVERBOARD

VACUUMRELIEF

GROUNDSERVICE

AUXILIARYPUMP

PRESSURESWITCH

HYDPUMP

1,700PSI

28 VDC

28 VDC

FIRE

FIRE

650PSI

Figure 13-2. Hydraulic System Schematic (SNs 24-297 and Subsequent, and 25-135, -181, and Subsequent)




ARMED ARMED

LO HYD PRES S

FIRE

SHUTOFF VALVEINDICATOR

PINHEAD LIGHTAIRPLANES EQUIPPED WITH

GLARESHIELD WARNING LIGHTS

AIRPLANES NOT EQUIPPED WITHGLARESHIELD WARNING LIGHTS

Figure 13-3. System Controls and Indicators (Typical)

HYDRAULICSUBSYSTEMSOperation of the hydraulic subsystems is pre-sented in Chapter 7, “Powerplant,” (ThrustReversers), Chapter 14 “Landing Gear andBrakes,” and Chapter 15, “Flight Controls,”(Flaps and Spoilers).




1. In the event of hydraulic system fail-ure, the optional LO HYD PRESS lightwill illuminate at approximately:A. 1,200 psiB. 1,500 psiC. 1,000 psiD. 850 psi

2. The auxiliary hydraulic pump will pro-vide approximately:A. 1,500 psiB. 1,650 psiC. 1,700 psiD. 1,400 psi

3. If DC electrical power is applied to theairplane, and residual hydraulic pres-sure is 1,450 psi:A. The auxiliary hydraulic pump will

not operate when the HYD PUMPswitch is on.

B. The LOW HYD light, if installed,will be out.

C. 1,450 psi will be shown on the HY-DRAULIC PRESSURE indicator.

D. All the above

4. On airplanes with glareshield warninglights, the hydraulic shutoff valves areclosed:A. With the FIRE switchesB. With the GEN switch in the OFF

positionC. Automatically when the FIRE warn-

ing light comes onD. With the BLEED AIR switches

5. On airplanes without glareshield warn-ing lights, the hydraulic shutoff valvesare closed:A. With the BLEED AIR switchesB. With the GEN switch in the OFF

positionC. With the guarded FIRE WALL

SHUTOFF switchesD. None of the above

6. Engine-driven hydraulic pumps are:A. Constant-volumeB. Variable-volumeC. Gear-typeD. Any of the above, depending on

model and serial number




QUESTIONS

14-i

CHAPTER 14LANDING GEAR AND BRAKES

CONTENTS

Page

INTRODUCTION ................................................................................................................. 14-1

GENERAL ............................................................................................................................ 14-1

LANDING GEAR................................................................................................................. 14-2

Indicating System .......................................................................................................... 14-2

Components ................................................................................................................... 14-3

Operation ....................................................................................................................... 14-6

BRAKES ............................................................................................................................. 14-10

General......................................................................................................................... 14-10

Normal Operation ........................................................................................................ 14-10

Antiskid Systems ......................................................................................................... 14-13

Emergency Brakes ....................................................................................................... 14-14

Parking Brakes............................................................................................................. 14-14

NOSEWHEEL STEERING ................................................................................................ 14-14

General......................................................................................................................... 14-14

Operation—Without Variable Authority (SNs 24-263 and Prior and 25-103 and Prior) ........................................................................................................ 14-15

Operation—With Variable Authority (SNs 24-258, -260, -264, and Subsequent and 25-104 and Subsequent) .................................................................... 14-16

MODEL 23 DIFFERENCES .............................................................................................. 14-17

Landing Gear ............................................................................................................... 14-17

Brakes .......................................................................................................................... 14-19

QUESTIONS....................................................................................................................... 14-20




14-iii

ILLUSTRATIONS

Figure Title Page

14-1 Gear Selector Switch and Position Lights (Typical) .............................................. 14-2

14-2 Gear Position Indications ....................................................................................... 14-3

14-3 Main Gear............................................................................................................... 14-4

14-4 Nose Gear............................................................................................................... 14-4

14-5 Nose Gear Centering Cams.................................................................................... 14-5

14-6 Landing Gear Retracted ......................................................................................... 14-7

14-7 Landing Gear Extended.......................................................................................... 14-8

14-8 Air Pressure Indicator ............................................................................................ 14-8

14-9 Emergency Gear Extension Lever.......................................................................... 14-9

14-10 Emergency Landing Gear Extension...................................................................... 14-9

14-11 Brake System Schematic (Model 25 Airplanes) .................................................. 14-11

14-12 Brake System Schematic (Model 23 and 24 Airplanes) ...................................... 14-12

14-13 Nosewheel Steering System (Airplanes without Variable Authority Steering) ... 14-15

14-14 Nosewheel Steering System (Airplanes with Variable Authority Steering) ........ 14-16

14-15 Landing Gear Control Panel (Model 23 Airplanes)............................................. 14-17




INTRODUCTIONThe landing gear is electrically controlled and hydraulically operated. The main gearincorporates dual wheels equipped with individual hydraulic brakes, and it retracts inboard.The single-wheel, self-centering nose gear incorporates an electrical steering system andretracts forward. Emergency gear extension and emergency braking are pneumatic. An antiskidsystem is incorporated into the normal hydraulic braking system.

GENERALThe landing gear has three air-hydraulic shockstruts. The outboard main gear doors aremechanically linked to the gear and move with it.The inboard doors are hydraulically operated andare closed when the gear is fully extended orretracted. An air bottle is provided for emergencygear extension and emergency braking. The gearactuators incorporate integral downlockingdevices; downlock pins are not required. Gearposition indications are displayed on the instru-ment panel.

The hydraulic brake system is controlled by fourvalves linked to the rudder pedals. Hydraulicsystem pressure is metered to the self-adjustingmultiple disc brake assemblies in proportion topedal deflection.

The antiskid system provides maximum decelera-tion without skidding the tires. When the systemis operating, wheel speed transducers (generators)furnish wheel speed information to a control unitwhich signals the antiskid valves to modulate orlimit braking pressure. The parking brake is set bypulling a handle on the throttle quadrant, whichtraps hydraulic pressure in the brake assemblies.

CHAPTER 14LANDING GEAR AND BRAKES




The electric nosewheel steering operates only onthe ground. When the system is engaged, steer-ing commands from a rudder pedal followup aresupplied to a computer-amplifier. The computer-amplifier determines the amount and directionthat a DC electric motor deflects the nosewheel.

Lear 20 series airplanes may have either variableor nonvariable authority nosewheel steering.

The following Landing Gear, Brakes, andNosewheel Steering sections describe theoperation of those systems. Model 23 differ-ences are presented later in the chapter.

LANDING GEAR

INDICATING SYSTEM

GeneralThe landing gear position indicating system con-sists of three red lights and three green lights, atest switch, and an aural warning horn.

Gear Position LightsThe three green LOCKED DOWN lights (Figure14-1) are illuminated by their respective down-lock switches on the gear actuators.

As each gear locks down, the correspondinggreen LOCKED DOWN light illuminates.During gear retraction, the lights extinguishwhen the downlocks are hydraulically released.

The nose gear red UNSAFE light is illuminatedwhen the gear is unsafe or in transit. When thenose gear is locked in either the up or the downposition, the light extinguishes.

The two main gear red UNSAFE lights illumi-nate whenever the respective inboard gear dooris unlocked. As each inboard door latches up(during extension or retraction), the correspond-ing red light extinguishes.

Indications for gear down-and-locked, up-and-locked, and in-transit conditions are shown inFigure 14-2.

If the gear is extended with the emergency airsystem, all three green lights and the two maingear red lights will be illuminated since bothmain gear doors will remain fully extended.

The landing gear position lights are tested byholding the TEST–MUTE switch on the landinggear control panel in the TEST position. All sixlights will illuminate and the warning horn willsound. The lights can be dimmed with the dim-ming rheostat (Figure 14-2), provided the navi-gation lights are on; otherwise they will be atmaximum intensity.

Circuitry related to the left and right main geargreen position lights may be common with thelanding/taxi light for that side. Confirmation ofmain gear downlocking (after bulb testing) canbe made by switching on the respective LDGLTS switch.

Nose gear green light circuitry may be commonwith the engine synchronizing system (ifinstalled). Confirmation of nose gear downlocking(after bulb testing) is made by positioning theENG SYNC switch on the pedestal to ENGSYNC (on) and observing that the amber ENGSYNC light on the annunciator panel illuminates.




Figure 14-1. Gear Selector Switch andPosition Lights (Typical)

Landing Gear Warning SystemThe aural warning horn will sound and three redUNSAFE lights will come on when the landinggear is not down and locked and either thrust leveris retarded below approximately 70% rpm.

Except on model 23 airplanes, the horn alsosounds when the flaps are extended beyond 25º ifthe landing gear is not down and locked, regard-less of thrust lever position. In this case, the horncannot be muted.

Holding the TEST–MUTE switch in the TESTposition illuminates all six-position indicator

lights and sounds the horn. Momentarily posi-tioning the switch to MUTE silences the hornwhen thrust levers are retarded and the gear isnot down and locked.

COMPONENTS

Main GearEach main gear consists of a conventional air-hydraulic shock strut, dual wheels, torque arm,squat switch, main gear actuator, inboard andoutboard doors, and an inboard door actuator(Figure 14-3).




Figure 14-2. Gear Position Indications

UNSAFE

NOSELEFT

BRT TEST

MUTE

RIGHT

GEARUP

GEARDOWN

LOCKED DOWN

UNSAFE

NOSELEFT

BRT TEST

MUTE

RIGHT

GEARUP

GEARDOWN

LOCKED DOWN

UNSAFE

NOSELEFT

BRT TEST

MUTE

RIGHT

GEARUP

GEARDOWN

LOCKED DOWN

UP AND LOCKED

IN TRANSIT

DOWN AND LOCKED

The main gear hydraulic actuator also serves as aside brace when the gear is extended. It featuresan integral downlock mechanism that can beunlocked only by hydraulic pressure on theretract side; therefore, no downlock pins are pro-vided. Each main gear torque arm link actuates asquat switch.

The main gear is hydraulically held in theretracted position and is enclosed by an outboardand an inboard door. The outboard door ismechanically linked to, and travels with, thegear. The inboard door is hydraulically actuated,electrically sequenced by microswitches, andheld retracted by a spring-loaded, overcenteruplatch that is released by a hydraulic actuator.

With a full fuel load and no passengers orbaggage aboard, 3 to 3 1/2 inches of bright

surface should be visible on the lower portion ofthe main gear strut.

Main Gear Wheel and TiresEach main gear wheel incorporates a fusibleplug that prevents tire blowout caused by exces-sive heat resulting from hard braking. Tires mustbe monitored closely when tread has worn to thebase of any groove at any location. Tires must bechanged any time cord is exposed.

Nose GearThe nose gear consists of an air-hydraulic shockstrut incorporating a self-centering device, anosewheel steering actuator, and mechanicallyoperated doors (Figure 14-4).




INBOARD DOOR

UPLOCK ROLLER

Figure 14-3. Main Gear

Figure 14-4. Nose Gear




The nose gear actuator incorporates an integraldownlock mechanism to maintain a positivedownlocked condition; therefore, a downlockpin is not required. As with the main gearactuator, the locking mechanism can be releasedonly by hydraulic pressure on the retract side.When retracted, the gear is retained by a spring-loaded uplock hook that engages the uplatchroller on the forward side of the strut. Theuplock hook is released by a hydraulic actuator.

When retracted, the nose gear is enclosed by twodoors that are linked to, and travel with, the gear.

An improperly centered nosewheel may jamin the wheel well; therefore, the nose strutincorporates a self-centering mechanism. Atliftoff, two cams are forced together by airpressure within the strut to center the wheel(Figure 14-5).

Since nosewheel centering depends on air pres-sure in the strut, proper inflation of the strut isespecially important. With a full fuel load and nopassengers or baggage aboard, 5 1/4 to 5 3/4inches of bright surface should be visible on thelower portion of the nose gear strut.

If the nosewheel is swiveled 180º from its normalposition, the cams may not be able to center thewheel prior to entering the wheel well. Therefore,the nose gear should be checked during the exteri-or inspection to ensure that the gear-uplock roller(Figure 14-4) is facing forward.

Nose Gear Wheel and TireThe nosewheel tire is chined to deflect water orslush spray (up to 3/4-inch deep) away from theengine intakes during takeoff or landing.

Nosewheel tire pressure should be maintained at105 psi.

Squat Switch FunctionsSNs 23-003 through 24-209 and25-003 through 25-045

Left Switch• Gear sequencing

• Pressurization

• Windshield overheat light

• Takeoff trim light

• Thrust reverser arming

• Antiskid (25 model)

• Stall warning test inhibit (Century III wing)

Right Switch• Gear sequencing

• Pressurization time delay (SNs 24-185, -188, and -190 through -209, and 25-027and -029 through -045)

• Windshield overheat light

• Stabilizer heat

• Thrust reverser armingFigure 14-5. Nose Gear Centering Cams


• Nose steering

• Stall warning test inhibit (Century IIIwing)


SNs 24-210 through 24-357 and SNs25-046 through 25-373

Left Switch• Gear sequencing

• Thrust reverser




• Squat switch relay box to ground

Right Switch• Gear sequencing

• Thrust reverser


• Stabilizer heat


• Squat switch relay box to ground


Individual Squat Switch FunctionsGear—Both squat switches must be in the airposition to allow gear retraction.

Stabilizer heat—Horizontal stabilizer heat is dis-abled on ground.

Takeoff trim—Takeoff trim light is disabledin flight.

Thrust reverser—Thrust reverser deployment inflight is disabled.

Stall test inhibit—Stall warning testing flight onaircraft with Century III wing disabled.

Antiskid—On model 25 aircraft, antiskid systemprevents brake application in flight.

Squat Switch Relay BoxFunctions

• Nose steering

• Pressurization

• Windshield overheat lights

• Inlet heater light

• Hourmeter

Relay Panel FunctionsWindshield overheat—Overheat light is con-trolled by low-limit thermoswitch on ground.

Nosewheel steering—Nose steering is disabledin flight.

Inlet heater light—Inlet heater light is disabledin flight.

Cabin pressure—Cabin pressurization on groundis prevented.

NOTEIf power is lost to panel or relays do not find a ground, systems fail toair mode.

OPERATIONLanding gear operation is controlled andsequenced by microswitches wired in series. Thesystem incorporates two solenoid-operatedhydraulic control valves—one for operation of themain gear doors and one for gear operation. Bothmain gear doors must be fully open before gearcan be extended or retracted.

The gear door control valve is energized to thedoor-open position when the landing gearselector switch is placed in either the GEAR UP







or GEAR DOWN position. It is energized to thedoor-close position by main gear-operatedswitches when the gear is fully retracted or extended.

The landing gear control valve is energized toextend or the retract position by switches locatedon the gear door actuators sensing the full openposition of both main gear inboard doors.During retraction, the circuit is completedthrough both squat switches to ensure that theairplane is off the ground prior to the valve being energized.

Normal RetractionPositioning the landing gear selector switch toGEAR UP energizes the door control valve tothe open position, directing hydraulic pressure torelease the main gear inboard door uplatches andto open the doors. The two main gear redUNSAFE lights illuminate simultaneously withuplatch release.

When the inboard doors are fully open, the door-open switches are actuated. This energizes thegear control valve to the retract position, andhydraulic pressure is directed to retract the land-ing gear (Figure 14-6). The three greenLOCKED DOWN lights extinguish and the nosegear red UNSAFE light illuminates.

When the gear has fully retracted, the nose gearred UNSAFE light extinguishes. When bothmain gear are fully retracted, switches are actu-ated and complete circuitry to energize the doorcontrol valve to the closed position. Hydraulicpressure closes the gear inboard doors, whichlock in position by spring tension on the dooruplatches, and the two main gear red UNSAFElights extinguish.

Normal ExtensionPositioning the landing gear selector switch toGEAR DOWN energizes the door control valveto the open position, thereby directing hydraulic

Figure 14-6. Landing Gear Retracted

TOEMERGENCY

BRAKES

OVERBOARD

GEAR EMERGENCY EXTENSION CONTROL VALVE

EMERAIR

BOTTLE

TO BRAKESYSTEM

UPLATCHACTUATOR

UPLATCH

NOSEGEARACTUATOR

EXTEND RETRACT

SOL SOL

SOL SOL

GEARCONTROL VALVE

DOORCONTROLVALVE

MAIN GEARACTUATOR

UPLATCH ACTUATOR

UPLATCH

DOORACTUATOR

GEAR INBOARD DOOR

DOORACTUATOR

UPLATCH

UPLATCHACTUATOR

PRIORITYVALVE

MAIN GEAR ACTUATOR

LEGEND

SYSTEM HYDRAULIC PRESSURE

RETURN

AIR PRESSURE

pressure to release the main gear inboard dooruplatches and to open the doors. The two maingear red UNSAFE lights illuminate simultane-ously with uplatch release.

When the inboard doors are fully open, the dooropen switches are actuated to energize the gearcontrol valve. This directs pressure to release thenose gear uplatch and extend the nose and maingear (Figure 14-7). The nose gear red UNSAFElight illuminates.

When the gear is fully down and locked, thethree green LOCKED DOWN lights illuminateand the nose gear red UNSAFE lightextinguishes. Circuitry is completed by bothmain gear downlock switches to energize thedoor control valve to the closed position.Pressure closes the gear inboard doors(Figure 14-7), which lock in position by springtension on the door uplatches, and the two maingear red UNSAFE lights extinguish.




Figure 14-7. Landing Gear Extended

TOEMERGENCY

BRAKES

OVERBOARD

GEAR EMERGENCY EXTENSION CONTROL VALVE

EMERAIRBOTTLE

TO BRAKESYSTEM

UPLATCHACTUATOR

UPLATCH NOSEGEARACTUATOR

EXTEND RETRACT

SOL SOL

SOL SOL

GEARCONTROL VALVE

DOORCONTROLVALVE

MAIN GEARACTUATOR

UPLATCH ACTUATOR

UPLATCHDOOR

ACTUATOR

GEAR INBOARD DOOR

DOORACTUATOR

UPLATCH

UPLATCHACTUATOR

MAIN GEAR ACTUATOR

LEGEND

SYSTEM HYDRAULIC PRESSURE

RETURN

AIR PRESSURE

PRIORITYVALVE

Figure 14-8. Air Pressure Indicator




Emergency ExtensionGeneralThe emergency gear extension system is pneu-matically operated by a bottle charged to1,800–3,000 psi with dry air or nitrogen locatedin the nose compartment. Bottle pressure isshown on the direct-reading EMERGENCY AIRindicator on the copilot’s instrument panel(Figure 14-8). The bottle also provides pressurefor emergency braking.

