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King Air C90A/B – The Training Workbook Copyright © 2012

Douglas S. Carmody and Executive Flight Training LLC are not liable for the accuracy, effectiveness or safe use of this workbook and do not warrant that this aircraft manual or publication contains current information and/or revisions. Aircraft manuals and publications required for any reason other than training, study or research purposes should be obtained from the original equipment manufacturer. Reference herein to any specific commercial products by trade name, trademark, manufacturer, or otherwise, is not meant to imply or suggest any endorsement by, or affiliation with that manufacturer or supplier. All trade names, trademarks and manufacturer names are the property of their respective owners. All illustrations are the property of Hawker Beechcraft Corporation and used with permission. Passages and examples reprinted from Beechcraft Hawker Corporation’s C90A maintenance manual, and POH are used with permission. No part of this book may be copied without the expressed written permission of Douglas Carmody. All rights reserved.

Published by Executive Flight Training LLC. Beaufort, SC

Table of Contents Chapter 1 – Airplane (General) INTRODUCTION TO THE KING AIR C-90A/B ................................................... 1 OBJECTIVES......................................................................................................... 1 HISTORY OF THE KING AIR C-90A/B ................................................................ .2 GENERAL............................................................................................................. ..2 NOSE SECTION............................................................................................... 3 COCKPIT ............................................................................................................. 3 LIGHTING SYSTEMS ........................................................................................... 4 CABIN CONFIGURATION .................................................................................... 5 CABIN WINDOWS................................................................................................ 9 EMERGENCY EXIT ............................................................................................ 10 INTERIOR DIVIDERS ......................................................................................... 10 AFT FUSELAGE ................................................................................................. 10 EMPENNAGE ..................................................................................................... 10 WINGS................................................................................................................ 11 ATTACH FITTINGS ............................................................................................ 12 POWER PLANT.................................................................................................. 12 ELECTRICAL SYSTEM ...................................................................................... 12 PROPELLER SYSTEM....................................................................................... 12 FUEL SYSTEM ................................................................................................... 13 ANTI-ICE/DE-ICE SYSTEMS ............................................................................. 13 ENVIRONMENTAL SYSTEM ............................................................................. 13 LIMITATIONS ..................................................................................................... 13 WEIGHT LIMITS ................................................................................................. 14 CENTER OF GRAVITY LIMITS .......................................................................... 14 MANEUVER LIMITS ........................................................................................... 15 FLIGHT LOAD FACTOR LIMITS (9650 POUNDS)............................................. 15 EMERGENCY PROCEDURES........................................................................... 15 GENERAL CHAPTER QUESTIONS................................................................... 16 Chapter 2 – Electrical Systems OBJECTIVES...................................................................................................... 18 ELECTRICAL POWER - DESCRIPTION AND OPERATION ............................. 19 BATTERY SYSTEM............................................................................................ 21 DC GENERATION - DESCRIPTION AND OPERATION .................................... 22 STARTER-GENERATORS ................................................................................. 22 GENERATOR CONTROL UNIT.......................................................................... 23 STARTER-GENERATOR PARALLELING .......................................................... 23 OVERVOLTAGE PROTECTION ........................................................................ 24 REVERSE CURRENT PROTECTION ................................................................ 24 OVER EXCITATION PROTECTION ................................................................... 24 FIELD FLASH CIRCUIT……………………………………………………...………24