Prior to using the system, the landing gearselector switch (see Figure 14-1) should beplaced in the GEAR DOWN position, and theGEAR circuit breaker on the right essential busshould be pulled. This will prevent inadvertentgear retraction subsequent to a successful exten-sion. The system is activated by depressing theemergency gear extension lever on the left side ofthe pedestal (Figure 14-9). The lever has a ratchetto keep it in the down position, once activated,and can be raised only by pulling an adjacentmetal tab while lifting the lever. Figure 14-9. Emergency Gear

Extension Lever

Figure 14-10. Emergency Landing Gear Extension

TOEMERGENCY

BRAKES

OVERBOARDGEAR EMERGENCY EXTENSION CONTROL VALVE

EMERAIRBOTTLE

TO BRAKESYSTEM

UPLATCHACTUATOR

UPLATCH NOSEGEARACTUATOR

EXTEND RETRACT

SOL SOL

SOL SOL

GEARCONTROL VALVE

DOORCONTROLVALVE

MAIN GEARACTUATOR

UPLATCHACTUATOR

UPLATCH

DOOR ACTUATOR

GEARINBOARD DOOR

DOOR ACTUATOR

UPLATCH

UPLATCHACTUATOR

MAIN GEAR ACTUATOR

LEGENDAIR PRESSURE

RETURN

PRIORITYVALVE

OperationPushing the emergency gear extension leverdown opens a valve to release air bottle pressureto position the gear control and door controlvalves to the extend position (Figure 14-10).This provides a return flow path for fluid in theretract side of the gear and door actuators.

The air pressure repositions the shuttle valves, which:

• Release the nose gear uplatch and themain gear door uplatches

• Open the main gear inboard doors

• Extend all three gear

Since no provision is made to close the maininboard doors, the two main gear red UNSAFElights will remain illuminated. The three greenLOCKED DOWN lights will illuminate.

After the gear is down and locked (providedthere is no hydraulic pressure indicated), airpressure can be relieved in the gear system bypulling an adjacent metal tab while lifting theemergency gear extension lever to the normalposition. This closes the valve on the emergencyair bottle and isolates the remaining air pressurefrom the gear system, preventing a possible leakin the gear system from depleting air pressurethat might be required for emergency braking.Attempting to retract the landing gear after usingthe emergency extension system may causeexcessive air pressure to be introduced into thehydraulic system return lines, thereby rupturingthe reservoir.

BRAKES

GENERALThe brake system is powered by hydraulic sys-tem pressure from the nose gear down (extend)line. The brakes can be applied by either pilot.The system has four multidisc, self-adjustingbrake assemblies, one for each main gear wheel,operated by power brake valves linked to the topsection of the rudder pedals. Braking force is indirect proportion to pedal application unless theantiskid system is activated. The antiskid system

permits stopping in the shortest possible distancefor any given runway condition. Parking brakescan be set by pulling a handle located on the leftside of the throttle quadrant. A pneumatic emer-gency brake system is used to stop the airplane ifhydraulic pressure is lost. Neither antiskid pro-tection nor differential braking is available dur-ing emergency braking.

The brake assemblies are designed to stop for-ward movement of the airplane. If the brakes areapplied with the airplane rolling backward, thebrakes may be damaged.

NORMAL OPERATIONWhen either pilot depresses both brake pedals,the two associated brake valves meter hydraulicsystem pressure (from the nose gear down line)through shuttle valves (one in each main pres-sure line), parking brake valves, antiskid valves,brake fuses, and a second set of shuttle valves,one for each set of four brake assemblies(Figures 14-11 and 14-12). The pilot applyingthe most pressure has control of the brakes.

Pistons in each brake assembly move a pressureplate, forcing the stationary and rotating discstogether against a backing plate to produce thebraking action. Depressing one pedal appliesboth brakes on the corresponding main gear;therefore, differential braking is available, if required.

Releasing pedal pressure repositions the brakevalve, and springs in the brake assembly forcefluid back through the brake valves to the reser-voir, thereby releasing the brakes.

During gear retraction, a restrictor in the returnline creates back pressure on the brakes, whichis sufficient to stop the wheels from rotatingprior to their entering the wheel well.

When taxiing through slush or snow, frequentbrake application creates friction heat, whichmay prevent the brakes from freezing.







PILOTBRAKE VALVE

PILOTBRAKE VALVE

COPILOTBRAKE VALVE

COPILOTBRAKE VALVE

TO RESERVOIR

FROM NOSE GEARDOWN LINE

PARKING BRAKEVALVES

ANTISKIDDISCONNECTSWITCH

ANTISKID

ON

OFF

OVERBOARD

EMERBRAKINGVALVE

GEAREMERGENCYEXTENSIONCONTROL VALVE

SERVO

ANTISKIDVALVE

SERVO

26 VAC

BRAKE FUSE

BRAKEFUSE**

TORESERVOIR

SOLENOIDSHUTOFF

SOLENOIDSHUTOFF

SQUAT SWITCH SQUAT SWITCH

SERVO

ANTISKIDVALVE

SERVO

ANTISKIDCONTROL BOX

ANTISKID GENANTISKID TEST*

*25-204 AND PRIOR, EXCEPT -197**25-184 AND PRIOR

LEGENDSYSTEM PRESSURE

METERED BRAKEPRESSURE

RETURN

EMERGENCY BRAKEAIR PRESSURE

ELECTRICAL

ELECTRICALMECHANICAL

Figure 14-11. Brake System Schematic (Model 25 Airplanes)




PILOTBRAKE VALVE

PILOTBRAKE VALVE

COPILOTBRAKE VALVE

COPILOTBRAKE VALVE

TO RESERVOIR

FROM NOSE GEARDOWN LINE

PARKING BRAKEVALVES

PARKING BRAKEVALVE

ANTISKID

ON

OFF

OVERBOARD

EMERGBRAKINGVALVE

GEAREMERGENCYEXTENSIONCONTROL VALVE

ANTISKIDVALVE

ANTISKIDVALVE

ANTISKIDVALVE

26 VAC

BRAKEFUSE

BRAKEFUSE

ANTISKIDPRINTED CIRCUIT

BOARD

ANTI-SKID GENTESTANTISKID TEST

* 23-003 THROUGH 23-049 ** 23-099 AND PRIOR; 24-310 AND PRIOR † ANTI-SKID GEN LIGHTS, 24-112 AND SUBSEQUENT

LEGENDSYSTEM PRESSURE

METERED BRAKEPRESSURE

RETURN

EMERGENCY BRAKEAIR PRESSURE

ELECTRICAL

MECHANICAL

*

**

†

ANTISKIDVALVE

Figure 14-12. Brake System Schematic (Model 23 and 24 Airplanes)


If a takeoff is made in slush or snow, the wheelsshould be allowed to spin down for approximate-ly one minute prior to gear retraction. This mini-mizes the possibility of brake freezing by sling-ing off accumulated slush. If frozen brakes aresuspected after the gear is extended for landing,brakes should be applied several times to breakup any possible ice formations.

On model 25 airplanes, this requires placing theANTISKID switch in OFF prior to brake actua-tion, then back to ON.

ANTISKID SYSTEMS

GeneralThe antiskid system limits braking on each maingear wheel independently to allow maximumbraking under all runway conditions without tireskidding. One of two antiskid systems may beinstalled. Model 25 airplanes have the Hydroairsystem. Model 23 and 24 airplanes are equippedwith the Goodyear system. On all models, anti-skid is inoperative during emergency braking.

Hydroair SystemThis fully modulated system consists of twoantiskid valves, four wheel speed transducers(one on each main wheel), a control box, fourred ANTI-SKID GEN lights, and a lever-lockingANTISKID switch on the center instrumentpanel. Model 25 airplanes through –204, except–197, have an ANTISKID TEST switch. Model25 airplanes with a Century III wing do not havea test switch. System malfunction is indicated byillumination of the ANTI-SKID GEN lightswhen the ANTISKID switch is on.

The antiskid system is not required to be opera-tional for flight. However, if a malfunction isindicated by illumination of a red ANTI-SKIDGEN light(s), it must be assumed that antiskidprotection is lost on the associated wheel(s).Takeoff and landing data must be computedaccordingly.

OperationAntiskid operation requires:

• The ANTISKID switch must be on.

• Each squat switch must be in the groundmode for its respective brake to function.

• The parking brakes must be released.

• The wheels must be rotating at least150 rpm.

Above 150 rpm, with the ANTISKID switch ONand brakes applied, the control box receives andanalyzes wheel speed inputs from the transduceron each main wheel (see Figure 14-11). If thewheel deceleration rate is higher than a predeter-mined limit, the applicable servo valve will indi-vidually regulate braking force on the corre-sponding brake by releasing braking pressureinto the return line.

A fault in the system is indicated by illuminationof the respective ANTI-SKID GEN light.Cycling the ANTISKID switch to OFF thenback to ON may clear the fault. All four lightsilluminate if power to the control box is lost or ifthe ANTISKID switch is OFF. An ANTISKIDTEST switch is located on the test switch panelforward of the thrust lever on airplanes withoutCentury III wings. Refer to the approved AFMfor test procedures.

At low taxi speeds (wheel speed below 150rpm), the antiskid system is inoperative. The sys-tem is automatically disengaged when the park-ing brakes are set; however, the red ANTI-SKIDGEN lights will not illuminate. To apply thebrakes when airborne, place the ANTISKIDswitch to OFF.

Goodyear System (All Model 23and 24 Airplanes)This partially modulated system consists of foursolenoid-operated antiskid valves, four wheelspeed generators (one on each main wheel), aprinted circuit board, monitor lights (24-112 andsubsequent), and a lever-locking ANTISKIDswitch on the instrument panel.



OperationAntiskid operation requires:

• The ANTISKID switch must be on.

• The airplane must be moving at least5 mph.

Above 5 mph, with the ANTISKID switch ONand brakes applied, the printed circuit boardreceives and analyzes wheel speed inputs fromthe transducer on each main wheel (Figure 14-12). If the wheel deceleration rate is higher thana predetermined limit, the applicable solenoid-operated antiskid valve will individually regulatebraking force on the corresponding brake byreleasing braking pressure into the return line.

Four green ANTI-SKID GEN lights (24-112 andsubsequent) will illuminate at approximately 5mph to indicate wheel generator operation. Afault in the system is indicated by the respectiveANTI-SKID GEN light extinguishing. All fourlights go out if power to the circuit board is lostor if the ANTI-SKID switch is OFF. An ANTI-SKID TEST switch is provided on the testswitch panel forward of the thrust levers. Forsystem testing, refer to the approved AFM.

The brakes may be applied in flight to break upsuspected accumulations of ice on the brakes.

At low taxi speeds (5 mph or less), the antiskidsystem is inoperative, and the ANTI-SKID GENlights will be out.

EMERGENCY BRAKESEmergency airbrakes are provided for use in theevent of normal brake system failure. Antiskidprotection, differential braking, and parkingbrakes are not available while using the emer-gency brakes.

To apply brakes with the emergency system, theEMER BRAKE handle (see Figures 14-11 and14-12) must be carefully pushed down until thedesired braking force is obtained. This meterspressure from the emergency air bottle throughfour shuttle valves to the brake assemblies inproportion to handle movement. A maximum of

600 psi may be applied. Releasing the handlestops flow from the bottle and allows applied airpressure to be vented overboard, releasing the brakes.

PARKING BRAKESNormal hydraulic system pressure from eitherengine-driven pump or the auxiliary pump canbe used to set the parking brakes. Pulling thePARK BRAKE handle on the left side of thepower lever quadrant mechanically closes bothparking brake valves (see Figures 14-11 and 14-12). The closed valves allow pressure from thepilot or copilot brake valves to be trapped in thebrake assemblies.

To set the parking brakes, pull the PARKBRAKE handle rearward. Pedal pressure may beapplied during or after handle movement.Setting the parking brake on model 25 airplanes(see Figure 14-11) disconnects the antiskid sys-tem and prevents inadvertent loss of brake pres-sure. Failure to fully release the parking brakemay result in the antiskid not being operational.

To release the parking brakes, the PARKBRAKE handle must be pushed fully forward.

NOSEWHEEL STEERING

GENERALTwo types of electrically operated nosewheelsteering systems are installed: with or withoutvariable authority. System components include aDC steering actuator, a computer-amplifier, rud-der pedal followups, and switches for systemengagement. In addition, airplanes with variableauthority steering utilize signals from the wheelspeed transducers. AC and DC power are sup-plied through NOSE STEER circuit breakers onthe pilot’s and copilot’s circuit-breaker panels.Airborne, the nosewheel steering system is deen-ergized by the squat switch on airplanes withouta relay box.







The steering actuator, mounted on top of thenose strut, steers the nosewheel through a gear-box and an electrical clutch. The clutch engageswhenever DC power is applied to the airplaneelectrical system, allowing the steering actuatorto function as a shimmy damper, even withsteering disengaged. If DC power is lost or theDC NOSE STEER circuit breaker is out, thenosewheel is free to swivel.

Prior to towing, electrical power must beremoved from the airplane, or the 28-VDCNOSE STEER circuit breaker must be pulled toprevent damage to the system. Since the wheelcan swivel 360°, the nose gear strut must beinspected to ensure that the nose gear uplockroller is pointing forward prior to taxiing.

OPERATION—WITHOUT VARIABLE AUTHORITY (SNS24-263 AND PRIOR AND 25-103AND PRIOR)On airplanes without variable authority, nose-wheel steering can be operated by the primary,wheel master, or steer lock switches.

Primary mode is engaged by depressingand holding either PRI STEER switch(Figure 14-13).

This provides steering capability of 40 to 50° ineither direction for low speed taxi up to10 knots.

Engaging the wheel master or steer lock switch-es permits steering capability up to 10° in eitherdirection. Steering should be disengaged above45 knots.

The wheel master switch on either control wheelpermits nose steering while held depressed.

Steer lock mode is engaged by momentary actu-ation of either STEER LOCK switch, and can bedisengaged by depressing either wheel masterswitch or the PRI STEER switch.

With steering engaged in any mode, the greenSTEER ON annunciator illuminates. Rudderpedal movement provides the displacement anddirectional signals to the computer-amplifierwhich drives the steering actuator in the appro-priate direction until it is stopped by a signalfrom a followup located in the drive gearbox.

WHEEL MASTER SWITCH

Figure 14-13. Nosewheel Steering System (Airplanes without Variable Authority Steering)

If the steering system is inoperative, differentialpower and braking can be used to taxi the air-plane. Pulling the 28 VDC NOSE STEER circuitbreaker will make it easier to steer, but must bereset prior to takeoff or landing to provide shim-my dampening.

OPERATION—WITH VARIABLEAUTHORITY (SNS 24-258, -260,-264, AND SUBSEQUENT AND25-104 AND SUBSEQUENT)Steering authority varies with airplane model.On model 24 airplanes, steering varies from 45°in either direction below 12 knots to 8° at 20knots. On model 25 airplanes, steering variesfrom 50° below 10 knots to 8° at 45 knots.

Variable-authority steering commands to thecomputer-amplifier are controlled by wheelspeed transducers. The system reverts tomaximum travel and should not be used above10 knots if the following conditions exist:

• Model 24 airplanes—Both inboard greenANTI-SKID GEN lights are off.

• Model 25 airplanes—Two or more of thefollowing red ANTI-SKID GEN lights areon: Left inboard, right inboard, and rightoutboard.

Nosewheel steering is engaged by actuating thewheel master or STEER LOCK switch.

To engage wheel master mode, depress andhold the wheel master switch on either controlwheel; disengage it by releasing the switch(Figure 14-14).




Figure 14-14. Nosewheel Steering System (Airplanes with Variable Authority Steering)

WHEELMASTERSWITCH




Engage steer lock mode by momentary actuationof the STEER LOCK switch. Depressing eitherwheel master switch disengages steer lock mode.

With steering engaged in either mode, the greenSTEER ON light illuminates. Rudder pedalmovement provides the displacement and direc-tional signals to the computer-amplifier. Signalsfrom the wheel speed transducers providegroundspeed information. The computer-ampli-fier applies a signal to the steering actuator inthe appropriate direction until it is stopped by asignal from a followup located in the drive gear-box. In the event the steering system is inopera-tive, differential power and braking can be usedto taxi the airplane.

MODEL 23 DIFFERENCES

LANDING GEAR

IndicationOn models 23-003 through 23-009, the landinggear control panel located on the pilot’s instru-ment panel has one red UNSAFE light instead ofthree (Figure 14-15).

If the UNSAFE light remains on at the comple-tion of the retract cycle, one or both main gearinboard doors are not latched or the nose gear isnot locked.

With the nose gear up and locked and the maingear inboard door latched, all lights on the panelare out. With all three gear down and locked, thegreen LOCKED DOWN lights will be on andthe UNSAFE light out.

A warning horn sounds if the thrust levers areretarded below 70% rpm with the gear up. TheTEST–MUTE switch positioned to MUTE willsilence the horn, but the UNSAFE light willremain on.

After an emergency extension, the three greenLOCKED DOWN lights will illuminate, and thered UNSAFE light will be off.

Priority and Boost ValvesOn SNs 23-003 through 23-025, a priority valveand a pressure boost valve are installed in thenose gear retract line. The priority valve pro-vides instantaneous pressure to the pressureboost valve, which, in turn, increases pressure toaccelerate nose gear retraction.

Gear Control ValveThe gear control valve is a four-way, two-position, solenoid controlled valve. The valve,electrically controlled by the gear selectorswitch on SNs 23-003 through 23-079 and thedoor-down microswitches on all other Learjets,directs hydraulic fluid to retract or extend thelanding gear. A pneumatic cylinder is providedon the valve for positioning to “gear down”when using the pneumatic extension system.

Door Control Valve (SNs 23-080and Subsequent)The gear door control valve is a solenoid-operat-ed valve controlled by the gear selector switchand the gear down-and-locked or gear-upswitches. It directs pressure to close the maingear inboard doors after extension or retraction.

Figure 14-15. Landing Gear Control Panel (Model 23 Airplanes)

Directional Control Valves (SNs23-003 through 23-079)A rotary three-position, four-way directionalcontrol valve is installed in the hydraulic pres-sure line to each main landing gear. The purposeof the valves is to close the main gear inboarddoors after the landing gear has extended orretracted. The valves are repositioned by a mechanical linkage connected to the main gear struts.