COMPONENT LOCATION ................................................................................. 25 BUS TIE SYSTEM……………………………………………………………….……25 AC GENERATION .............................................................................................. 26 EXTERNAL POWER .......................................................................................... 26 AVIONIC MASTER SWITCH .............................................................................. 28 CIRCUIT BREAKERS ......................................................................................... 28 STATIC DISCHARGING - DESCRIPTION AND OPERATION........................... 29 ELECTRICAL SYSTEM LIMITATIONS............................................................... 29 EMERGENCY ELECTRICAL PROCEDURES ................................................... 30 ABNORMAL ELECTRICAL PROCEDURES ...................................................... 31 EXPANDED ELECTRICAL PROCEDURES ....................................................... 31 ELECTRICAL SYSTEM QUESTIONS ................................................................ 32 Chapter 3 – Annunciator System OBJECTIVES...................................................................................................... 35 ANNUNCIATOR SYSTEM .................................................................................. 35 ANNUNCIATOR EMERGENCY PROCEDURES ............................................... 37 ANNUNCIATOR ABNORMAL PROCEDURES .................................................. 37 ANNUNCIATOR SYSTEM QUESTIONS ............................................................ 38 Chapter 4 – Fuel System OBJECTIVES...................................................................................................... 39 FUEL SYSTEM - DESCRIPTION AND OPERATION ......................................... 40 FUEL GAUGES .................................................................................................. 41 FUEL DRAIN VALVES........................................................................................ 42 FIREWALL SHUTOFF VALVES ......................................................................... 42 FUEL VENTS...................................................................................................... 42 FUEL PUMPS ..................................................................................................... 42 FUEL TRANSFER PUMPS................................................................................. 43 FUEL FILTERS ................................................................................................... 44 ENGINE FUEL CONTROL LINE HEATER ......................................................... 45 FUEL HEATER ................................................................................................... 45 CROSSFEED...................................................................................................... 45 FUEL PURGE SYSTEM ..................................................................................... 46 FUEL LIMITATIONS ........................................................................................... 46 FUEL MANAGEMENT ........................................................................................ 47 EMERGENCY FUEL SYSTEM PROCEDURES ................................................ 48 ABNORMAL FUEL PROCEDURES ................................................................... 49 FUEL SYSTEM EXPANDED PROCEDURES .................................................... 50 FUEL SYSTEM QUESTIONS ............................................................................. 50 Chapter 5 - Engine System OBJECTIVES...................................................................................................... 53 GENERAL ENGINE DESCRIPTION................................................................... 53 TURBOPROP ENGINE SYMBOLS AND THEIR MEANINGS ............................ 55 ENGINE FUEL SYSTEM .................................................................................... 61 STARTING AND IGNITION SYSTEM................................................................. 63 AUTO IGNITION ................................................................................................. 63 FIRE DETECTION SYSTEM .............................................................................. 63

FIRE EXTINGUISHING SYSTEM ....................................................................... 64 POWERPLANT LIMITATIONS ........................................................................... 65 EMERGENCY ENGINE PROCEDURES ............................................................ 66 ABNORMAL ENGINE PROCEDURES ............................................................... 69 ENGINE SYSTEM QUESTIONS ........................................................................ 71 Chapter 6 – Propeller System OBJECTIVES...................................................................................................... 74 GENERAL........................................................................................................... 74 BASIC PRINCIPLES ........................................................................................... 75 PROPELLER GOVERNOR ................................................................................ 75 PRIMARY GOVERNOR...................................................................................... 75 LOW PITCH STOP ............................................................................................. 76 SECONDARY LOW PITCH STOP...................................................................... 77 OVERSPEED GOVERNOR................................................................................ 77 FUEL TOPPING GOVERNOR............................................................................ 78 PROPELLER FEATHERING .............................................................................. 78 AUTOFEATHER ................................................................................................. 78 PROPELLER BETA AND REVERSING ............................................................. 79 PROPELLER SYNCHROPHASER ..................................................................... 80 PROPELLER CARE ........................................................................................... 81 LIMITATIONS ..................................................................................................... 81 PROPELLER EMERGENCY PROCEDURES .................................................... 81 PROPELLER ABNORMAL PROCEDURES ....................................................... 82 PROPELLER EXPANDED PROCEDURES ....................................................... 82 PROPELLER SYSTEM QUESTIONS................................................................. 82 Chapter 7 – Pressurization and Environmental Systems OBJECTIVES...................................................................................................... 84 INTRODUCTION ................................................................................................ 84 HEATING, COOLING AND PRESSURIZATION - DESCRIPTION AND OPERATION..................................................................................................... 85 MANUAL HEAT OPERATION ............................................................................ 87 ELECTRIC HEAT................................................................................................ 87 FRESH AIR VENTILATION ................................................................................. 88 COOLING - DESCRIPTION AND OPERATION ................................................. 88 AIR CONDITIONING TEMPERATURE CONTROL - DESCRIPTION AND OPERATION................................................................................................... 89 PRESSURIZATION - DESCRIPTION AND OPERATION ................................ 90 PRESSURIZATION LIMITATIONS ................................................................... 94 EMERGENCY ENVIRONMENTAL SYSTEM PROCEDURES ......................... 94 ABNORMAL ENVIRONMENTAL PROCEDURES ............................................ 96 ENVIRONMENTAL SYSTEM EXPANDED PROCEDURES ............................ 97 OXYGEN DURATION …………………………………………………………….. 98 ENVIRONMENTAL SYSTEM QUESTIONS ..................................................... 98 Chapter 8 – Landing Gear, Tires, and Brake System OBJECTIVES.................................................................................................... 101 GENERAL......................................................................................................... 101