Operation (SNs 23-003 through23-079)Main Gear-Up CycleMain hydraulic system pressure enters the landinggear control valve where it is directed into the uplines. A 1,100- to 1,150-psi priority valve,installed in the landing gear actuator retract line,diverts the initial hydraulic fluid through the doordirectional control valves to the main gear inboarddoor actuators and to the door uplock actuators.Pressure unlatches the door uplocks and opens theinboard gear doors. When the lock actuators anddoor operating cylinders “bottom out,” the pres-sure rises above 1,100 psi, the priority valves open,allowing fluid to enter the retract side of the maingear actuators, and the main gear retracts. Duringthe last few degrees of main gear retraction,mechanical linkage reverses the door directionalcontrol valve and applies pressure to the close sideof the inboard gear door actuators which closes thedoors. The spring-loaded uplatch engages, com-pleting the main landing gear-up cycle.

Main Gear-Down CycleMain hydraulic system pressure enters the landinggear control valve and is directed into the landinggear-down lines. Pressure is applied simultaneous-ly to the uplock actuator, the door actuator, and the

gear actuator; however, the return fluid from thegear actuator is restricted by an orifice. The uplockopens, the inboard doors open, and the gear slowlyextends. During the last few degrees of main geartravel, mechanical linkage reverses the position ofthe directional control valves. Pressure is appliedto the up side of the inboard door actuators and thedoors close. The downlock in the actuator engagesby spring tension, completing the main landinggear-down cycle.

Operation (SNs 23-003 through23-025)

Nose Gear-Up CycleMain hydraulic system pressure enters the land-ing gear control valve and is directed through apriority valve to the nose gear boost valve. Theboost valve increases the pressure from 1,500 psito approximately 2,340 psi to the up side of thenose gear actuator. The nose gear retracts andcloses the nose gear doors. The spring-loadeduplock engages to complete the nose gear-upcycle. Subsequent airplanes do not incorporate apriority valve and boost valve in the nose gearretract lines.

Emergency ExtensionOn SNs 23-003 through 23-024, emergencyextension is initiated with the emergency gearand brake control handle on the right side ofthe pedestal. Pushing the handle down appliesair to the brakes and also opens a valve thatdirects air pressure to reposition shuttle valvesand extend the landing gear. On SNs 23-003through 23-024 with SK 23/24-43, 23-025 andsubsequent, emergency extension is initiatedwith an emergency extension lever on the leftside of the pedestal.







BRAKES

Parking BrakesOn SNs 23-003 through 23-049, the parkingbrake can be set only by the pilot. Fluid reservechambers in the parking brake valves cause diffi-culty in parking brake release. Depress the ped-als slowly while holding the PARK BRAKEhandle in the off position.

AntiskidModel 23 and SNs 24-100 and subsequent areequipped with the Goodyear antiskid system. Thesystem operates as described earlier in the chap-ter, with one exception—model 23 airplanes donot have ANTI-SKID GEN lights; model 24-112and subsequent do have the lights.

1. If a green LOCKED DOWN light isburned out, main gear down-and-lockedcondition can be confirmed by:A. GND IDLE light illuminatedB. ENG SYNC light illuminatedC. Illumination of the corresponding land-

ing lightD. Red UNSAFE lights illuminated

2. Emergency air pressure can be usedfor:A. Gear extension and parking brakeB. Gear, flaps, spoilers, and brakesC. Gear extension and brakesD. Gear extension, flaps, and brakes

3. Prior to takeoff, the EMERGENCYAIR pressure indicator should indicate:A. 1,800 to 3,000 psiB. Minimum 1,700 psiC. 3,000 to 3,350 psiD. Maximum 1,750 psi

4. Automatic brake snubbing is providedby restricting return fluid from the:A. Antiskid systemB. Engine-driven pumpsC. Nose gear steering systemD. Landing gear system

5. Normal brake pressure is provided bythe:A. Main hydraulic system through the

nose gear downlineB. Brake accumulatorC. Emergency air bottle through the anti-

skid control valvesD. Emergency air bottle

6. On airplanes without variable authoritynosewheel steering, the amount ofsteering available in either directionwith the steer lock mode engaged is:A. 20ºB. 6 to 10ºC. 30ºD. 45º

7. On airplanes with variable authoritynosewheel steering, steering mode(s)available are:A. Steer lock onlyB. Wheel master onlyC. Primary onlyD. Wheel master and steer lock

8. All models except model 23—After anemergency gear extension, the gearposition light indication will be:A. Three greenB. Three green, two redC. Three red, two greenD. Three red, three green

9. Model 23—After an emergency gearextension, the gear position lightindication will be:A. One red, three greenB. Three greenC. Three red, two greenD. Three red, three green

10. The nose gear red UNSAFE light willbe on when:A. The nose gear is unsafe or in transit.B. Nosewheel steering is inoperative.C. The nose gear doors are open.D. The nose gear doors are closed.




QUESTIONS

11. Model 23—The red UNSAFE light willbe on when:A. The nose gear is unsafe or in transit.B. A main gear is unsafe or in transit.C. A and B are correct.D. None of the above

12. Three red UNSAFE lights will be onand the gear warning horn will soundwhen the:A. Gear is retracted and no green

LOCKED DOWN lights are on.B. Gear is down and the thrust levers are

above approximately 70% rpm.C. Gear is up and the thrust levers are

below approximately 70% rpm.D. Flaps are extended below 25°.

13. MODEL 23—The red UNSAFE lightswill be on and the gear warning hornwill sound when the:A. Gear is retracted and no green

LOCKED DOWN lights are on.B. Flaps are extended below 25°C. Gear is down and the thrust levers are

above approximately 70% rpm.D. Gear is up and the thrust levers are

below approximately 70% rpm.

14. Illumination of a main gear redUNSAFE light may indicate:A. The corresponding main gear is not

down and locked.B. The corresponding main gear is not up

and locked.C. The corresponding main gear inboard

door is not fully closed.D. The corresponding main gear inboard

door is locked in the closed position.

15. Model 23—Illumination of the redUNSAFE light may indicate:A. One of the main gear inboard doors is

not fully closed.B. One of the main gear is not down and

locked.C. A and B are correct.D. None of the above.




15-i

CHAPTER 15FLIGHT CONTROLS

CONTENTS

Page

INTRODUCTION ................................................................................................................. 15-1

GENERAL ............................................................................................................................ 15-1

PRIMARY FLIGHT CONTROLS........................................................................................ 15-3

Elevators......................................................................................................................... 15-3

Ailerons.......................................................................................................................... 15-6

Rudder............................................................................................................................ 15-9

TRIM SYSTEMS .................................................................................................................. 15-9

General........................................................................................................................... 15-9

Rudder Trim................................................................................................................... 15-9

Aileron Trim .................................................................................................................. 15-9

Pitch Trim ...................................................................................................................... 15-9

SECONDARY FLIGHT CONTROLS................................................................................ 15-12

Flaps............................................................................................................................. 15-12

Spoilers ........................................................................................................................ 15-14

YAW DAMPERS ................................................................................................................ 15-16

General......................................................................................................................... 15-16

Single Yaw Damper System ........................................................................................ 15-16

Dual Yaw Damper System........................................................................................... 15-16

STALL WARNING SYSTEMS.......................................................................................... 15-18

General......................................................................................................................... 15-18

Stall Warning Indications ............................................................................................ 15-18




Operation—Airplanes with Standard Wing................................................................. 15-19

Operation—Airplanes with Century III Wing............................................................. 15-20

Operation—Airplanes with MKII Wing...................................................................... 15-21

MACH OVERSPEED WARNING/STICK PULLER—ALL AIRPLANES EXCEPT MODEL 23.......................................................................................................... 15-22

General......................................................................................................................... 15-22

Operation ..................................................................................................................... 15-22

Test............................................................................................................................... 15-22

DIFFERENCES .................................................................................................................. 15-22

Wheel Master Switch................................................................................................... 15-22

Control Wheel Trim Switch......................................................................................... 15-23

QUESTIONS....................................................................................................................... 15-24



15-ii FOR TRAINING PURPOSES ONLY

15-iii

ILLUSTRATIONS

Figure Title Page

15-1 Flight Control Surfaces .......................................................................................... 15-2

15-2 Flight Controls Gust Lock...................................................................................... 15-3

15-3 Elevator Control System ........................................................................................ 15-4

15-4 Aileron Tabs ........................................................................................................... 15-6

15-5 Aileron Control System............................................................................................15-7

15-6 Rudder Control System.......................................................................................... 15-8

15-7 Trim Systems Controls and Indicators................................................................. 15-10

15-8 Flap System Schematic ........................................................................................ 15-13

15-9 Spoiler System ...................................................................................................... 15-14

15-10 Spoiler System Schematic.................................................................................... 15-15

15-11 Yaw Damper System............................................................................................ 15-17

15-12 Angle-of-Attack Vane .......................................................................................... 15-18

15-13 Stall Warning System—Airplanes with Standard Wing ...................................... 15-19

15-14 Stall Warning System—Airplanes with Century III Wing .................................. 15-20

15-15 Stall Warning System—Airplanes with MKII Wing ........................................... 15-21




INTRODUCTIONThe manually operated primary flight controls on Learjet 20 series airplanes incorporate elec-trical trim in all three axes. Secondary flight controls consist of hydraulically actuated spoilersand flaps. Other systems related to flight controls are the yaw damper, stall warning, and Machoverspeed warning systems.

GENERALThe primary flight controls, ailerons, elevator,and rudder are direct linkage systems withpushrods, bellcranks, and cables connecting thecontrol surfaces to cockpit controls with nohydraulic or power boost. A rudder/aileron interconnect provides control coordination during turns.

The ailerons incorporate mechanical balancetabs to provide aerodynamic assistance. Roll,yaw, and pitch trim systems are electrically oper-ated and controlled. The trim tabs are installedon the left aileron and the rudder. The movablehorizontal stabilizer provides pitch trim.

The flaps and spoilers are electrically controlledand hydraulically actuated.

20

20 20

105

510 10

5

5

LO

C

GS



CHAPTER 15FLIGHT CONTROLS


On models 23 and 24, a single yaw damper sys-tem provides yaw stability. Model 25 airplaneshave a dual yaw damper system.

A dual stall warning system provides an indica-tion of impending stall by vibrating the controlcolumn and, if no corrective action is taken,induces a forward control column movement toreduce the airplane angle of attack.

A Mach overspeed warning system warns ofoverspeed and on all airplanes induces an aftcontrol column movement to raise the nose ofthe airplane when MMO is exceeded. Model 23airplanes have overspeed warning but do nothave a stick puller.

All flight control surfaces are shown inFigure 15-1.

A flight controls gust lock (Figure 15-2) is pro-vided to prevent wind gust damage to the prima-ry flight control surfaces. When installed, thelock holds full rudder and ailerons, and full-down elevators.




HORIZONTALSTABILIZER

ELEVATOR

FLAP

AILERON TRIM TAB

AILERON BALANCE TAB(BOTH WINGS)

AILERON

VERTICAL STABILIZER

RUDDER

RUDDERTRIM TAB

SPOILER

Figure 15-1. Flight Control Surfaces

PRIMARY FLIGHTCONTROLS

ELEVATORSThe elevators are hinged to the aft edge of thehorizontal stabilizer and are positioned by fore-and-aft movement of the control column. Threescuppers are located near the aft edge of eachelevator for moisture drainage, and static dis-charge wicks are attached to the trailing edge ofeach elevator.

The elevators can also be positioned through anelectrically actuated pitch servo.

On SNs 24-139 and subsequent, all model 25airplanes, and airplanes modified for 45,000-footoperation, a bob weight attached to the controlcolumn and a downspring assembly in the eleva-tor control linkage are incorporated to enhancepitch stability. Airplanes prior to SNs 24-139have an elevator downspring located in the afttailcone but do not have a bob weight.

Figure 15-3 shows the elevator control system.

Pitch ServoThe pitch servo (torquer) is DC operated. It ismechanically connected to the elevator controllinkage through a capstan mechanism incorpo-rating an electric clutch and a mechanical slipclutch. Three flight control systems use the pitchservo to operate the elevators:

• Autopilot—When engaged, the autopilotcan alter noseup or nosedown attitude bycommanding the servo to position the ele-vator up or down, as required.

• Stall warning systems—When engaged,either system can cause the servo to posi-tion the elevator to decrease the angle ofattack in the event of an impending stall(stick pusher).

• Nudger—On airplanes with a two-speedtrim system, the stick pusher applies apusher force in conjunction with shakeractuation.

• Stick puller (all airplanes except model23)—Operating through the L STALLWARNING switch, the system can com-mand the servo to position the elevator fornoseup.




PILOT’S CONTROL WHEEL

PILOT’S RIGHT RUDDER PEDAL

Figure 15-2. Flight Controls Gust Lock




DOWNSPRING ASSEMBLY

AUTOPILOTPITCH SERVO

AFT ELEVATOR SECTOR

FORWARD ELEVATOR SECTOR

FORCE SENSOR

BOB WEIGHT

SNs 140 AND SUBSEQUENT, ALL MODEL 25 AIRPLANES, AND AIRPLANES MODIFIED FOR 45,000-FOOT OPERATION

Figure 15-3. Elevator Control System (Sheet 1 of 2)




SNs PRIOR TO 24-140 EXCLUDING AIRCRAFTMODIFIED FOR 45,000-FOOT OPERATION

Figure 15-3. Elevator Control System (Sheet 2 of 2)

The electric clutch must be engaged to couplethe servo to the elevator linkage for elevatorpositioning. The clutch engages when any oneof the following switches are positioned on:

• L STALL WARNING

• R STALL WARNING

• Autopilot ENGAGE

With all three of the above switches in the offposition, the electric clutch is disengaged, dis-connecting the servo from the elevators.

By exerting sufficient force on the control columnto slip the mechanical clutch, the pilot can over-ride any undesirable servo inputs to the elevators.

Autopilot operation is described in Chapter 16,“Avionics.”

AILERONSAileron control is provided manually by dual-control wheels and automatically by the autopi-lot. Balance tabs are installed on both ailerons todecrease the force required at the control wheel.An aileron trim tab is installed on the leftaileron, and an aileron-rudder interconnectassists turn coordination.

Figure 15-5 shows the aileron control system.

Roll Servo(Autopilot Function Only)The ailerons can also be positioned by theautopilot roll servo. The roll servo is similar tothe pitch servo but does not incorporate an elec-tric clutch. A mechanical slip clutch allows thepilot to override undesired roll servo inputswhen the autopilot is engaged.

Balance TabThe balance tab on each aileron (Figure 15-4)provides aerodynamic assistance in moving the aileron, thus reducing control wheel forces.The tab is attached with one pushrod on model23 airplanes and with dual pushrods on all other airplanes.

Trim TabThe aileron trim tab is attached to the inboardtrailing edge of the left aileron (Figure 15-4).The tab is positioned by either control wheeltrim switch. Aileron trim tab position is indicat-ed by a trim indicator on the pedestal.




AILERONWING

BALANCETAB

TRIM TAB

Figure 15-4. Aileron Tabs




AUTOPILOT ROLLSERVOACTUATOR

Figure 15-5. Aileron Control System




PRIMARY YAW DAMPERSERVOACTUATOR

SERVO SECTOR

RUDDER BELLCRANK

SECONDARY YAW DAMPER

AILERON/RUDDER INTERCONNECT

AIRPLANES CERTIFIED FOR 45,000-FOOT OR HIGHER OPERATION

SERIAL NUMBERS PRIOR TO 24-140EXCLUDING AIRCRAFT MODIFIED FOR 45,000-FOOT OPERATION

RUDDER SECTOR

Figure 15-6. Rudder Control System

RUDDERThe rudder is manually positioned with either setof rudder pedals or by a yaw damper servo. Theyaw damper can be overridden through amechanical slip clutch in the event of a malfunc-tion. The yaw damper can be interrupted onmodel 23 airplanes by depressing and holdingeither wheel master switch. The wheel masterswitch completely disengages the yaw damperon airplanes with the two-speed trim system.

Figure 15-6 shows the rudder control system.

Rudder Trim TabA trim tab, mounted on the bottom trailing edgeof the rudder, is controlled by a trim switch onthe aft end of the center pedestal. Trim positionis indicated by the rudder trim indicator.

TRIM SYSTEMS

GENERALThe airplane is trimmed in the pitch axis bychanging the angle of attack of the movable hori-zontal stabilizer. A dual-motor (normal andemergency) actuator moves the leading edge ofthe horizontal stabilizer up or down in responseto pitch trim inputs. Controls and indicators forthe trim systems are shown in Figure 15-7.

The trim position indicators for pitch, roll, andyaw are all DC powered through the TAB &FLAP PN circuit breaker.

RUDDER TRIM

ControlRudder trim is controlled by the YAW TRIMrudder switch on the aft end of the centerpedestal, spring-loaded to the OFF position.

The switch is split into an upper and a lowerhalf. Both halves must be pressed simultaneouslyto initiate rudder trim tab motion. Model 23 air-planes and SNs 24-100 through 24-169 may

have a single toggle or a rocker switch for ruddertrim. The rudder trim system is DC poweredthrough the YAW circuit breaker.

AILERON TRIM

ControlAileron (roll) trim is controlled with either con-trol wheel trim switch located on the outboardhorn of each control wheel (Figure 15-7). Eachcontrol wheel trim switch is a dual-function (trimand trim arming) switch which controls roll andnormal pitch trim. Each switch has four posi-tions—LWD, RWD, NOSE UP, and NOSEDOWN—and is spring-loaded to the neutralposition. The arming button on top of the switchmust be depressed and held while simultaneous-ly moving the trim switch in the direction ofdesired trim action. Actuation of either controlwheel trim switch to LWD or RWD (with armingbutton depressed) will signal the trim tab actua-tor motor in the left aileron to move the trim tabin the appropriate direction. Actuation of thepilot’s trim switch will override actuation of thecopilot’s switch.

The aileron trim motor is DC powered throughthe ROLL circuit breaker.

Aileron Trim IndicatorAileron trim tab position indication is providedby the AIL trim indicator (Figure 15-7).

PITCH TRIM

GeneralPitch trim is accomplished by repositioning thehorizontal stabilizer to the desired trim settingwith a dual-motor (normal and emergency) actu-ator that operates in three modes:

1. Normal pitch trim mode—Normal trimmotor

2. Emergency pitch trim mode—Emergencytrim motor

3. Autopilot pitch trim mode—Emergencytrim motor







TRIMSWITCH

ARMINGSWITCH

WHEELMASTERSWITCH

PILOT’S CONTROLWHEEL (COPILOT’S

SIMILAR)

Figure 15-7. Trim Systems Controls and Indicators

The pilot-operated normal pitch trim and emer-gency pitch trim systems are electrically inde-pendent systems. Mode selection (normal oremergency) is made with the P TRIM selectorswitch (Figure 15-7). On model 23 airplanes,emergency trim is available with the P TRIMswitch in either NORM or EMER position. Onall subsequent airplanes, emergency trim isavailable with the P TRIM switch in EMERposition only.

Normal pitch trim is controlled through either ofthe control wheel trim switches; emergencypitch trim is controlled through the EMERGEN-CY PITCH trim toggle switch on the aft end ofthe pedestal (Figure 15-7).

All airplanes other than model 23 incorporate atwo-speed NORMAL trim motor and an audibleclicker that signals trim in motion. On these air-planes, trim occurs at the normal rate when theflaps are extended 3º or more. With the flaps up(or extended less than 3º), trim occurs at approx-imately 1/4 the normal rate (slow trim). Anamber PITCH TRIM OVSP annunciator illumi-nates if trim occurs at the normal rate with flapsup or extended less than 3º. The trim in motionclicker indicates movement of the horizontal sta-bilizer and sounds whenever the stabilizer is inmotion for longer than one second with the flapsextended less than 3º.

Autopilot operation uses the emergency motor toadjust pitch trim.

Horizontal stabilizer position is displayed on thepitch trim indicator.

Stabilizer ActuatorThe stabilizer actuator is operated by either oftwo DC-powered motors. The normal trim motorand control circuits are powered through thePITCH circuit breaker. The emergency trimmotor and control circuits are powered throughthe EMERG TRIM circuit breakers.

The emergency trim motor operates at approxi-mately 1/2 the speed of the normal trim motor.

P TRIM Selector SwitchThe P TRIM (pitch trim) selector switch pro-vides the normal and emergency mode selections(Figure 15-7). In the NORM position, normalpitch trim is available from both of the controlwheel trim switches. In the OFF position, bothtrim motors and control circuits are deenergized.In the EMER position, emergency pitch trim isavailable from the EMERGENCY PITCH trimswitch (Figure 15-7). The pilot’s normal trimfunction is rendered inoperative.

The EMERGENCY PITCH trim switch isspring-loaded to the center position.

The autopilot utilizes the emergency trim motorwith the P TRIM selector switch in the NORMor EMER position; however, if either controlwheel trim switch is actuated with the armingbutton depressed (Figure 15-7) or if the EMER-GENCY TRIM switch is actuated, the autopilotdisengages. On model 23 airplanes, actuating thepitch trim switch without arming disengages theautopilot.

In the event of normal pitch trim runaway on air-planes with the two-speed trim system, depress-ing and holding the wheel master switch willstop both the normal and the emergency trimmotors. On model 23 airplanes, the emergencytrim is not interrupted.

It is not a requirement that takeoff trim be setwithin the green band for takeoff. Trim settingsforward of the green band should be used fortakeoff at aft center-of-gravity loadings. Trimsettings near the aft portion of the green bandshould be used for takeoff at forward cg load-ings. However, in no event should takeoff beattempted with the TAKE-OFF TRIM annuncia-tor illuminated.

The TAKE-OFF TRIM annunciator is disabledin flight by the squat switch.




SECONDARY FLIGHTCONTROLS

FLAPS

GeneralThe single-slotted flaps are electrically con-trolled and hydraulically actuated on all air-planes except SNs 23-003 through 23-029,which have mechanically positioned flap controlvalves. The left and right flaps are interconnect-ed by cable to prevent flap asymmetry.

Position switches mechanically connected toeach flap provide flap position information tothe landing gear and stall warning systems.Except on model 23 airplanes, a warning hornsounds if the flaps are extended beyond 25º andthe gear is not down and locked.

Flap Selector SwitchThe flap selector switch is located on the rightside of the throttle quadrant. A schematic of theflap system is shown in Figure 15-8. On SNs 23-003 through 23-089 (unless modified by AMK73-5), there is a single flap actuator.

Flap Position IndicatorA FLAP position indicator is located on the cen-ter instrument panel. Flap position signals areelectrically transmitted to the indicator which isDC powered through the TAB & FLAP PN circuit breaker. The indicator will read UP withloss of electrical power, regardless of actual flap position.

OperationWhen the flap selector switch is placed in theDN position, a hydraulic control valve is posi-tioned to direct pressure to extend both flap actu-ators. The flaps may be stopped at any interme-diate position by placing the selector switch tothe center, neutral position. This removes electri-cal power from the control valve, which movesto the neutral position, trapping fluid between

the control valve and the actuators to hold theflaps in the selected position. Full flap extensionis 40º.

When extended, the flaps are protected fromexcessive airloads by a 1,650-psi relief valve inthe downline. When the relief valve opens,hydraulic pressure is relieved.

If the flap selector switch is left in the DN posi-tion, the control valve will remain energized,hydraulically maintaining the flaps in that posi-tion. If hydraulic pressure is lost under theseconditions, a check valve near the control valvetraps pressure between the control valve and theactuators, holding the flaps in the down position(SNs 24-297 and subsequent, and SNs 25-181and subsequent).

Placing the selector switch in the FLAP UP posi-tion electrically energizes the control valve todirect pressure to the retract side of both actuators.

The flaps on airplanes with the HowardRaisbeck MKII wing are modified for greaterlift, and maximum extension is reduced to 38º.They operate as described above. Optional pre-select flaps may also be installed on these airplanes.

Operation (Preselect Flaps)When the flap selector switch is placed in thedesired position (10º, 20º, DN), the flaps areextended by hydraulic pressure. If increased air-speeds impose excessive airloads on the flaps, arelief valve in the downline opens, allowing par-tial retraction of the flaps. However, the flapswill return to the selected position when the air-load is reduced.

If the flaps move in either direction from the pre-selected position, a series of switches activatesthe flap control valve to hydraulically return theflap to the preselected position.

When the selector switch is moved toward theFLAP UP position, an intermediate stop isencountered at the 20º position to facilitateretraction in a go-around situation. Furthermovement of the selector switch toward FLAP







Figure 15-8. Flap System Schematic

FLAPACTUATOR

INTERCONNECTCABLE

FLAPLIMITSWITCH(PRESELECT)

POSITIONTRANSMITTER

RELIEF VALVE

EXTEND RETRACT

FLAP CONTROLVALVE

*

*

**

*

NORMAL HYDRAULICSYSTEM PRESSURE

RETURN

STATIC

MECHANICAL

ELECTRICAL

LEGEND

AIRPLANES WITH MKII WING

SNs 24-297, 25-181, AND SUBSEQUENT

OPTIONAL ON ALL AIRPLANES WITH MKII WING

*

UP results in further flap retraction.Simultaneous extension of the flaps and spoilersin flight is prohibited because of possible dam-age to the flaps.

SPOILERSThe spoilers, one located on the upper surface ofeach wing forward of the flaps, may be used inflight for rapid descent, or on the ground for liftreduction. The spoilers are hydraulically actuat-ed and electrically controlled by a switch on thecenter pedestal. DC control power for the systemis supplied through the SPOILER circuit breakeron the copilot’s circuit-breaker panel.

OperationWhen the SPOILER switch (Figure 15-9) ispositioned to EXT, the spoilers begin to extendand the red spoiler light illuminates steady. Fullextension is approximately 40º. Returning theswitch to RET causes the spoilers to fully retractand the spoiler light to extinguish.

Spoiler deployment causes significant nosedownpitching and should be anticipated and offsetwith control pressure and pitch trim.

When RET is selected, the spoilers retract andare locked down, and the spoiler light goes out.The resultant noseup pitch motion can be coun-tered with control pressure and trim.

Figure 15-10 shows a schematic of the spoiler system.




Figure 15-9. Spoiler System




SOL SOL

SPOILER CONTROL VALVERETRACT EXTEND

(SHOWN IN EXTEND POSITION)

ACTUATOR ACTUATOR

PRESSURE

RETURN

LEGEND

SPOILER

LH SPOILER RH SPOILER

RESTRICTORRESTRICTOR

Figure 15-10. Spoiler System Schematic

YAW DAMPERS

GENERALA single yaw damper is installed on model 23and 24 airplanes, dual yaw dampers on model 25airplanes. The system(s) provides full-time yawdamping in flight by applying rudder againsttransient motion in the yaw axis. Yaw deviation,sensed by rate gyros and lateral accelerometers,results in damper servos positioning the rudderas required. The damper servos are reversibleDC motors in the cable control system to therudder. They can be overridden by the pilot atany time, if necessary, by the application of suf-ficient force of the rudder pedals.

The yaw damper should be disengaged whiletrimming the rudder, then reengaged.

SINGLE YAW DAMPER SYSTEM

OperationModel 23 and 24 airplanes are equipped with asingle (primary) yaw damper system. A rategyro and a lateral accelerometer provide yawsignals to the computer-amplifier, which trans-mits the necessary drive commands to the yawdamper servo. Airplanes SNs 23-003 through24-129 have two rate gyros with a selectorswitch on the center instrument panel.

The system is engaged by turning on theAUTOPILOT MASTER switch and the YAWDAMPER switch on the subpanel on model 23airplanes or by depressing the YAW DAMPERON button on the autopilot controller (Figure15-11) on model 24 airplanes.

On model 23 airplanes, the yaw damper is inter-rupted while the control wheel master switch isdepressed and held. On all other airplanes,depressing the wheel master completely disen-gages the yaw damper and a warning tonesounds. Turning off the AUTOPILOT MASTERswitch or the YAW DAMPER switch on thepilot’s subpanel or depressing the YAWDAMPER OFF button on the autopilot con-troller disengages the yaw damper.

DUAL YAW DAMPER SYSTEMAll 25 series airplanes are equipped with dual(primary and secondary) yaw damper systems.Each system is completely independent with itsown rate gyro, lateral accelerometer, computer/amplifier, and damper servo. Yaw deviation sig-nals from the rate gyro and lateral accelerometerare processed by the computer/amplifier, anddrive signals are transmitted to the applicabledamper servo for rudder deflection.

Prior to engagement of either the primary or secondary damper, the respective force indica-tors should be centered. The primary damperforce indicator is on the autopilot force indica-tor panel on the pilot’s instrument panel. Thesecondary damper force indicator is located onthe yaw damper panel on the center pedestal(Figure 15-11).

To engage the primary yaw damper, position theselector switch on the yaw damper panel to PRIYAW DAMPER, then depress the YAWDAMPER ON button on the autopilot flight con-troller. Disengage the primary damper bydepressing the YAW DAMPER OFF button, bydepressing the wheel master switch on eithercontrol wheel, or by selecting the secondary yaw damper.

To engage the secondary yaw damper, positionthe selector switch to SEC YAW DAMPER, thenposition the SEC YAW DAMPER switch toENGAGE. Disconnect the damper by movingthe SEC YAW DAMPER switch out of theENGAGE position, by depressing the wheelmaster switch on either control wheel, or byselecting the primary yaw damper.

On SNs 25-363 and subsequent, the yaw damperis automatically disengaged by the squat switch-es at touchdown.

Only one yaw damper can be engaged at a time.







PRIMARY YAW DAMPER SERVO

SECONDARY YAW DAMPER SERVO **

**

**

*

*

SELECTOR SWITCH

AUTOPILOT EFFORTINDICATORS

AUTOPILOT FLIGHTCONTROLLER

ALL MODELS

25 SERIES AIRPLANES

Figure 15-11. Yaw Damper System

STALL WARNING SYSTEMS

GENERALOne of three stall warning systems is installed,depending on the type of wing used. Each systemprovides visual and tactile warning of an impend-ing stall, and includes the following major com-ponents: left and right stall vanes, transducers,computer/amplifier, stick shaker motors, angle-of-attack indicators, L and R STALL WARNINGswitches, and red L and R STALL warning lights.All three systems use the autopilot pitch servo forstick pusher operation. All three systems are hotwired to the battery bus.

STALL WARNING INDICATIONSThe systems present indications of approachingstall and full stall with ANGLE OF ATTACKindicators, red L and R STALL lights, and theuse of control column shakers, nudgers (on air-planes other than model 23), and a stick pusher.

Stall Vanes and TransducersDuring flight, angle-of-attack vanes (Figure 15-12) on both sides of the forward fuselage alignwith the local airstream. Transducers operatedby these vanes produce voltages proportional toairplane angle of attack. These voltage signalsare processed by the system computer/amplifier.

ANGLE OF ATTACK IndicatorsThe computer/amplifier translates signals fromthe stall vane transducers into visual indicationsof stall margin on the ANGLE OF ATTACKindicators. The indicator face is divided intothree colored segments—green, yellow, and red.The green segment represents the normal operat-ing range, the yellow segment warns of anapproaching stall, and the red segment indicatesthat the angle of attack is at or just above aero-dynamic stall. On airplanes with the Century IIIwing, the left and right ANGLE OF ATTACKindicators are controlled by their respective stallvane. On airplanes with the MKII or standard

wing, each indicator is controlled by the vane onthe opposite side of the fuselage.

Stick (Control Column) ShakerShaker motors are attached to the front side ofeach control column. Actuation of the shakerscauses low-frequency, high-amplitude vibrationin the control columns.

Stick PusherThe stick pusher function utilizes the autopilotpitch servo to reduce angle of attack by decreas-ing airplane pitch attitude. Pusher activation pro-vides elevator down motion, causing a suddenabrupt forward movement of the control column.Elevator servo actuation during pusher engage-ment can result in up to 80 pounds of forceapplied to the control column if activated simul-taneously by both systems. However, the pushercan be engaged by either system independently.In this event, pusher force is limited to 50pounds, which is diminished linearly by anaccelerometer to maintain 0.5 g. In the event ofinadvertent pusher engagement due to malfunc-tion, the pilot can override the pusher.

NudgerOn airplanes other than model 23, a nudger isincorporated into the stall warning system. Asangle of attack increases to the point of shaker




Figure 15-12. Angle-of-Attack Vane

motor operation (but prior to pusher operation), apush command is applied to the pitch servo withapproximately 1/4 the force of the pusher.

OPERATION—AIRPLANESWITH STANDARD WINGThe vane-operated transducer provides a voltage,proportional to airplane angle of attack, to thecomputer/amplifier. If angle of attack increasesuntil airspeed is 7% above stick pusher actuationspeed, one or both stick shakers are energized and

one or both red STALL lights begin to flash(Figure 15-13). When both vanes simultaneouslyincrease to 5% above aerodynamic stall condition,the computer/amplifier signals the autopilot to dis-engage and the pitch servo to apply nosedown ele-vator (stick pusher). When angle of attack decreas-es below the stall point, the pusher releases.

Stick shaker actuation occurs as the pointersweeps 1/3 to 1/2 into the yellow segment of theindicator. The stick pusher actuates as the indica-tor moves into the red segment.




Figure 15-13. Stall Warning System—Airplanes with Standard Wing

ANGLE-OF-ATTACKINDICATOR

PUSHER ACTUATION

SHAKER ACTUATION

NORMAL ACCELEROMETER(1/2 g LIMITER)

ACCELEROMETERCUTOUT BOX

BIASBOX

COMPUTER/AMPLIFIER

TO LHANGLE-OF-ATTACKINDICATOR

VANEVANE STALL WARNING

L R


SHAKERMOTOR

SHAKERMOTOR

L STALL R STALL

*

* FLAP POSITION INPUT

OPERATION—AIRPLANESWITH CENTURY III WINGAs airplane angle of attack increases, the vane-operated transducers feed a voltage signal intothe computer/amplifier, along with flap positionand altitude information (closing of the altitudeswitches at 22,500 feet increases the stick pusheractuation point by approximately 15 knots aboveairplane stall speed). The computer/amplifiersums the signals and drives the pointer of theANGLE OF ATTACK indicators accordingly(Figure 15-14).

If angle of attack increases until airspeed is 7%above pusher actuation, the red L and R STALL

lights flash, the indicator pointer enters the yel-low segment, and the stick shakers actuate. Ifangle of attack continues to increase until air-speed is approximately 3 knots above airplanestall speed, the autopilot disengages and thestick pusher applies nosedown elevator.

Testing of the system is accomplished on theground with the stall warning test switch on thecenter pedestal. Holding the switch to R TESTcauses the pointer on the right ANGLE OFATTACK indicator to sweep across the greensegment. As the pointer enters the yellow seg-ment, the right shaker actuates and the R STALLlight flashes. As the pointer sweeps into the redsegment (STALL lights on steady, airplanes with




PUSHER ACTUATION

SHAKER ACTUATION



BIASBOX

RATESENSOR

COMPUTER/AMPLIFIER


L R


SHAKERMOTOR

SHAKERMOTOR

L STALL R STALL

*

* BIAS INPUTS:• FLAP POSITION• ALTITUDE

ANGLEOF

ATTACK

ANGLEOF

ATTACK

Figure 15-14. Stall Warning System—Airplanes with Century III Wing

two-speed trim system), the stick pusher actu-ates, then releases as the pointer recedes into theyellow segment. Repeat the test, positioning theswitch to the L TEST position. The test switch isdisabled in flight by its respective squat switch.

OPERATION—AIRPLANESWITH MKII WINGAs angle of attack increases, vane-operatedtransducers feed a voltage signal proportional toairplane angle of attack to the computer/amplifi-er. If angle of attack increases until speed isapproximately 7% above airplane stall speed, thecomputer/amplifier causes the stick shakers toenergize and the L and R STALL lights to flash(Figure 15-15).

When both vane-operated transducers simultane-ously detect an angle of attack beyond stall con-ditions, the computer/amplifier disengages theautopilot and signals the pitch servo to applynosedown elevator (stick pusher). When angle ofattack decreases below the stall point, the stickpusher releases.

Stick shaker actuation occurs when the pointerof the indicator is in the yellow segment. Thestick pusher actuates when the pointer is in thered segment, and, except for model 23 airplanes,the STALL lights are on steady.




PUSHER ACTUATIONSHAKER ACTUATION



BIASBOX

COMPUTER/AMPLIFIER


L R


SHAKERMOTOR

SHAKERMOTOR

L STALL R STALL

*

* FLAP POSITION INPUT

161412108

64

2

18

0

MK IIANGLE

OFATTACK

161412108

64

2

18

0

MK IIANGLE

OFATTACK

Figure 15-15. Stall Warning System—Airplanes with MKII Wing

MACH OVERSPEEDWARNING/STICKPULLER—ALL AIRPLANES EXCEPTMODEL 23

GENERALThe stick puller system provides audible over-speed warning and noseup elevator through theautopilot pitch servo in the event airspeed reach-es MMO (.82 Mach).

The stick puller utilizes the autopilot pitch axiscircuitry to control the elevator servo forceapplied. The resultant noseup force on the con-trol column during puller actuation is approxi-mately 18 pounds. If the autopilot is engaged,puller actuation cancels any selected flight direc-tor pitch modes and inhibits autopilot use of thepitch servo until the puller is released. Systemcontrol circuits require 28 VDC suppliedthrough the L STALL WARN circuit breaker.Power for the stick puller system is controlledthrough the L STALL WARNING switch. Thesystem will be inoperative if the switch is in theOFF position.