GROUND HANDLING TOWING....................................................................... 102 PARKING.......................................................................................................... 102 NOSE LANDING GEAR.................................................................................... 102 DESCRIPTION AND OPERATION - LANDING GEAR..................................... 103 LANDING GEAR WARNING SYSTEM ............................................................. 106 TIRES ............................................................................................................... 107 HYDRAULIC BRAKE SYSTEM ........................................................................ 108 SHOCK STRUTS.............................................................................................. 109 LANDING GEAR LIMITATIONS ....................................................................... 109 EMERGENCY LANDING GEAR SYSTEM PROCEDURES ............................. 109 ABNORMAL LANDING GEAR PROCEDURES................................................ 109 LANDING GEAR EXPANDED PROCEDURES ................................................ 110 LANDING GEAR SYSTEM QUESTIONS ......................................................... 111 Chapter 9 - Pneumatic and Vacuum System OBJECTIVES.................................................................................................... 113 DESCRIPTION ................................................................................................. 113 PNEUMATIC - DESCRIPTION AND OPERATION........................................... 113 VACUUM SYSTEM - DESCRIPTION AND OPERATION ................................ 114 ENGINE BLEED AIR CONTROL…………. ...................................................... 115 DOOR SEAL SYSTEM ..................................................................................... 115 FLIGHT HOUR RECORDER ............................................................................ 116 PNEUMATIC LIMITATIONS ............................................................................. 116 PNEUMATIC SYSTEM EMERGENCY PROCEDURES ................................... 116 PNEUMATIC SYSTEM ABNORMAL PROCEDURES ...................................... 116 PNEUMATIC SYSTEM EXPANDED PROCEDURES ...................................... 116 PNEUMATIC AND VACUUM SYSTEM QUESTIONS ...................................... 117 Chapter 10 – Anti-Icing System OBJECTIVES.................................................................................................... 118 DESCRIPTION ................................................................................................. 118 ICE AND RAIN PROTECTION - DESCRIPTION AND OPERATION ............... 118 INERTIAL ICE SEPARATION SYSTEM ........................................................... 121 AIR INTAKE ANTI-ICE LIP ............................................................................... 122 ENGINE FUEL CONTROL HEAT ..................................................................... 122 WINDOWS AND WINDSHIELDS ..................................................................... 122 PROPELLER DEICING .................................................................................... 123 PITOT HEAT..................................................................................................... 125 STALL WARNING VANE HEAT ....................................................................... 125 FUEL VENTS.................................................................................................... 125 FUEL HEAT ...................................................................................................... 125 ICING LIMITATIONS ........................................................................................ 126 EMERGENCY ICING SYSTEM PROCEDURES .............................................. 126 ABNORMAL ICING SYSTEM PROCEDURES ................................................. 126 ICING EXPANDED PROCEEDURES ............................................................ 127 ENCOUNTERING ICING CONDITIONS…………………………………………128 ANTI-ICING SYSTEM QUESTIONS ................................................................. 130