OPERATIONThe overspeed warning horn is functional when-ever the airplane electrical system is poweredand either WRN LTS circuit breaker is closed.The stick puller system becomes functionalwhen the L STALL WARNING switch is turnedon. The STALL WARNING switches shouldremain on at all times in flight except as indicat-ed in the approved AFM. With the stick pullerinoperative, speed is limited to 0.74 Mach onairplanes with a Century III wing. On airplaneswith a MKII or standard wing, an inoperativestick puller limits the airplane to 30,000 feet or below.

TESTWith the BATTERY switch(es) on, test the stickpuller by placing the L STALL WARNINGswitch on and actuating the MACH TEST switchon the test switch panel.

DIFFERENCES

WHEEL MASTER SWITCHDepressing the wheel master switch performsthe following functions:

Model 23• Interrupts normal pitch, roll, and yaw trim

while depressed

• Disengages the autopilot

• Permits maximum of 10º nosewheel steer-ing on the ground

• Interrupts yaw damper (while helddepressed)

• Centers nose gear when airborne if gear isextended and switch is held

Airplanes SNs 24-100 andSubsequent, and 25-003 andSubsequent

• Disengages the autopilot

• Permits maximum of 10º nosewheel steer-ing (Nonvariable-authority airplanes)

• Disengages steer lock function

• Disengages the yaw damper

• Interrupts pusher and puller whiledepressed

• Interrupts normal and emergency pitchtrim while depressed

• Centers nose gear when airborne if gear isextended and switch is held




CONTROL WHEEL TRIMSWITCHOn model 23 airplanes it is possible to trim theailerons with the autopilot engaged. On airplaneswith two-speed trim, any trimming action disen-gages the autopilot completely.




1. The airplane systems that use the pitchservo to position the elevator are:A. Autopilot and Mach trimB. Autopilot, stick pusher, and stick pullerC. Pusher and Mach trimD. Yaw damper and stick pusher

2. The airplane is trimmed in the pitchaxis by:A. Movable trim tabs on the elevatorsB. CanardsC. The movable horizontal stabilizerD. A downspring on the elevators

3. To enable pitch trim through the con-trol wheel trim switches, the P TRIMselector switch must be in the:A. NORM or EMER positionB. NORM, OFF, or EMER positionC. NORM positionD. EMER position

4. The systems that can function with theP TRIM selector switch in the EMERposition are:A. Normal pitch trim and Mach trimB. Emergency pitch trim and Mach trimC. Emergency pitch trim and normal pitch

trimD. Emergency pitch trim and autopilot

pitch trim

5. In the event of airplane electrical fail-ure, the flap position indicator will:A. Be powered by the emergency battery

and indicate actual position of the flapsB. Not be powered and will freeze at last

flap positionC. Not be powered and will go to fullscale

up deflection regardless of flap positionD. None of the above

6. The electrical power source for the stallwarning system is provided by the:A. Battery busB. Battery-charging busC. Main DC busesD. Emergency battery

7. If either L or R stall warning system isfound to be inoperative before takeoff:A. The airplane can be flown provided the

circuit breaker is pulled for the inopera-tive system.

B. The airplane may be flown provided thepilot has an ATP rating.

C. The airplane may be flown provided theautopilot and yaw damper systems areoperating.

D. The airplane must not be flown.

8. The switch used to turn the stick pullersystem on and off is the:A. STICK PULLER switchB. AUTOPILOT master switchC. L STALL WARNING switchD. R STALL WARNING switch




QUESTIONS

16-i

CHAPTER 16AVIONICS

CONTENTS

Page

INTRODUCTION ................................................................................................................. 16-1

GENERAL ............................................................................................................................ 16-1

AIR DATA SYSTEM ............................................................................................................ 16-3

Pitot-Static System......................................................................................................... 16-3

Air Data Sensor.............................................................................................................. 16-4

Ram-Air Temp Gage...................................................................................................... 16-4

AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS) ...................................................... 16-5

General........................................................................................................................... 16-5

Component Description ................................................................................................. 16-5

Flight Controller Operation ........................................................................................... 16-7

Stability System (AFC/SS) ............................................................................................ 16-9

FC 110 AFCS Emergency Condition Procedures.......................................................... 16-9

FLIGHT DIRECTOR SYSTEM ........................................................................................... 16-9

General........................................................................................................................... 16-9

Attitude Director Indicator (ADI)................................................................................ 16-11

Horizontal Situation Indicator (HSI) ........................................................................... 16-11

Mode Selector.............................................................................................................. 16-11

Operation ..................................................................................................................... 16-11

RVSM SYSTEM ................................................................................................................. 16-13

General......................................................................................................................... 16-13

System Control ............................................................................................................ 16-13

Indications ................................................................................................................... 16-13

Operating Procedures .................................................................................................. 16-14

QUESTIONS....................................................................................................................... 16-18




16-iii

ILLUSTRATIONS

Figure Title Page

16-1 Pitot-Static System (Typical) ................................................................................. 16-2

16-2 Pitot Head (Typical) ............................................................................................... 16-3

16-3 Pitot-Static Drains (Typical)................................................................................... 16-3

16-4 Static Ports (Typical).............................................................................................. 16-3

16-5 ALTERNATE STATIC SOURCE Valve ................................................................ 16-4

16-6 RAM AIR TEMP Gage.......................................................................................... 16-5

16-7 Autopilot Flight Controller .................................................................................... 16-5

16-8 Control Wheel Switches (Pilot’s) ........................................................................... 16-6

16-9 Autopilot Effort Indicators ..................................................................................... 16-7

16-10 Flight Director Panels ............................................................................................ 16-9

16-11 ADI and HSI Symbology..................................................................................... 16-10

16-12 Altimeter System ................................................................................................. 16-12

16-13 Air Data Panel...................................................................................................... 16-13

16-14 Air Data Panel Indications ................................................................................... 16-14

16-15 Air Data Display Unit (ADDU)........................................................................... 16-14




INTRODUCTIONThe Learjet 20 series avionics consists of, but is not limited to, the air data system, theautomatic flight control system, the navigation system, and the communication system.This chapter includes information on the air data, automatic flight control, and flightdirector systems used in the Learjet 20. The user should consult applicable supplementsin the approved AFM and vendor manuals for additional information and informationon specific systems not included in this chapter.

GENERALThe air data system consists of the pitot-staticsystem, the air data sensor (if installed) or au-topilot altitude controller, and the RAM AIRTEMP gage.

The automatic flight control system includesthe autopilot and yaw damper. The standardautomatic flight control system installed on

the Learjet 20 series is the Jet Electronicsand Technology, Inc., (JET) FC 110. The yawdamper system operates independently of theautopilot and may be engaged with or with-out the autopilot engaged. The yaw dampersystem is described in Chapter 15, “FlightControls.”

CHAPTER 16AVIONICS







LH SHOULDERSTATIC PORT

AUTOPILOTALTITUDE

CONTROLLERALTITUDE

PRESSURESWITCH 1

DRAIN VALVE 2

DRAIN VALVE 2

EPRSYSTEM 2

ROC(P) 4

PRESSMODULE OR

OUTFLOW VALVE 6

ALTPRESSSW 1

DRAIN VALVE 2

DRAIN VALVE 2

MACH TRIMSWITCH 3

CABALT 5

AIR DATA SENSOR 7

DRAINVALVE

RH PITOT HEAD

ALT

HO

LD

ALT

RAT

E

IAS

MA

CH

RH FWDSTATIC PORT

RH SHOULDERSTATIC PORT

RH AFTSTATIC PORT

MACH WARNING ANDOVERSPEED SWITCH

ALTIMETER(COPILOT)

ROC(COP)

RH CENTERSTATIC PORT

EPRDRAINVALVE

AIRSPEED ANDMACH INDICATOR

(COPILOT)

CABINSAFETY

VALVE

REAR PRESSUREBULKHEAD

STATIC PORT

FORWARDPRESSUREBULKHEADSTATIC PORTINSTRUMENT


VALVE

AIRSPEED AND MACHINDICATOR (PILOT)ALTIMETER

(PILOT)

LH AFTSTATIC PORT

LH FWDSTATIC PORT

LH PITOT HEAD

STATIC DEFECTCORRECTION

MODULE

ALT STATIC PORT(NOSE COMPT)

PILOT’S PITOT

COPILOT’S PITOT

PILOT’S STATIC

LEGEND

COPILOT’S STATIC

ALTERNATE STATIC

OTHER STATIC

NOTES:

1. CENTURY III RAS EQUIPPED ONLY

2. SNs 24-138 AND SUBSEQUENT, 25-003 AND SUBSEQUENT, AND AIRPLANES WITH HEATED STATIC PORTS

3. ALL AIRPLANES WITH AFC/SS

4. MAY BE ELECTRICAL

5. AIRPLANES WITH IDC PRESSURIZATION SYSTEM

6. OUTFLOW VALVE—IDC PRESSURIZATION SYSTEM PRESSURIZATION MODULE—GARRETT PRESSURIZATION SYSTEM

7. SN 25-270 AND SUBSEQUENT, AND THOSE AIRPLANES WITH ENGINE STALL WARNING SYSTEM

Figure 16-1. Pitot-Static System (Typical)

AIR DATA SYSTEM

PITOT-STATIC SYSTEMThe pitot-static system supplies pitot and staticair pressure for operation of the airspeed andMach indicators, the Mach trim switch, theMach warning and overspeed switch, the EPRsystem (on airplanes certified for flight inknown icing conditions), the air data sensoror autopilot altitude controller, and the staticdefect correction module. Static pressure issupplied to the rate-of-climb indicators, al-timeters, cabin differential pressure indicator,pressurization control module, outflow valve,and the cabin safety valve. The equipment in-stalled varies with airplane model and serialnumber. Figure 16-1 shows a typical pitot-static system incorporating all equipmentwhich may be installed on the Lear 20.

A heated pitot head is located on each side ofthe fuselage just forward of the cockpit (Figure16-2). Left and right PITOT HEAT switchesare located on the pilot’s subpanel. (SeeChapter 10, “Ice and Rain Protection,” for ad-ditional information.)

Condensate drains (Figure 16-3), locatednear the left and right nose wheel well doors,are provided for each pilot and static systemon all airplanes SNs 24-138 and subsequentand all 25 models.

The static systems provide independent sourcesof static pressure to the pilot’s and copilot’s

instruments. Each static system has one staticport on each side of the airplane nose (Figure16-4).

The dual pickups are provided for redundancyand to reduce sideslip effects on the instru-ments which use static air.

The left forward and right center static ports areconnected to the pilot’s instruments. The left rearand right forward static ports are connected




Figure 16-3. Pitot-Static Drains (Typical)

Figure 16-4. Static Ports (Typical)

Figure 16-2. Pitot Head (Typical)

to the copilot’s instruments. The right rear staticport is connected with an alternate static portinside the nose compartment to provide thepressurization module or outflow valve with astatic source. The static source goes to the pres-surization module on models with the Garrettpressurization system, or to the outflow valveon airplanes with the IDC pressurization sys-tem. A static source inside the tailcone on all 20series airplanes provides static pressure to thecabin safety valve. Refer to Chapter 12,“Pressurization,” for additional information.

Two shoulder static ports are located on topof the fuselage nose in front of the windshield.These ports provide static pressure to the au-topilot altitude controller on the air data sen-sors and an a l t i tude pressure swi tch onairplanes with the Century III wing.

On airplanes SNs 24-138 and subsequent andall 25 models, the two forward static ports oneach side are heated. Circuit protection isthrough the PITOT HEAT circuit breakers.On those airplanes with the Learjet CenturyIII RAS, the shoulder static ports are alsoheated. Refer to Chapter 10, “Ice and RainProtection,” for additional information.

An ALTERNATE STATIC SOURCE valve islocated below the pilot’s instrument panel(Figure 16-5) and connects the pilot’s staticsystem to an alternate source located on frame5. For normal operation, the lever remainsdown (CLOSED); for alternate air, the leveris moved up (OPEN).

When the ALTERNATE STATIC SOURCEvalve is positioned to OPEN, the pilot’s staticinstruments are connected to an alternate portinside the unpressurized nose section. See theAFM for corrections to be applied to the air-speed and the altimeter when using the alter-nate static source.

AIR DATA SENSORAn air data sensor is installed on airplanesSNs 25-270 and subsequent and those air-planes modified per AMK 81-12 (installationof engine stall warning system). The air datasensor is installed in the nose compartment justforward of the nose wheel well box. It re-ceives a static input from the shoulder staticports. The pitot input is from the copilot’sp i t o t sy s t em. The air data sensor provides altitude hold, alti-tude rate, indicated airspeed, and Mach num-ber output signals. High altitude and lowaltitude switch outputs and synchronous airdata output are also available.

RAM-AIR TEMP GAGERam-air temperature is displayed on the RAMAIR TEMP gage, located on the copilot’s sub-panel (Figure 16-6). The gage is calibratedin degrees Celsius. For conversion to outsideair temperature (OAT), refer to the Ram-AirOutside Temperature Conversion figure in the“Performance” section of the approved AFM.




Figure 16-5. ALTERNATE STATIC SOURCEValve

AUTOMATIC FLIGHTCONTROL SYSTEM(AFCS)

GENERALThe JET FC 110 automatic flight control sys-tem (AFCS) reduces the pilot’s workload byflying the airplane to and holding desiredheadings, attitudes, and altitudes. The AFCScan be used for capture and track of VOR/ILSradio signals and provides a back course fea-ture. The FC 110 system uses 115 VAC and 28VDC.

The autopilot, when engaged, has control in twoaxes—pitch and roll. The yaw damper systemoperates independently of the autopilot andmay be engaged with or without the autopilotengaged. Detailed information on the yawdamper system is given in Chapter 15, “FlightControls.”

COMPONENT DESCRIPTIONComputer Amplifier—The computer amplifierfor the AFCS is located on the floor of thecockpit beneath the pilot’s seat.

Servos—The 28-VDC servos are mounted inthe following locations:

• Roll—under the floor in the forwardpart of the cabin

• Pitch—in the vertical stabilizer (leftside of tail compartment above the bat-teries on SNs 23-003 through 24-138)

• Yaw—in the vertical stabilizer (rightside of tail compartment above the bat-teries on SNs 23-003 through 24-138)

• Secondary yaw (all 25 series)—in thetailcone

The servos receive electrical signals from theautopilot computer and move the control sur-faces as required.

Flight Controller—The autopilot flight con-troller is mounted on either the instrumentpanel or the center pedestal (Figure 16-7).

The flight controller contains all of the AFCSmode buttons, engage button, on–off lights,roll turn knob, and pitch command wheel. Yawdamper buttons are included. All mode buttonsare solenoid engaging and stay depressed untilcancelled by other operations.




Figure 16-6. RAM AIR TEMP Gage

Figure 16-7. Autopilot Flight Controller

Control Switches—Control of the autopilotand the yaw damper is by means of the AUTOPI-LOT MASTER switch on the pilot’s subpanelon SNs 23-003 through 23-099. On SNs 24-100and subsequent, when airplane power is on,the yaw damper receives power directly throughthe AFCS circuit breakers. On 23 models, athree-position YAW DAMPER ON/OFF YAWLOCK switch is located on the pilot’s sub-panel. Yaw lock, when selected, energizes bothservo clutches (motor not running). This main-tains the rudder position at the pilot’s selectedposition for a yaw damper failure. On all model24 and 25 airplanes, the yaw damper engage/dis-engage buttons are located on the flight con-troller.

Wheel Master Buttons—The wheel master but-tons are momentary switches on all models(Figure 16-8). When either switch is depressed,autopilot roll, pitch, and yaw axes disengage.

Maneuver Control Button (MANUV R/P)—The maneuver control buttons are momentaryswitches mounted on the right horn of the pilot’scontrol wheel and on the left horn of the copi-lot’s control wheel. Holding this switch de-pressed momentarily disengages the roll andpitch axes of the AFCS and allows the airplaneto be manually flown with the control wheel.

If a new attitude is established and the maneu-ver switch is then released, the roll and pitchaxes of the AFCS automatically reengage, main-taining the new pitch attitude (within limits) andlevels the wings. If maneuver limits are ex-ceeded and the switch is released, the autopi-lot will return the airplane to an attitude withinthe maneuver limits, wings level, and 15° pitch.Previously engaged roll and pitch axes modeswill be disengaged (HDG, ALT, and NAV, G/S,REV CRS, 1/2° Bank, SPD). If flying faster than0.78 MI (0.79 MI on 24 series airplanes), the au-topilot warning horn will sound with theMANUV R/P switch engaged.

Pitch Sync Switch (not installed on Model23)—When the pitch sync switch between thethrust levers (labeled “A/P PITCH REL” onairplanes with go-around [GA] switch) is de-pressed, autopilot pitch axis will disengageand pitch attitude changes may be made bymoving the control column or trimming. Whenthe pitch sync switch is released, the autopilotwill maintain airplane pitch attitude (withinlimits) existing when the switch is released.Depressing the pitch sync switch will disengageSPD, ALT, or G/S ARM modes if previously se-lected. The overspeed warning horn will soundif the pitch sync switch is depressed at speedsabove 0.78 MI (0.79 MI 24 series).




WHEELMASTERSWITCH

FOUR-WAYTRIM

SWITCH

MICROPHONESWITCH (NOT

SHOWN)

MANEUVERCONTROLSWITCH

FD 108/FD 109PITCH SYNCSWITCH

Figure 16-8. Control Wheel Switches (Pilot’s)

Pitch Trim Monitor Switch—This switch is in-stalled on the pedestal switch panel and isused to test the autopilot disengage functionwhen an improper electrical signal is intro-duced to the pitch function. The switch hasthree positions—DOWN, OFF, and UP, and isspring-loaded to the OFF position.

Roll Monitor Switch (when installed)—Thisswitch is installed on the pedestal switch paneland is used to test the autopilot disengage func-tion when an improper electrical signal is in-troduced to the roll function. The switch has twopositions—hold to TEST and a spring-loadedOFF position.

Autopilot Roll Monitors—The roll monitorswill disengage the autopilot if the bank angleexceeds 40° or if the roll rate exceeds 20° persecond for more than approximately one-halfsecond. When either monitor disengages theautopilot, the disengage tone will sound.

Autopilot Pitch Trim Monitor—The autopilotmaintains pitch trim using the airplane’s emer-gency pitch trim motor. Whenever the autopi-lot is engaged and secondary trim runs in adirection opposite the elevator servo force, themonitor will disengage the autopilot and the dis-engage tone will sound.

Autopilot Pitch Monitor (24 Series)—With theautopilot engaged, the autopilot pitch moni-tor will cause the elevator to streamline (nullservo force) whenever the g level reaches 0.45g. The pitch axis does not disengage but willmaintain the elevator streamlined. Previouslyengaged pitch modes will remain engaged.When the airplane is again above the g limit,the pitch axis will resume elevator inputs.

Autop i l o t /S t i ck Nudger /S t i ck Pu l l e rInterface—If the autopilot is engaged and thestick nudger or puller actuates, any selectedpitch mode will disengage and the autopilotwill maintain a synchronous standby modeuntil the stick nudger or puller releases. Whenthe stick nudger or puller releases, the autopi-lot will maintain the existing pitch attitude.

Autopilot Effort Indicators—An autopilot ef-fort monitor senses the autopilot output signalsfor pitch, roll, and yaw. These are displayed on

the autopilot effort indicators (Figure 16-9)labeled “EL” (pitch), “AIL” (roll), and “RUD”(yaw). Deflection from indexed positions in-dicates that force is being applied to the asso-ciated autopilot servoactuator.

Figure 16-9. Autopilot Effort Indicators

FLIGHT CONTROLLEROPERATIONLights—The autopilot amber OFF light is il-luminated anytime the autopilot has DC poweravailable and is not engaged. The autopilotgreen ON light is illuminated only when the au-topilot pitch and/or roll axes are engaged. Whenthe engage button or any mode button is de-pressed, a light inside the button will illuminateand extinguish when the button is disengaged.

Engage Button—The autopilot ENG buttonengages the pitch and/or roll axes and main-tains heading using the pilot’s HSI (called theHDG hold function) and maintains wings level.Commands to the AFCS are initiated by the rollturn knob and pitch command wheel. The au-topilot can be disengaged by any of the follow-ing buttons or switches, and a tone will soundindicating disengagement.

• Either four-way trim switch (up or down)(UP, DOWN, LWD, RWD with the arm-ing button engaged on airplanes withtwo-speed trim system)

• Either wheel master button







• Emergency pitch trim switch (up ordown)

• Pilot’s VG erect button

• Pilot’s DG slave/free switch (FREE po-sition)

• Trim monitor test switch

• Roll monitor test switch

• Pitch trim selector switch in the OFF po-sition

Turn Knob—The turn knob must be in detentbefore the autopilot will engage. Once en-gaged, it will command bank angles up to 30°,but will disengage any horizontal modes(HDG, NAV, REV CRS) that were engaged.

Pitch Command Wheel—When the pitch com-mand wheel on the autopilot controller is rotated,the autopilot will change airplane pitch attitudein response to the pitch command rotation. Pitchattitude change rate is approximately 1° persecond. Pitch command authority is limited to10° nosedown and 20° noseup. Rotating thepitch command wheel will disengage SPD, ALT,or G/S ARM modes if they were previouslyselected.

Navigation Mode—The navigation (NAV) modebutton is pressed to activate the autopilot func-tion, which captures and tracks navigationalradio beams such as VOR and ILS. For NAVmode functions, the No. 1 navigation receivermust be tuned to the frequency of the selectednavigation aid, and the course pointer on the HSIset to the desired course. Activation of the NAVmode will cause the autopilot to intercept andtrack the desired course.

REV CRS (Reverse Course Select)—For REVCRS mode operation, the navigation receivermust be tuned to a localizer frequency, and theNAV mode must be engaged. When the REVCRS button on the autopilot controller is de-pressed, the localizer course information to theautopilot is reversed, and the glide-slope signalto the autopilot is locked out. REV CRS shouldbe used for back course localizer approachesonly. The REV CRS function is inoperative onVOR frequencies, and disengaging the NAVmode also disengages the REV CRS function.When NAV and/or REV CRS buttons are de-

pressed, the HDG mode will disengage. Turningthe turn knob out of detent will disengage NAVand/or REV CRS modes.

Heading Mode—The heading (HDG) modebutton is pressed to activate the heading selectfunction of the autopilot. This mode com-mands the autopilot to turn the airplane asnecessary and fly a heading selected by the po-sition of the heading bug on the HSI. Maximumbank angle available using this mode is 28°.This button will cancel NAV and REV CRSmodes if engaged. Turning the turn commandknob out of detent will disengage heading(HDG) select mode.

Altitude Hold Mode—The altitude hold (ALT)mode is selected to maintain the existing baro-metric altitude. Using the pitch commandwheel will disengage altitude hold.

G/S ARM (Glide-slope Select)—When the G/SARM button on the autopilot controller is de-pressed, the autopilot will capture and track theILS glide-slope signal. For G/S ARM to func-tion, the navigation receiver must be tuned to alocalizer frequency, an active glide-slope signalmust be present, and REV CRS mode must notbe engaged.

a. If ALT mode is not engaged, and theG/S ARM button is depressed, the glide-slope function will engage if any glide-slope signal is present.

b. If ALT mode is engaged and the G/S ARMbutton is depressed, the glide-slope func-tion will not engage until the glide-slopepointer on the horizontal situation indi-cator (HSI) is in the neutral (glide-slopecentered) position. When the glide-slopecenter is intercepted and the glide-slopefunction engages, the previously selectedALT mode will disengage.

1/2 Bank Mode—1/2 bank mode is applicableto airplanes SNs 25-270, 25-272, and subse-quent. The push–on/push–off 1/2 BANK but-ton on the autopilot controller is used inconjunction with the HDG mode to limit bankangles to half their normal value. The 1/2 bankmode has no effect on the turn command orNAV mode.

Speed (SPD) Hold Mode—The speed (SPD)hold mode, when pressed, will maintain the air-plane at the speed existing when the SPD modewas engaged. Above 29,000 feet, the autopi-lot will maintain Mach number indicated;below 29,000 feet, the autopilot will main-tain indicated airspeed SNs 25-270, -272, andsubsequent.

STABILITY SYSTEM (AFC/SS)The purpose of the AFC/SS system is to allowthe autopilot to automatically trim the airplanein the pitch axis using EMER pitch trim at thehigher Mach numbers to help prevent highspeed tuck.

The AFC/SS consists of a pressure switch inthe copilot’s pitot system that activates a warn-ing horn when 0.78 MI (0.79 MI, 24 series) isexceeded and the autopilot is not engaged.

FC 110 AFCS EMERGENCYCONDITION PROCEDURESRefer to Pitch Axis Malfunction and Roll orYaw Axis Malfunction procedures, AFM,Section III.

FLIGHT DIRECTORSYSTEM

GENERALThe most common flight director installationis the Collins FD 108. Each flight directorsystem includes an attitude director indicator(ADI), a horizontal situation indicator (HSI),a mode selector panel, annunciators, andcourse and heading select controls located onthe HSI. In addition, the flight director has apitch sync switch located on the controlwheels.

The basic attitude reference mode is ener-gized when AC and DC power is applied to theairplane, but with no modes selected on themode selector panel. It provides indication ofairplane heading on the HSI and roll and pitchattitude and sideslip on the ADI. The ADIcommand bars (V-bars) are biased out of view.

The flight director system provides visualcommands to the pilot to manually maintaina desired altitude, capture and maintain a de-sired heading, capture and maintain a desiredVOR course, and capture and maintain an ILScourse and glide path.




HDG

NAVARM

NAVCAPT

REV

GSARM

GSCAPT

EXT

GA

MODE SELECTOR

FLT

SEL

ON

ALT HDG NAVLOC APPR

Figure 16-10. Flight Director Panels




10 10

2020

200

100

10 10

H D G

MILESCOURSE

5.2 01 1

COURSE

ATTITUDEDISPLAY

HORIZONLINE

GLIDE-SLOPEPOINTER

AIRPLANESYMBOL

RUNWAYSYMBOL

BANK POINTER

BANK INDEX

RADARALTITUDE

COURSEDISPLAY

HEADINGMARKER (BUG)

AZIMUTHCARD

LATERALDEVIATIONBAR (CDI)

DISTANCEDISPLAY

COURSEARROW

TO–FROMARROW

GLIDE-SLOPEPOINTER

AIRPLANESYMBOL

HEADINGKNOB

COURSEKNOB

COMMANDBARS

INCLINOMETER

N

33

30W

24

21 S

15

12E

6

3

Figure 16-11. ADI and HSI Symbology




Figures 16-10 and 16-11 show the flight direc-tor mode selector and annunciators and theflight director symbology.

ATTITUDE DIRECTORINDICATOR (ADI)The ADI provides a presentation of the airplaneattitude, localizer and glide-slope deviation,bank angle, and airplane slip or skid. When aflight director mode has been selected, commandbars on the ADI appear in view to provide thecomputed roll and pitch commands. These barsmove up or down to command pitch, and moveclockwise or counterclockwise to commandbank. The pilot must maneuver the airplane sothat the airplane symbol is “flown into” thecommand bars and the two are aligned to sat-isfy the commands.

HORIZONTAL SITUATIONINDICATOR (HSI)The HSI displays airplane position and head-ing with respect to magnetic north, selectedheading, slant range DME distance from a se-lected station, lateral deviation from a se-lected VOR or localizer course, and verticaldeviation from the center of a glide-slopebeam. Two controls are located on the HSI: theheading (HDG) knob and the COURSE knob(Figure 16-11).

Rotating the HDG knob sets the headingmarker (bug) to the desired heading as read onthe azimuth card. The COURSE knob rotatesthe course arrow to a magnetic heading and/orselects a VOR or localizer course.

MODE SELECTORThe mode selector includes push–on/push–offswitches for altitude (ALT), heading (HDG),navigation localizer (NAV LOC), and approach(APPR) modes (see Figure 16-10).

OPERATIONALT Mode—Altitude hold is selected whenthe ALT button is pushed ON and HDG,NAV/LOC, or APPR mode has been selected.

If selected mode(s) is subsequently disen-gaged, the command bars will stow. Pitch holdis selected when the ALT button is pushed off.

HDG Mode—When the HDG selector ispushed ON, the command bars will displaysteering commands selected by the headingmarker on the HSI.

NOTEWhen heading changes of greater than135° are desired, begin by moving theheading marker on the HSI 135° in thedirection of the desired turn. When theairplane completes 120° of the de-sired turn, select the heading markerto the desired heading.

NAV/LOC Mode—Pushing the NAV/LOC but-ton ON engages the NAV ARM mode. Thedisplayed steering commands will permit theairplane to follow the heading set by the head-ing marker to produce the desired VOR or lo-calizer intercept angle. A capture point iscomputed relative to deviation from the desiredcourse, closing speed of the airplane to thecourse, and course intercept angle. Whencourse capture is achieved, the NAV ARM an-nunciator will extinguish and the NAV CAPTannunciator will illuminate, and the commandbars will indicate a bank to intercept the se-lected course. The command bars will displaycommands to turn and track the selected radiocourse. Crosswind correction is automaticallycomputed. A cone suppression circuit ensuressmooth station passage for the VOR.

APPR Mode—Localizer capture in the APPRmode is identical with the NAV/LOC mode. Inaddition, the APPR mode provides for glide-slope arm and capture.

When the APPR mode is selected and a validglide-slope signal exists, the GS ARM annun-ciator will illuminate. The glide-slope captureis independent of localizer capture. Glide-slopecapture is possible when the airplane is ap-proaching the glide slope from below or abovethe actual glide slope. Following capture, the GSARM light will extinguish and the GS CAPTlight will illuminate.




ALT

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RVSM SYSTEMRVSM system aircraft are equipped withSTC STO 1195WI manufactured by AvconIndustries.

GENERALThe existing pilot (servo pneumatic or pneu-matic) altimeters have been replaced withIS&S air data display units (ADDUs) thatcombine the function of the basic altimeterwith those of the traditional altitude alerter.The ADDU is a self-sensing unit that has pitotand static system connections. The existing au-topilot air data sensor was removed and re-placed with the analog interface unit (AIU) thatconverts digital data from the altimeters toanalog signals to interface with the existingFC-110 autopilot (Figure 16-12).

The AIU also provides the outputs to otheroptional aircraft systems (i.e., long range NAV,SAT/TAS indicator). The altitude alerter wasremoved and replaced with the air data panel.Th i s pane l cons i s t s o f an ADC1/ADC2switch/annunciator, an AIU FAIL annunciator,a standby altimeter, and the static source staticport switch (Figure 16-13).

The existing pitot-static system pitot tubes,static ports, and plumbing are replaced withRosemount pitot-static probes and plumbing.

References to static ports in the basic AirplaneFlight Manual (AFM) are to be disregarded.References to pitot heads shall mean pitot-static probes. The pressurization static port hasnot changed in this configuration. A standbyaltimeter (pneumatic) has been installed andplumbed to the aircraft existing shoulder staticports.

SYSTEM CONTROLThe following circuit breakers are added toprovide protection for the air data equipment:

ALT1—Is a 2-amp, 28-VDC circuit breaker onthe LH main bus. It protects ADDU 1.

ALT2—Is a 2-amp, 28-VDC circuit breaker onthe RH main bus. It protects ADDU 2.

ADC AIU—Is a 1-amp, 28-VDC ci rcui tbreaker on the RH main bus. It provides pro-tection to the AIU.

ADC AIU—Is a 1-amp, 26-VAC circu i tbreaker on the RH 26 VAC bus. It also providesprotection to the AIU.

STATIC SOURCE—Is a 5-amp, 28-VDC cir-cuit breaker on the LH main bus. It providesprotection to the isolation valves.

STBY ALT—Is a 2-amp, 28-VDC circuitbreaker on the RH main bus. It protects thestandby altimeter vibrator.

INDICATIONSRefer to Figure 16-14 for the fol lowing indications:

ADC1/ADC2—The ADC1/ADC2 switch/an-nunciator captions are displayed in green. Theswitch/annunciator selects the air data sourcefor transponder altitude reporting, altitudepreselect, altitude alerting and air data inputsto the analog interface unit (AIU). The se-lected air data source becomes the active mas-ter ADDU.




Figure 16-13. Air Data Panel

AIU FAIL—The AIU FAIL is an amber cap-tion that illuminates to advise that the analoginterface unit (AIU) is no longer outputtingvalid data.

Static Port Switch—The static port L and R an-nunciators illuminate (amber) to advise whenthe air data static sources are isolated to thepilot Rosemount probe (L illuminated) or tothe copilot probe (R illuminated). In normaloperations, the switch is centered on BOTH,and neither static port annunciator is illumi-nated. This indicates that the air data sourcesare cross coupled.

For the following, refer to Figure 16-15:

A—Is on the ADDU LCD. This annunciator in-dicates that the ADDU is selected as the ac-tive master ADDU. An absence of this symbolindicates that the ADDU is operating as theslave ADDU.

COM—Is on the ADDU LCD. This annunci-ation indicates that the ADDU is no longerreceiving RS-422 data from the other ADDU.Altitude preselect is disabled on the slaveADDU when this occurs.

OFF—Is on the ADDU LCD. When this cap-tion is in view, all other indications will beblank. This indicates an internal ADDU fault.

STBY—Is on the ADDU LCD. This annunci-ator indicates that the ADDU is no longer ap-plying a static source error correction andaltitude reporting information is no longeravailable on this ADDU.

Refer to Figure 16-13 for the following:

VIB OFF—Is on the standby altimeter. The an-nunciation indicates that the standby altime-ter internal vibrator is off.

NOTEThe COM, OFF, and STBY annun-ciations can possibly be cleared bycycling the appropriate circuit break-ers for the affected ADDU.

OPERATING PROCEDURES

ADDUThe ADDU is capable of displaying altitudein feet or meters and barometric scale ininches of mercury (in. Hg) or hectopascals(Hpa). At initial power-up, the ADDU de-faults to the setting that was effective at thelast power interruption.




Figure 16-14. Air Data Panel Indications

Figure 16-15. Air Data Display Unit (ADDU)

Changing the barometric scale from in. Hg toHpa:

• Depress and hold the barometric knobfor more than 4 seconds, but less than 8seconds. Each depression will togglethe baro scale between in. Hg and Hpa.

Changing the altitude readout from feet tometers:

• Depress the barometric knob for morethan 8 seconds. Each depression willtoggle the altitude readout between feetand meters.

NOTEWhen returning from Hpa/meters toin.Hg/feet, depress the barometricknob for more than 8 seconds tochange altitude readout from metersto feet. Depress the barometric knoba second time for more than 4 sec-onds, but less than 8 seconds, to re-turn the barometric scale from Hpato in. Hg.

• Set the barometric pressure byturning the barometric knob oneach ADDU. Momentarily press-ing the barometric knob will set29.92 in. Hg (1,013 Hpa).

NOTEThe master and slave barometric setknobs are totally independent anddifferent units. Barometric settingsare possible between the pilot andcopilot ADDUs.

• Both ADDUs must be selected tothe same units (feet or meters) orthe altitude preselect function willonly opera te f rom the masterADDU.

The altitude position correction—Flaps (0°),gear up, pilot and copilot ADDU charts in the“Performance” section of the AFM Supplement

only applies when the STBY annunciator is il-luminated on the ADDU display. Under nor-mal conditions, the corrections are negligible.The pilot and copilot altitude position correc-tion charts in the AFM Supplement are to be ap-plied under all applicable flight conditions.

Altitude ReportingAltitude reporting may be supplied from ei-ther ADDU. Selecting ADC1 provides alti-tude information from the pilot ADDU foreither transponder. Selecting ADC2 providesaltitude information from the copilot ADDUfor either transponder. The TFR 1–2 transferswitch is on the transponder control panel.Selecting TFR1 transmits altitude informa-tion from the LH transponder supplied by theselected ADDU. Selecting TFR2 transmits al-titude information from the RH transpondersupplied by the selected ADDU.

Autopilot InterfaceEither ADDU may be used to supply inputs tothe AIU and hence the autopilot. SelectingADC1 uses the pilot ADDU as the air data sourcefor the AIU. Selecting ADC2 uses the copilotADDU as the air data source for the AIU. Theselected ADDU will have the active master in-dicator “A” illuminated and will be the unitused for transponder, autopilot air data input,and other auxiliary inputs (Figure 16-15).

Altitude AlerterSelect the desired alerter altitude by rotatingthe ALT SEL knob (Figure 16-15). A clockwiserotation causes the selected altitude to in-crease and a counterclockwise rotation causesthe selected altitude to decrease. Knob selec-tivity will be 100 feet per detent (30 meters inmetric mode).




The slave unit will display the same selectedaltitude as the master (“A” illuminated) and ad-justing its ALT SEL knob will change the se-lected altitude on both the slave and the master,as long as the same units (feet or meters) areselected. If different altitude units of measure(feet or meters) are selected on the ADDUs,the preselect display on the slave ADDUs willblank and its ALT SEL knob will be disabled.

Static Port SwitchThe static port switch is capable of isolatingthe pilot and copilot air data static sources(see Figure 16-14). In the normal position(centered on BOTH), the LH and RH pilot andcopilot static sources are cross coupled andneither the L or R static port switch annunci-ators on the air data panel will be illuminated.