Chapter 11– Flight Controls OBJECTIVES.................................................................................................... 132 FLIGHT CONTROLS ........................................................................................ 132 ELEVATOR TRIM ............................................................................................. 133 CONTROL LOCKS ........................................................................................... 134 GROUND MOORING/TOWING........................................................................ 134 WING FLAPS.................................................................................................... 135 YAW DAMPER ................................................................................................. 136 RUDDER BOOST…………………………………………………………………. 136 STALL WARNING SYSTEM ............................................................................. 137 FLIGHT CONTROL LIMITATIONS ................................................................... 137 FLIGHT LOAD FACTOR LIMITS ...................................................................... 137 FLIGHT CONTROL EMERGENCY PROCEDURES ........................................ 137 FLIGHT CONTROL ABNORMAL PROCEDURES ........................................... 138 FLIGHT CONTROL EXPANDED PROCEDURES ............................................ 139 FLIGHT CONTROLS QUESTIONS .................................................................. 140 Chapter 12 – Pitot Static System OBJECTIVES.................................................................................................... 141 PITOT AND STATIC PRESSURE SYSTEM..................................................... 141 OUTSIDE AIR TEMPERATURE ....................................................................... 142 PITOT STATIC LIMITATIONS .......................................................................... 142 PITOT STATIC SYSTEM EMERGENCY PROCEDURES................................ 143 PITOT STATIC SYSTEM ABNORMAL PROCEDURES................................... 143 PITOT STATIC SYSTEM QUESTIONS ............................................................ 143 Chapter 13 – Oxygen System OBJECTIVES.................................................................................................... 145 OXYGEN SYSTEM - DESCRIPTION AND OPERATION................................. 145 OXYGEN LIMITATIONS ................................................................................... 146 OXYGEN EMERGENCY PROCEDURES ........................................................ 146 OXYGEN ABNORMAL PROCEDURES ........................................................... 147 OXYGEN SYSTEM QUESTIONS ..................................................................... 147 Profiles and Power Settings…………………………………………..…149

King Air C-90A/B Training Manual 1

Chapter 1

Airplane –General

INTRODUCTION TO THE KING AIR C-90A/B

This training and informational workbook describes the airframe, engines and systems of the King Air C-90A. It is a compilation of operating information and techniques gathered over 20 years of King Air Training. It covers serial numbers LJ-1063 through LJ-1299. The C-90B covers serial numbers LJ-1302 and subsequent. The following are changes differentiating the C-90B from the C-90A: reduced cabin sound; a gated ground fine power lever position; follow-up type flap selector switch; approach chart holder on pilot’s and copilot’s control wheels; changed chip detect warning annunciation to caution annunciation; and incorporates 26 electronically tuned dynamic vibration absorbers mounted in strategic locations on specific fuselage frames. It is an excellent refresher program but it is intended for training purposes only and is not a substitute for the POH. The Pilot’s Operating Handbook shall take priority over anything written here.

OBJECTIVES

After completion of this chapter, the student should be able to: Locate and Describe: Entry Door/Emergency Exit Baggage Area Avionics Area Cabin Section Fuselage Wing Section Lights


HISTORY OF THE KING AIR C-90A

The King Air C-90A was introduced in 1984 beginning with serial number LJ – 1063. The airplane has a max gross weight of 9650 pounds and holds 384 gallons of fuel. With a 5.0 pressurization differential and a 235 knot cruise speed, the C-90A was a perennial best-seller. Improvements included dual bleed air and an electric heater. The aircraft also included vertical engine instruments and an hydraulically operated landing gear. The aircraft is equipped with two Pratt & Whitney PT 6A – 21 engines rated at 550 hp. The C-90A production run ended with serial number LJ-1299. The C-90B was introduced. The C-90B includes better sound proofing and an upgraded interior.