To isolate the pilot static air data source, po-sition the toggle switch to the left, and the Lannunciator illuminates amber. Under thiscondition, the pilot ADDU now receives itsstatic source from the pilot forward static portwhile the copilot ADDU receives its staticsource from the pilot aft static port. Similarly,to isolate the copilot static air data source,position the toggle switch to the right, andthe R annunciator illuminates amber. Underthis condition, the pilot ADDU receives itsstatic source from the aft copilot static pres-sure port, while the copilot ADDU receives itsstatic source from the forward copilot staticpressure port.

NOTEWhen the static port switch is usedto isolate the pilot static air datasource, the copilot Mach/airspeedindicator (MASI) indications shouldbe considered suspect because thecopilot pitot source could be ques-tionable. Similarly, when isolatingthe copilot static air data source, thepilot MASI indications should beconsidered suspect.

Selecting the L or R static port switch couldalso affect altitude position corrections as pre-sented in the “Performance” section of theAFM Supplement. Therefore, the pilot andcopilot position error correction charts shouldnot be used when either the L or R static portswitch positions are selected. However, thestandby altimeter position error chart is stillvalid.

Standby AltimeterThe standby altimeter is plumbed to the air-craft static shoulder ports. Power for the al-t ime t e r l i gh t i ng i s p rov ided f rom theemergency battery. The standby altimeter is ca-pable of displaying altitude in feet and baro-metric scale in in. Hg or Hpa. Refer to the“Performance—Altimeter Position Error:Standby Altimeter” in the Airplane FlightManual Supplement to determine correct al-titude for current flight conditions.

Pitot-Static ProbesAir data sensing is accomplished by utilizingtwo heated Rosemount pitot-static probes.The pilot total pressure is obtained from thepilot side pitot-static probe. The copilot totalpressure is acquired from the copilot sideprobe. The pilot static pressure is acquired bycross coupling the pilot side forward staticport with the copilot side aft static port.Similarly, the copilot static pressure is ac-quired by cross coupling the pilot side aftstatic port with the copilot side forward staticport.

Prior to all flights in RVSM airspace, the air-plane skin surface in the pitot-static probe re-g ion (RVSM cr i t i c a l r eg ion ) and t hepitot-static probes must be inspected for ob-vious damage or deformation. If damage orsurface irregularities are found, the airplaneis not permitted to operate in RVSM airspace.




PerformanceThe following charts in the basic AFM are su-perceded by charts contained in the RVSMsupplement:

• All airspeed and Mach calibration chartsexcept ground airspeed calibrations

• All altitude position correction charts

• All stall speed charts

• Low speed buffet boundary chart

• Takeoff safety speed (V2) charts

• Enroute climb speed schedule

• Approach and landing climb speed charts

• Landing approach speed (VREF) chart




1. Pitot and static drain valves are located:A. In the nose wheel wellB. On the skin below the entrance doorC. On the nose section near the nose

gear wheel well doorsD. On the nose section at the lower cen-

terline

2. The pilot’s pitot system supplies im-pact pressure to the:A. Mach overspeed switch and

Mach/airspeed indicatorB. Cabin pressure controller and

Mach/airspeed indicatorC. Mach/airspeed indicator and autopi-

lot controllerD. Mach/airspeed indicator

3. An alternate static source is selectablefor the:A. Autopilot altitude controllerB. Pilot’s flight instruments onlyC. Mach airspeed overspeed switchD. Copilot’s flight instruments only

4. When the FC 110 manual pitch com-mand (thumbwheel) is operated, it willdisengage:A. The altitude hold modeB. All available modesC. The heading and altitude hold

modesD. The heading mode

5. In accordance with the AFM supple-ments, the FC 110 must be:A. Engaged when flying at speeds

above Mach .78 or .79B. Disengaged before using the spoilersC. Flown with altitude hold off in se-

vere turbulence

D. Operated as outlined in A, B, or C

6. The basic attitude reference mode isenergized when:A. AC and DC power is applied to the

airplane.B. The AUTO PILOT engage button is

depressed.C. The TEST switch is depressed.D. Both A and B above

7. When using autopilot back coursemode, the:A. Reciprocal of the front course must

be set in the HSI course window.B. Glide-slope receiver signal is cap-

tured.C. Published inbound (front) course

must be set in the HSI course win-dow.

D. Both B and C are correct.

8. To control the airplane in the pitch axis, the autopilot uses the:A. Pitch servo onlyB. Pitch servo and trim tabs on the

elevatorsC. Horizontal stabilizer trim actuator

onlyD. Pitch servo and emergency pitch trim

motor

9. When the pedestal disconnect lever isdepressed and held, it willA. Completely disengage the autopilotB. Disengage the pitch and roll axes

while the lever is depressedC. Disengage the pitch axis while the

lever is depressedD. Disengage the roll axis while the lever

is depressed




QUESTIONS

17-i

CHAPTER 17MISCELLANEOUS SYSTEMS

CONTENTS

Page

INTRODUCTION ................................................................................................................. 17-1

GENERAL ............................................................................................................................ 17-1

OXYGEN SYSTEM.............................................................................................................. 17-2

General........................................................................................................................... 17-2

Oxygen Cylinder............................................................................................................ 17-5

Overboard Discharge Indicator...................................................................................... 17-5

OXYGEN PRESSURE Gage ........................................................................................ 17-6

Crew Distribution System.............................................................................................. 17-6

Passenger Distribution System ...................................................................................... 17-7

Oxygen Duration............................................................................................................ 17-9

DRAG CHUTE ................................................................................................................... 17-10

General......................................................................................................................... 17-10

Preflight ....................................................................................................................... 17-11

Operation ..................................................................................................................... 17-11

QUESTIONS....................................................................................................................... 17-12




17-iii

ILLUSTRATIONS

Figure Title Page

17-1 Oxygen Schematic No. 1........................................................................................ 17-2



17-4 Oxygen Cylinder and Overboard Discharge Indicator........................................... 17-5

17-5 OXYGEN PRESSURE Gage................................................................................. 17-6

17-6 Crew Oxygen Mask................................................................................................ 17-6

17-7 OXY–MIC Panels (Typical)................................................................................... 17-7

17-8 Passenger Oxygen System Controls....................................................................... 17-8

17-9 Passenger Mask...................................................................................................... 17-8

17-10 Passenger Oxygen Compartment Door Manual Release ....................................... 17-9

17-11 Tailcone Access Door and Drag Chute ................................................................ 17-10

17-12 Drag Chute Riser Attachment Hook .................................................................... 17-10

17-13 Drag Chute Handle .............................................................................................. 17-10

TABLE

Table Title Page

17-1 Average Time of Useful Consciousness................................................................. 17-9




INTRODUCTIONMiscellaneous systems used on the Lear 20 series include the oxygen system and the dragchute. The airplane uses high-pressure oxygen stored in a cylinder located in the right nosesection or the dorsal fin. The drag chute is offered as optional equipment.

GENERALThe 20 series oxygen system consists of thecrew distribution system and the passengerdistribution system. Oxygen is available to thecrew whenever the shutoff valve is open and canbe made available to the passengers eitherautomatically above 14,000 feet cabin altitude ormanually at any altitude by the cockpit controls.The system is primarily designed for use in theevent of rapid decompression or pressurization

system failure. It is not designed for plannedextended unpressurized flight at high cabinaltitudes requiring the use of oxygen.

The optional drag chute is used to improvedeceleration on the ground. It is most effectivewhen deployed at higher speeds, but it can still beeffective when deployed at speeds below 60 knots.

1612

8

40

NO 1 FUELTRANS

NO 1 FUELLOW

BATTHOT

ANTI-ICEON

OIL

BLOWEROFF

ENG 1CHIP

NO 1 FUELFILTER

NO 1 BATTSYS

XMSNOIL

90° BOXOIL

GEN 1HOT

TEST

RESET

CHAPTER 17MISCELLANEOUS SYSTEMS




OXYGEN SYSTEM

GENERALThe oxygen system components include anoxygen storage cylinder, a cylinder shutoff

valve and regulator assembly, an overboarddischarge indicator, an oxygen pressure gage,and the distribution systems for the crew andpassengers. Figures 17-1 through 17-3 depictthe crew and passenger oxygen distributionsystems instal led in the Lear 20 seriesairplanes.




MANUAL SHUTOFF VALVE

REGULATOR

RELIEF VALVE(70 PSIG)

FILLER VALVE

CREW OXYGEN MASK

SOLENOID VALVENORMALLY CLOSED

ANEROID SWITCH

MANUAL DOORRELEASE*

MASK STORAGE DOORRELEASE VALVE

PASSENGER OXYGEN MASK

OXYGEN CYLINDER

CYLINDER RELIEF VALVE(3,350 PSIG)

DISCHARGE INDICATOR

RELIEF VALVE(275 PSIG)

PRESSURE GAGE(ON INSTRUMENT PANEL)

CREW OXYGEN MASK

PASS OXY VALVEAUTO—OFFNORMALLY OPEN (AUTO)

PASS MASK VALVEMAN—AUTONORMALLY CLOSED (AUTO)

LANYARD

MASK STORAGE DOOR

LANYARD-OPERATED VALVE

* NOT INSTALLED ON SNs 23-003 THROUGH 24-138

23-003 THROUGH 24-20025-003 THROUGH 25-038

EFFECTIVITY:

SUPPLY PRESSURE

REGULATED PRESSURE

LEGEND

Figure 17-1. Oxygen Schematic No. 1




FILLER VALVE

SOLENOID VALVE

ANEROID SWITCHNORMALLY CLOSED

MANUAL DOORRELEASE

MASK STORAGE DOORRELEASE VALVE

PASSENGER MASK

OXYGEN CYLINDER

MANUAL SHUTOFF VALVEAND REGULATOR(NORMALLY OPEN)

DISCHARGE INDICATOR


CREW OXYGEN MASK



MASK STORAGE DOOR

LANYARD

* PROVISION FOR OPTIONAL OXYGEN MASKS

24-201 THROUGH 24-28825-039 THROUGH 25-182

EFFECTIVITY:

SUPPLY PRESSURE

REGULATED PRESSURE

LEGEND

*

*

*

*

FLOW DETECTOR





* PROVISIONS FOR OPTIONAL OXYGEN MASKS


EFFECTIVITY:

SUPPLY PRESSURE

REGULATED PRESSURE

LEGEND

**

*

*

MANUAL SHUTOFF VALVEAND REGULATOR

(NORMALLY OPEN)

FILLER VALVE

CREW OXYGEN MASK

SOLENOID VALVE

ANEROID SWITCHNORMALLY CLOSED

DOOR LATCH

LANYARD ACTUATORVALVE

FLOW DETECTOR

OXYGEN CYLINDER

DISCHARGE INDICATOR


CREW OXYGEN MASK



PASSENGER MASK

STORAGECOMPARTMENT

LANYARD


OXYGEN CYLINDERThe system is supplied with oxygen from astorage cylinder located in the right nose sectionon 23 and 24 models and model 25 through SN25-024.

Except as noted above, the oxygen cylinder is located in the dorsal fin on all 25 models(Figure 17-4).

Each oxygen cylinder has a storage capacity of38 cubic feet at 1,800 psi. System pressureshould be between 1,800 and 1,850 psi. Theshutoff valve and pressure-regulator assembly isattached to the storage cylinder and provides forpressure regulation and servicing. Oxygenpressure for the passenger and crew distributionsystem is regulated at 60 to 80 psi. The cylinder,along with its shutoff valve and regulatorassembly, can be reached through an accessdoor. Under normal conditions, this valve shouldalways be left in the on (open) position.

The pilot should be aware that if the oxygencylinder shutoff valve is closed, oxygen pressurewill still be read on the OXY PRESS gage in thecockpit. During the interior preflight inspection,ensure that the shutoff valve is open by checkingfor oxygen flow through both crew oxygenmasks, using the 100%, or EMER, position.

OVERBOARD DISCHARGEINDICATORThe overboard discharge indicator (greenblowout disc) (Figure 17-4) provides the pilotwith a visual indication that there has not beenan overpressure condition in the oxygen storagecylinder. The disc blows out if the cylinderpressure reaches 2,700 to 3,000 psi, releasing alloxygen pressure. The green blowout disc islocated on the right side of the dorsal fin or thelower right side of the nose section.




Figure 17-4. Oxygen Cylinder and Overboard Discharge Indicator

OXYGEN PRESSURE GAGEThe OXYGEN PRESSURE gage (Figure 17-5)provides a direct reading of oxygen cylinderpressure, which is necessary to ensure that anadequate supply of oxygen is aboard. The gage,located at the top of the copilot’s instrumentpanel, is marked as follows:

• Yellow arc ............................. 0–300 psi

• Green arc .....................1,550 –1,850 psi

• Red line .................................. 2,000 psi

CREW DISTRIBUTION SYSTEMThe crew distribution system (see Figures 17-1through 17-3) consists of the pilot’s and copilot’soxygen masks with mask-mounted regulators fordiluter-demand or 100% operation. Oxygen isavailable to the crew any time the storage bottleshutoff valve is open and the masks are plugged in.

The crew masks (Figure 17-6) are stowed on thepilot’s and copilot’s sidewalls. The mask oxygenlines are connected to quick-disconnectreceptacles located on the cockpit sidewalls.Optional oxygen-flow detectors may be installedin the mask oxygen lines.

NOTEHeadsets, eyeglasses, or hats worn bycrewmembers may interfere with thequick-donning capabilities of theoxygen mask.

On 24E/F-350 and subsequent and 25D-227 andsubsequent, crew masks are pressure demandwith normal, 100%, and pressure positions thatcan be selected by the crew.

NOTEUse oxygen if pressurization irregu-larities are encountered in flightsabove 10,000 feet.

Following are descriptions of three oxygenmasks/regulators commonly used on Lear 20aircraft. Consult the applicable AFM forspecifics on mask/regulator operation andoxygen duration charts.

The ZMR 100 diluter-demand mask regulatorhas a NORMAL–100% oxygen selector lever.With NORMAL selected, the regulator deliversdiluted oxygen, on demand, up to 20,000 feet




Figure 17-5. OXYGEN PRESSURE Gage

Figure 17-6. Crew Oxygen Mask

cabin altitude. Above 20,000 feet, 100% oxygenmust be selected. With the selector in the 100%position, 100% oxygen is delivered at any cabinaltitude. The 100% position should be usedwhen smoke or fumes are present in thepressurized compartment.

The Robertshaw regulator has two controls–theNORM–EMERGENCY selector on thepressure-regulator control, and the 100% leveron the diluter control. With NORMAL selected,the regulator delivers diluted oxygen, ondemand, up to 30,000 feet cabin altitude. Above30,000 feet, the regulator automatically delivers100% oxygen and maintains a slight positivepressure. With EMERGENCY selected (at anyaltitude) and the 100% lever depressed, theregulator delivers 100% oxygen and maintains aslight positive pressure for respiratory protectionfrom smoke and fumes.

The Puritan-Bennett regulator has two controlson the pressure-regulator control—theNORM–EMERGENCY selector and the 100%lever. With NORM selected, the regulatordelivers diluted oxygen, on demand, up to33,000 feet cabin altitude. Above 33,000 feet,the regulator automatically delivers 100%oxygen. At 39,000 feet it provides positive-pressure breathing. To obtain 100% oxygen atany time, 100% must be selected on thepressure-regulator control. With EMER selected,the regulator delivers 100% oxygen andmaintains a slight positive pressure in the maskcup at all times for respiratory protection fromsmoke and fumes. At 39,000 feet, automaticpressure breathing is supplied.

Each mask assembly includes a microphone andhas an electrical cord which is plugged into theOXY–MIC jack on the respective OXY–MICpanel (Figure 17-7) located on each side panel.To operate the mask microphone, the OXY–MICswitch must be in the ON position and themicrophone keyed, using the microphone switchon the outboard switch on the outboard horn onthe control wheel. Communication betweencrewmembers can be accomplished by using theINPH function of the audio control panel andincreasing the MASTER VOL level.

PASSENGER DISTRIBUTIONSYSTEMThe passenger distribution system (see Figures17-1 through 17-3) is used to provide oxygen tothe passengers in case of a pressurization systemfailure or any other time that oxygen is required.Oxygen is available in the crew oxygendistribution lines whenever the oxygen cylindershutoff valve is open; however, oxygen is notavailable to the passenger distribution systemuntil required.

Oxygen supply to the passengers’ system iscontrolled with three valves. Two valves aremanually operated with control knobs on the




Figure 17-7. OXY–MIC Panels (Typical)

pilot’s sidewall (Figure 17-8), and the third issolenoid-operated by an aneroid switch. Themanually controlled PASS OXY (forward) valveis normally in the NORM (open) position, whichallows oxygen up to the manually controlledPASS MASK (aft) valve and to the aneroid-controlled solenoid valve. Oxygen can beadmitted to the passenger distribution systemthrough either of these passenger mask valves,both of which are normally in the closed position.

With the PASS OXY valve in the OFF (closed)position, oxygen will not be available to thepassenger distribution system in any event. Thisposition may be used only when no passengersare being carried.

With the PASS OXY valve in the NORM (open)position, oxygen will be automatically admitted tothe passenger distribution system through theaneroid-controlled solenoid valve if the cabinreaches 14,000 ±750 feet. The aneroid switchopens the solenoid valve and deploys the passengermasks. On SNs 24-256 and subsequent and 25-095and subsequent, it also illuminates the cabinoverhead lights.

In the event of aircraft electrical failure, automaticdeployment of the passenger masks is not possible.The oxygen solenoid valve requires DC powerthrough the OXY VAL circuit breaker forautomatic mask deployment.

With the PASS OXY valve in the NORM (open)position, rotating the PASS MASK valve fromAUTO to the MAN position admits oxygen into thepassenger distribution system and causes thepassenger oxygen masks to drop. This position canbe used to deploy the passenger masks at anyaltitude, but will not cause the cabin overhead lightsto illuminate.




PASS MASK VALVE

PASS OXY VALVE

Figure 17-8. Passenger Oxygen System Controls

ELASTIC STRAP

OXYGEN MASK

REBREATHER BAG

Figure 17-9. Passenger Mask

The passenger oxygen masks (Figure 17-9) arestowed in compartments in the conveniencepanels above the passenger seats. Thecompartments may contain as many as threemasks, depending on the airplane seatingconfiguration. There will be at least one moremask than passenger seats available.

The passenger mask storage compartment doorsare held closed by latches. When oxygen isadmitted into the passenger distribution system,the oxygen pressure causes the door latches(plungers) to open each compartment door.When the doors open, the passenger masks fallfree and are available for passenger use. As thepassenger pulls down on his mask to don it, anattached lanyard withdraws a pin from thesupply valve which releases oxygen into themask breather bag at a restricted, constant-flowrate. The breather bag may seem to inflateslowly, but this is normal.