GENERAL The King Air C-90A is a high performance, all metal, low wing aircraft. It is approved for day and night IFR/VFR flight operations as well as flight into known icing. (If properly equipped) The airplane has fully cantilevered wings and a conventional tail. The fuselage is pressurized to the skin between pressure bulkheads. The control cables, torque shafts, plumbing and wiring connections that pass through pressure walls are installed with fitted seals or plug connectors to minimize leakage. The King Air 90 fuselage is of semimonocoque construction and is fabricated from, frames, bulkheads and keels reinforced by


FOR TRAINING PURPOSES ONLY www.Kingairtraining.com

longerons and stringers. It is powered by two 550 SHP Pratt & Whitney turboprop engines. The C-90A is equipped with two PT6A-21 engines. The aircraft has been approved for numerous STC’s for various other engine configurations. The engines incorporate a three-stage axial and a single stage centrifugal compressor which is driven by a single-stage reaction turbine. The engine has proven to be extremely reliable. Unscheduled engine shutdowns occur approximately once every 300,000 hours. Depending on the interior configuration, the airplane can accommodate up to 10 people, although the normal corporate configuration is 6 passengers.

NOSE SECTION

The nose section of the airplane houses the radar antenna and the avionics bay. The radome is constructed of a composite material allowing radar waves to pass through easily. The nose section also contains the hydraulic brake fluid reservoir, the vacuum system inlet and the air conditioner. (Including the compressor) The nose section is un-pressurized and is accessed via removable panels on each side of the compartment. This compartment is limited to 350 pounds, which includes the weight of any avionics equipment installed within the compartment.

COCKPIT Seats The pilot and copilot seats are adjustable both fore and aft, as well as vertically. The seat adjustment lever is located under the front inboard corner of the seat. When held in the up position, the seat can be moved forward or aft as required. Lifting the release lever under the front outboard corner of the seat allows vertical adjustments to be made. Consistently good landings can be made by adjusting the vertical position of the seat to create an eye level at the center point of the windshield. The armrests pivot and can be raised or lowered as required.

Seat Belts The shoulder harness installation incorporates an inertia reel attached to the back of the seat. The two straps are worn with one strap over each shoulder and fastened into the lap belt. Spring loading at the inertia reel keeps the harness snug, but still allows normal movement required during flight. The inertia reel is designed to lock during sudden deceleration.

Oxygen Masks The quick donning oxygen masks for the pilot and copilot are stored on the bulkhead behind the pilots. Newer aircraft are equipped with masks stowed directly above the crew. On aircraft not equipped with quick donning masks, the crew oxygen mask can be located in the seat back pocket or underneath the pilot’s seat.

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PILOT TIP Beards and mustaches should be trimmed so that they do not interfere with the

proper sealing of the oxygen mask.

LIGHTING SYSTEMS

Cockpit Lights An overhead-light control panel, easily accessible to both pilot and copilot, incorporates a functional arrangement of all lighting systems in the cockpit. Each light group has its own rheostat switch placarded BRT - OFF. The MASTER PANEL LIGHTS - ON - OFF switch controls the overhead light control panel lights, fuel control panel lights, engine instrument lights, radio panel lights, subpanel and console lights, pilot and copilot instrument lights, and gyro instrument lights. The instrument indirect lights in the glareshield and overhead map lights are individually controlled by separate rheostat switches. The push-button FREE AIR TEMP switch, located on the left sidewall panel next to the gage, turns ON and OFF the lights near the outside air temperature gage.

Cabin Lights A three-position switch on the copilot's subpanel placarded CABIN LIGHTS - START BRIGHT - DIM - OFF, controls the fluorescent cabin lights. The switch to the right of the interior light switch activates the cabin NO SMOKING/FASTEN SEAT BELT signs and accompanying chimes. This three- position switch is placarded CABIN LIGHTS - NO SMOKE & FSB - FSB - OFF. The baggage-area light is controlled by a two-position switch just inside the airstair door aft of the door frame and is connected to the hot battery bus. A threshold light is located forward of the airstair door at floor level, and an aisle light is located at floor level aft of the spar cover. A switch adjacent to the threshold light turns both these lights on and off. The switch also turns the exterior entry light on and off. When the airstair door is closed, all the lights controlled by the threshold light switch will extinguish. If the master switch is on, the individual reading lights along the top of the cabin may be turned on or off by the passengers with a push-button switch adjacent to each light.

Exterior Lights Switches for the landing lights, taxi lights, wing ice lights, navigation lights, recognition lights, rotating beacons, and wing-tip and tail strobe lights are located on the pilot's sub-panel. They are appropriately placarded as to their function. Tail floodlights, if installed, are incorporated into the horizontal stabilizers and are designed to illuminate both sides of the vertical stabilizer. A switch for these lights,




placarded LIGHTS - TAIL FLOOD - OFF, is located on the pilot's subpanel. A flush-mounted floodlight forward of the flaps in the bottom of the left wing may be installed. This entry light provides illumination of the area around the airstair door, to provide passenger convenience at night. It is controlled by the threshold light switch just inside the door on the forward door frame, and will extinguish automatically whenever the cabin door is closed.

PILOT TIP In fog or low visibility conditions, landing and taxi lights should be left off to

reduce light reflections.

CABIN CONFIGURATION

Various configurations of passenger seats and couches can be installed. The standard airplane seats two pilots and six passengers. All seats are equipped with seat belts and headrests. Some passenger seats can be moved fore and aft by lifting the horizontal release bar that extends laterally under the front of adjustable seats. The seatbacks can be adjusted to any angle from fully upright to fully reclining, by depressing the release tab located on the side of the seat at the front inboard corner. W hen the tab is depressed and the passenger leans against the seatback, the seatback will slowly recline until the tab is released, or until the fully reclining position is attained. When no weight is placed against the seatback and the tab is depressed, the seatback will rise until the tab is released, or until the fully upright position is reached. The seatbacks of all occupied seats must be upright for takeoff and landing. An optional lateral-tracking passenger seat may be installed. These seats have a flat, rectangular release lever located underneath the front inboard corner of the seat. When this lever is lifted, the seats can be adjusted fore and aft, as well as laterally. When occupied these seats must be positioned against the cabin wall for takeoff and landing. The armrests can be raised and lowered by lifting the release tab located under the front end of the armrest.

Hand held fire extinguishers are mounted in the cockpit beneath the copilot seat and in the passenger cabin beneath the last seat on the left side of the airplane.

Toilet The aircraft is equipped with a chemical or electrically operated toilet that is normally installed in the aft baggage compartment. The forward facing unit is equipped with a hinged cushion cover turning the toilet into an additional passenger seat. The seat belt and shoulder harness for the toilet incorporates a single adjustable strap attached to the aft bulkhead.




Relief tubes are located on the left cabin side wall forward of the toilet and in the cockpit under the pilot’s seat .

Aft Baggage Compartment The 53.4 cubic foot aft cabin baggage compartment can be separated from the cabin by a partition or a folding curtain. It includes provisions for hanging bags as well and providing for up to 350 pounds of baggage storage. All baggage and cargo must be properly secured with the webbing provided. Any item stored in the baggage compartment is accessible in flight.

Storage and Dispensing Cabinetry A large pyramid cabinet is located just behind the left cockpit

partition. It provides storage for coffee, water, liquor decanters, trash, cold beverages and ice. Additional storage space is also available in the pull-out drawers installed next to the side facing jump seat.