NOTEThe individual compartment doors canbe opened manually (Figure 17-10) ifselected use is required, but pulling themask lanyard will not start oxygenflow until the PASS MASK valve isturned to MAN.

Should the doors be inadvertently opened fromthe cockpit, oxygen pressure must be bled fromthe passenger distribution system before themasks can be restowed. This is accomplished bypulling one of the passenger mask lanyards afterensuring that the PASS MASK valve is closed(AUTO). If the doors open due to malfunction ofthe solenoid-operated valve, the PASS OXYvalve must be turned off to permit stowage of thepassenger masks.

The compartment doors can be opened manuallyfor mask cleaning and servicing.

No smoking when oxygen is being used.

OXYGEN DURATIONBefore an overwater flight is made, the pilotshould plan oxygen requirements to provide asufficient quantity for all occupants in case of apressurization failure. It may be necessary tocarry additional oxygen on board to ensure thatboth oxygen duration and fuel requirements are met.

The pilot should also remember thatunpressurized flight above 30,000 feet, eventhough oxygen is being used, can result inunconsciousness. Therefore, if a pressurizationproblem is encountered at a high altitude, adescent should be initiated immediately. Table17-1 lists the average time of useful con-sciousness at various altitudes.

WARNING




Figure 17-10. Passenger OxygenCompartment Door ManualRelease

15,000 to 18,000 feet . . . . . . . . . . 30 minutes or more22,000 feet . . . . . . . . . . . . . . . . . . . . . . 5 to 10 minutes25,000 feet . . . . . . . . . . . . . . . . . . . . . . . 3 to 5 minutes28,000 feet . . . . . . . . . . . . . . . . . . . . 2 1/2 to 3 minutes30,000 feet . . . . . . . . . . . . . . . . . . . . . . . 1 to 2 minutes35,000 feet . . . . . . . . . . . . . . . . . . . . 30 to 60 seconds40,000 feet . . . . . . . . . . . . . . . . . . . . 15 to 20 seconds45,000 feet . . . . . . . . . . . . . . . . . . . . . 9 to 15 seconds

Table 17-1. AVERAGE TIME OFUSEFUL CONSCIOUSNESS

For cabin altitudes of 10,000 feet and above, theoxygen duration times listed in the AFM includecabin altitude ascent time from 8,000 feet to thefinal stabilized cabin altitude (assuming a fullycharged system of 1,850 psi). The pilot may usethe following formula to calculate the oxygenduration for a partially charged oxygen system:

System PressureDuration (per AFM) X

1,850

NOTEThis is an estimate only. Flight shouldbe planned to arrive at an altitude notrequiring supplemental oxygen whengage indicates 300 psi.

DRAG CHUTE

GENERALThe drag chute is an optional decelerationdevice. It, like the thrust reverser system, may beused to produce shorter stopping distances. Thegreatest deceleration rate is produced at thehighest speed; however, the chute is stilleffective at speeds as low as 60 knots.

The chute is stored in a removable canisterwhich is mounted inside the tailcone access door(Figure 17-11). The canister lid is released fromthe canister when the drag chute handle ispulled, allowing the pilot chute to deploy. The

pilot chute then pulls the main chute canopy outof the canister.

The main chute riser attaches to the airplane atthe chute control mechanism just forward of thecanister (Figure 17-12). The loop at the end ofthe main riser slips over a recessed metal pinwhich is held in position by spring pressurewhen the drag chute is stowed. Therefore, if thechute should inadvertently deploy (handle instowed position), the main chute riser will slipfree of the pin and separate from the airplane.When the drag chute handle (Figure 17-13) ispulled, the pin is mechanically locked in positionto retain the chute riser. Simultaneousfunctioning of the mechanical canister controlmechanism releases the canister lid, therebydeploying the chute.




Figure 17-13. Drag Chute HandleFigure 17-11. Tailcone Access Door and

Drag Chute

Figure 17-12. Drag Chute RiserAttachment Hook

The drag chute can be used:

• When landing on wet or icy runway

• During any landing emergency involvingno-flap hydraulic or brake failure, or lossof directional control

• During takeoff if the decision is made toabort

Do not deploy the drag chute under thefollowing conditions:

• In flight

• If the nose gear is not on the ground

• When the indicated airspeed is above 150 knots

• With thrust reversers deployed

If chute deployment occurs above 150KIAS or if jettison or failure occursabove 100 KIAS, the drag chutesystem and adjacent structure in thevicinity of the tailcone access doormust be thoroughly inspected fordamage before the next flight.

PREFLIGHTThe following procedures may be accomplishedduring the preflight inspection to check the dragchute for proper installation:

• Open the tailcone access door.

• Pull the chute deploy handle in thecockpit to full extension. Observe thatboth handle grip safeties must bedepressed to pull the handle. Also notethat the chute riser attachment hook locksand that the control mechanism crank andring function.

• Return the chute deploy handle to thestowed (down) position.

• Ensure that the riser attaching loop is insertedin the riser attachment hook and that thehook is unlocked; ensure that the riser passesthrough the loop on the riser keeper.

• Check chute riser for proper routing andstowage as the tailcone access door isclosed.

OPERATIONAs the nosewheel touches down, the copilot, onthe pilot’s command, deploys the drag chute bysqueezing the drag chute control handle andpulling it up to its full extension (a pull force ofapproximately 40 pounds may be required).With the chute deployed, the pilot should keepthe airplane well clear of the runway andtaxiway lights, markers, and obstructions bylanding and taxiing on the upwind side. Taxiingdownwind should always be avoided.

NOTEThe drag chute trails approximately 37feet behind the airplane. If the airplaneis taxiing downwind with a deployeddrag chute, the chute may overtake andbecome entangled with the airplane.

The drag chute can be jettisoned afterdeployment at any time. Normally, the pilotheads the airplane into the wind as much aspossible to jettison the chute after the airplaneclears the runway. The copilot jettisons the dragchute by squeezing the control handle gripsafeties and pushing the handle down to thestowed position to release the chute. If the chutehas collapsed prior to jettisoning, the chute risermust be pulled free after stowing the handle.Because the possibility always exists thatjettisoning the chute might be required duringthe landing roll, any planned deployment shouldbe coordinated with the control tower.

CAUTION




1. During preflight, the pilot can determineif the oxygen bottle is turned on by:A. Reading the pressure indicated on the

oxygen pressure gage in the cockpitB. Selecting 100% on the mask regulators

and taking several deep breaths throughthe mask

C. Placing the OXY–MIC switch to theOXY position

D. Visually checking for the green flowindicator on the mask supply hose

2. With the PASS OXY valve in theNORM position, selecting MAN on thePASS MASK valve:A. Causes passenger masks to drop and

turns on the cabin overhead lightsB. Prevents oxygen from entering the

passenger oxygen distribution linesC. Causes passenger masks to drop, but

does not turn on the cabin overheadlights

D. Disables the solenoid-operated oxy-gen valve

3. With the PASS OXY valve in theNORM position and the PASS MASKvalve in the AUTO position:A. Oxygen is supplied to the passenger

masks if the cabin altitude reaches10,000 feet.

B. Passenger masks will automaticallydeploy in the event of electrical failure.

C. Passenger masks will automaticallydeploy and the cabin overhead lightswill illuminate if the cabin altitudereaches 14,000 feet.

D. The aneroid-controlled passengermask drop valve is disabled.

4. The OXY PRESS gage reads:A. Direct pressure of the cylinderB. Electrically derived system high

pressureC. Direct pressure of the pilot’s supply lineD. Elec t r i ca l ly de r ived sys tem low

pressure

5. The maximum demonstrated crosswindcomponent for drag chute deployment is:A. 10 knotsB. 15 knotsC. 20 knotsD. 25 knots

6. The drag chute is deployed by:A. Squeezing the control handleB. Rotating the control handle fully

clockwise and pulling it up to its fullextension

C. Squeezing the control handle andpulling it up to its full extension

D. Squeezing the control handle andpushing it completely forward

7. The maximum indicated airspeed fordrag chute deployment is:A. 120 knotsB. 130 knotsC. 140 knotsD. 150 knots




QUESTIONS

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WA-3FOR TRAINING PURPOSES ONLY



8. NOSEWHEEL AND TIRE—CONDITION AND NOSE GEAR UPLOCK FORWARD

9. RADOME ALCOHOL DISCHARGE PORT—CLEAR OF OBSTRUCTIONS

10. RADOME AND RADOME EROSION SHOE—CONDITION

12. RIGHT PITOT HEAD—COVERS REMOVED, CLEAR OF OBSTRUCTIONS

11. OXYGEN SUPPLY VALVE—ON

OXYGEN PRESSURE RELIEF DISC—INTACT

14. RIGHT STATIC PORTS—CLEAR OF OBSTRUCTIONS

13. RIGHT STALL WARNING VANE—FREEDOM OF MOVEMENT, LEAVE IN DOWN POSITION

②

③

WA-4 FOR TRAINING PURPOSES ONLY



15. RIGHT PITOT-STATIC DRAIN VALVES (2)—DRAIN

16. RIGHT SHOULDER STATIC PORT—CLEAR OF OBSTRUCTIONS

17. COPILOT’S WINDSHIELD DEFOG OUTLET—CLEAROF OBSTRUCTIONS

19. EMERGENCY EXIT—SECURE

20. UPPER FUSELAGE ANTENNAE—CONDITION

18. LOWER FUSELAGE ANTENNAE, ROTATING BEACON LIGHT AND LENS—CONDITION

21. ROTATING BEACON LIGHT AND LENS (ON VERTICAL FIN)—CONDITION

④




22. RIGHT ENGINE INLET—CLEAR OF OBSTRUCTIONSAND CONDITION

25. RIGHT MAIN GEAR LANDING LIGHT—CONDITION

26. RIGHT MAIN GEAR WHEELS, BRAKES, AND TIRES—CONDITION

23. FUEL CROSSOVER, LEFT WING SUMP, LEFT ENGINE FUEL, RIGHT WING SUMP, AND RIGHT ENGINE FUEL DRAIN VALVES —DRAIN

27. STALL STRIP (WING LEADING EDGE) AND STALLFENCE—CONDITION

24. RIGHT MAIN GEAR AND WHEEL WELL—HYDRAULIC/FUEL LEAKAGE AND CONDITION

28. RIGHT WING ACCESS PANELS (UNDERSIDE OFWING)—CHECK FOR FUEL LEAKAGE




29. RIGHT FUEL VENT (UNDERSIDE OF WING)—CLEAR OF OBSTRUCTIONS

30. VORTEX GENERATORS OR BOUNDARY LAYERENERGIZERS—CONDITION

32. RIGHT TIP TANK—CONDITION

31. RIGHT WING HEAT SCUPPER (UNDERSIDE OF WINGFORWARD)—CLEAR OF OBSTRUCTIONS

34. RIGHT TIP TANK SUMP DRAIN VALVE(S) (TWO ONAIRPLANES PRIOR TO 24-181 AND 25-070)—DRAIN

33. RIGHT TIP TANK RECOGNITION LIGHT AND LENS—CONDITION

⑤




35. RIGHT TIP TANK FUEL CAP—CONDITION AND SECURE

36. RIGHT TIP TANK NAVIGATION LIGHT, STROBE LIGHT, AND LENS—CONDITION

39. SCUPPER (UNDERSIDE OF RIGHT WING AFT)—CLEAR OF OBSTRUCTIONS, NO FUEL LEAKAGE

37. RIGHT TIP TANK FIN AND STATIC DISCHARGE WICKS (2)—CONDITION

40. RIGHT AILERON—CHECK FREE MOTION, BAL-ANCE TAB LINKAGE, BRUSH SEAL CONDITION

38. RIGHT TIP TANK FUEL JETTISON TUBE—CLEAR OFOBSTRUCTIONS

41. RIGHT SPOILER AND FLAP—CONDITION




45. FUEL VENT DRAIN VALVE, TRANSFER LINE DRAINVALVE, FUSELAGE TANK SUMP DRAIN VALVE—DRAIN

42. RIGHT ENGINE OIL TANK FILLER CAP AND ACCESSDOOR—SECURE

43. RIGHT ENGINE THRUST REVERSER— CONDITIONAND STOWED

46. LEFT AND RIGHT FUEL FILTER DRAIN VALVES—DRAIN

47. FUEL VENT RAM-AIR MAST (DRAG CHUTE IN-STALLED)—CHECKED

44. R IGHT ENGINE TURBINE EXHAUST AREA—CONDITION, CLEAR OF OBSTRUCTION

⑥




48. FUEL VENT RAM-AIR MAST (WITHOUT DRAGCHUTE INSTALLED)—CHECKED

49. TAILCONE ACCESS DOOR—OPEN 52. TAILCONE INTERIOR—CHECK FOR FLUID LEAKS,SECURITY, AND CONDIT ION OF INSTALLEDEQUIPMENT

50. TAILCONE INTERIOR—CHECK FOR FLUID LEAKS,SECURITY, AND CONDIT ION OF INSTALLED EQUIPMENT

51. TAILCONE INTERIOR—CHECK FOR FLUID LEAKS,SECURITY, AND CONDIT ION OF INSTALLEDEQUIPMENT

53. DRAG CHUTE—CHECK FOR PROPER INSTALLATION




54. DRAG CHUTE—CHECK FOR PROPER INSTALLATION

57. OXYGEN SERVICING DOOR—SECURE

58. OXYGEN DISCHARGE DISC—CONDITION

55. TA ILCONE ACCESS DOOR—CLOSED ANDSECURED

56. OXYGEN BOTTLE SUPPLY VALVE—OPEN

65. FIRE-EXTINGUISHER DISCS—CONDITION

59. RIGHT VOR/LOC ANTENNA—CONDITION

60. VERTICAL STABILIZER, RUDDER, HORIZONTALSTABILIZER, AND ELEVATOR—CONDITION,DRAIN HOLES CLEAR

61. STATIC DISCHARGE WICKS (6 ON HORIZONTALSTABILIZER, 1 ABOVE NAV LIGHT, 1 ON VERTI-CAL FIN)—CONDITION

62. VERTICAL FIN NAVIGATION LIGHTS, STROBELIGHT AND LENS—CONDITION

63. VLF H-FIELD ANTENNA—CONDITION

64. LEFT VOR/LOC ANTENNA—CONDITION

⑦




66. LEFT ENGINE THRUST REVERSER—CONDITIONAND STOWED (DEE HOWARD)

69. LEFT AILERON—CHECK FREE MOTION, BAL-ANCE, AND TRIM TABS LINKAGE, AND BRUSHSEAL CONDITION

67. LEFT ENGINE DIPSTICK FILLER CAP AND ACCESSDOOR—SECURE

70. LEFT TIP TANK FUEL JETTISON TUBE ( IF IN-STALLED)—CLEAR OF OBSTRUCTIONS

68. LEFT SPOILER AND FLAP—CONDITION

71. LEFT TIP TANK FIN AND STATIC DISCHARGEWICKS (2)—CONDITION

⑧




72. LEFT TIP TANK NAVIGATION LIGHT, STROBE LIGHTAND LENS—CONDITION

73. LEFT TIP TANK FUEL CAP—CONDITION AND SECURE

76. LEFT TIP TANK—CONDITION

74. LEFT TIP TANK SUMP DRAIN VALVE (TWO ON AIR-CRAFT PRIOR TO SNs 24-181 AND 25-067)—DRAIN

77. LEFT WING HEAT SCUPPER (UNDERSIDE OF WINGFORWARD)—CLEAR OF OBSTRUCTIONS

75. LEFT TIP TANK RECOGNITION LIGHT AND LENS(IF INSTALLED)—CONDITION

78. BOUNDARY LAYER ENERGIZERS(IF INSTALLED)—CONDITION

79. VORTEX GENERATORS—MISSING

➈




83. STALL FENCE (IF INSTALLED)—CONDITION

80. LEFT WING ACCESS PANELS (UNDERSIDE OFWING)—CHECK FOR FUEL LEAKAGE

81. LEFT FUEL VENT (UNDERSIDE OF WING)( IF INSTALLED)—PLUG REMOVED, CLEAR OF OBSTRUCTIONS

84. LEFT MAIN GEAR AND WHEEL WELL—HYDRAULIC/FUEL LEAKAGE AND CONDITION

82. STALL STRIP (WING LEADING EDGE)—CONDITION 85. LEFT MAIN GEAR LANDING LIGHT— CONDITION

86. LEFT MAIN GEAR WHEELS, BRAKES, AND TIRES—CONDITION




87. LEFT ENGINE INLET—CLEAR OF OBSTRUCTIONSAND CONDITION

WA

-15

FO

R T

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59

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60 61 62 63

5554

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APP-i

APPENDIX 15 HB

This appendix contains the following conversion tables:

Table Page

APP-1 Conversion Factors ............................................................................................. APP-1

APP-2 Fahrenheit and Celsius Temperature Conversion ............................................... APP-2

APP-3 Inches to Millimeters (0.001 Inch to 10 Inches) ................................................ APP-3

APP-4 Weight (Mass): Ounces or Pounds to Kilograms ............................................... APP-4

APP-5 Weight (Mass): Thousand Pounds to Kilograms................................................ APP-5




APP-1FOR TRAINING PURPOSES ONLY



Table APP-1. CONVERSION FACTORS

APP-2 FOR TRAINING PURPOSES ONLY



Tab

le A

PP

-2.

FA

HR

EN

HE

IT A

ND

CE

LS

IUS

TE

MP

ER

AT

UR

E C

ON

VE

RS

ION




Table APP-3. INCHES TO MILLIMETERS (0.0001 INCH TO 10 INCHES)

APP-4 FOR TRAINING PURPOSES ONLY



Table APP-4. WEIGHT (MASS): OUNCES OR POUNDS TO KILOGRAMS




Table APP-5. WEIGHT (MASS): THOUSAND POUNDS TO KILOGRAMS



ANN-iFOR TRAINING PURPOSES ONLY

ANNUNCIATOR PANEL

The Annunciator Panel section presents a colorrepresentation of all the annunciator lights inthe airplane.

Please unfold page ANN-1 to the right andleave i t open for ready reference as the annunciators are cited in the text.

AN

N-1

FO

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TAKE

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TRIM

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HTR

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VSP

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EDFI

RETA

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RELO

WHY

DFU

ELXF

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L LO OIL L

RAN

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R LO OIL

L. FU

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FUEL

L STA

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L FI

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R. FU

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IL PR

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BAT 1

40

BAT 1

60

PITO

T HT

L CAB

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R CA

B AI

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L. TI

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TIP

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L. GE

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ARM

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EXT

MON

MON

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GA

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GS ARM

GS CAPT

EXT

GA

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flight safety learjet 20 series pilot training manual volume 2

Documents

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