PILOT TIP Maximum content weight in each drawer is 30 pounds

Airstair Door The airstair entrance is attached to the airframe by a hinge at the bottom of the door. The door swings outward and downward when opened. A hydraulic damper allows the door to open slowly. As a result, it isn't necessary for a crew member to supervise when a passenger opens the door. A stairway forms an integral part of the door and provides for easy passenger access to the cabin. The internal door steps fold in when the door is closed and fold out automatically when the door is opened. While the door is open, it is supported by a plastic-encased cable, which also serves as a passenger handrail. Dual stair assist cables are available as an option. The forward assist cable is easily detachable to provide more room for loading large baggage or cargo into the airplane. Boarding lights built into the steps provide for passenger boarding at night. The door lights are powered by the hot battery bus so they can be controlled at a switch near the door without turning on the battery switch. Closing and latching the door will turn off the stair lights regardless of switch position. The door closes against an inflatable rubber seal which is installed around the opening in the door frame. Engine bleed air supplies pressure to inflate the door seal and provide a positive seal around the door. The door latching system incorporates 4 bayonet pins and 2 "J" hooks to insure structural integrity. Proper latching of the door can be verified by both observing an annunciator light in the




cockpit and by visually confirming position marks on the pins. A pressure lockout device prevents inadvertent unlocking of the door inflight.

CAUTION Only one person at a time should be on the door stairway.

Operation The door is operated by rotating the handle in the center of the door. The inside and outside handles are mechanically interconnected. To open the door from inside the airplane, push the safety release button and rotate the handle counter clockwise. The handle is turned clockwise to open the door from outside the airplane. The release button acts as a safety device to help prevent accidental opening of the door by requiring a deliberate two handed operation to open. As an additional safety measure, a differential-pressure-sensitive diaphragm is incorporated into the release-button mechanism. The outboard side of the diaphragm is open to atmospheric air pressure and the inboard side to cabin air pressure. As the cabin to atmospheric air pressure differential




increases, it becomes more difficult to depress the release button. The door is held securely to the airframe by two latch bolts at each side of the door and two latch hooks at the top of the door. These lock into the aircraft door frame to secure the airstair door when closed. The cabin DOOR UNLOCKED light in the annunciator panel remains illuminated until the cabin door is closed securely. When the door is closed and latched, the lower forward latch bolt compresses the switch mounted behind the latch plate in the doorway. When the handle is rotated to the locked position, a contact switch is actuated, removing current to the cabin DOOR UNLOCKED light.

CAUTION If the DOOR UNLOCKED annunciator illuminates in flight, do not attempt to check the security of the door! If you have any reason to suspect that the door may not be securely locked, depressurize the cabin at a safe altitude and instruct all passengers to remain seated with their seatbelts fastened. Only after the airplane has made a full-stop landing and the cabin has been depressurized member should you check the security of the cabin door.

To close the door from outside the airplane: 1. Lift up the free end of the airstair door and push it up against the door frame

as far as possible. 2. Grasp the door handle with one hand and rotate it clockwise as far as it will

go. The door will move into the closed position. 3. Rotate the handle counterclockwise as far as it will go. 4. The release button will pop out and the door handle should be pointing aft.

To close the door from inside the airplane: 1. Grasp the handrail cable and pull the airstair door up against the door frame. 2. Next, grasp the handle with one hand and rotate it counterclockwise as far

as it will go while pulling inward on the door. The door will move into the closed position.

3. Then turn the handle clockwise as far as it will go. The release button should pop out, and the handle should be pointing down.

4. Check the security of the door by attempting to rotate the handle counterclockwise without depressing the release button. The handle should not move.

5. Lift the folded stairs to reveal a placard adjacent to the round observation window. The placard presents a diagram showing how the arm and shaft should be positioned. A red pushbutton switch near the window turns on a light inside the door to illuminate the area.

6. Proceed to check the visual inspection ports, one of which is located near each corner of the door. A green stripe painted on the latch bolt should be aligned with the black pointer .

CAUTION If any condition specified in this door-locking procedure is not met, do not take off.




PILOT TIP

Only a crew member should close the door.

CABIN WINDOWS

Cabin Exterior Windows Each cabin window is made of a sheet of clear, stretched, acrylic plastic and is seated in the window frame. The windows are part of the pressurization vessel and are capable of withstanding maximum cabin pressure differential. The plastic windows should be kept clean and waxed at all times. Only approved Plexiglas cleaners such as Mirror Glaze, Permatex Plastic Cleaner or Parko Anti-Static Plastic Polish should be utilized. To prevent scratches and crazing, wash the windows carefully with plenty of mild detergent and water. Use the palm of the hand to feel and dislodge dirt and mud. A soft cloth, chamois or sponge may be used, but only to carry water to the window surface. Rinse the window thoroughly, and then dry it with a clean, moist chamois. Rubbing the surface of the plastic window with a dry cloth will serve only to build up an electrostatic charge that attracts dust. Remove oil and grease with a cloth moistened with kerosene. Never use gasoline, benzene, alcohol, acetone, carbon tetrachloride, fire extinguisher or anti-ice fluid, lacquer thinner or glass cleaner. These liquids will soften the plastic and may cause crazing. After removing all dirt and grease from the window, it should be waxed with a good grade of commercial wax. The wax will fill in minor scratches and help prevent additional scratches. Apply a thin, even coat of wax and bring it to a high polish by rubbing lightly with a clean, dry, soft flannel cloth. Never use a power buffer; the heat generated by the buffing pad may soften the plastic.

Polarized Interior Windows Two window panes composed of a film of polarizing material laminated between two sheets of acrylic plastic are installed on the inboard side of the window. The inner most pane rotates freely in the window frame and has a protruding thumb knob near the edge. Rotation of this pane changes the relative alignment between the polarizing films, thus providing any degree of light transmission from full intensity to almost none. Do not leave the windows in the polarized position while parked on the ramp. Intense sunlight will cause deterioration of the polarizing material.

Note: Some King Air models have shade type window blinds.

WARNING! Do not look directly at the sun, even though polarized windows

because eye damage could result.




EMERGENCY EXIT

The emergency exit door is located at the third cabin window on the right cabin side wall. Inside the airplane, the door is released with two hooks, a trigger button, and a latch-release pull-up handle. A placard on the emergency exit hatch release cover details how to operate the emergency exit. If the cabin is pressurized, a pressure lock out prevents the door from being opened. Pulling the hooks will override the pressure lock and allow the trigger button to be pushed. This releases the handle. When the handle is pulled up and the securing latches are released, a hinge at the bottom allows the hatch to swing outward and downward for emergency exit.

INTERIOR DIVIDERS

Interior dividers are provided by curtains or panels.

AFT FUSELAGE

The fuselage is designed and tested to meet fail-safe structural requirements. There is no scheduled retirement or replacement requirement for the fuselage. The aft fuselage area contains the oxygen bottle and filler port. The oxygen bottle is located in an unpressurized aft compartment. Access to the compartment is through a door Located on the bottom of the right side of the Fuselage. This large lockable door on the lower surface of the fuselage immediately aft of the pressure bulkhead provides access for mechanics to reach avionics, flight controls, and other systems. All conditioned air passing out of the cabin through the outflow valves is-ducted overboard rather than being expelled into the aft fuselage. This eliminates the potential for a large amount of moisture being condensed out into the fuselage area during flight.

EMPENNAGE

The empennage includes the rudder, horizontal stabilizer, vertical stabilizer, elevators, and the trim tabs. The airplane features a conventional empennage configuration. All empennage control surfaces are mechanically operated via control cables and bellcranks. The flight control cable assemblies are pre-stretched prior to installation in the airframe. This extra manufacturing process reduces the likelihood that cables will slacken or lose tension in service. Both manual and




electric trim are used for elevator trim. The elevators incorporate dual trim tab surfaces and actuators. Dual trim tabs provide symmetrical trim loading and system redundancy. The tabs are attached to the elevator with piano type hinges to improve strength and service life. The pneumatic de-ice boots are attached to the leading edges of the horizontal and vertical stabilizers.

WINGS

The airplane utilizes a NACA 23000 series wing shape. This airfoil series exhibits a balance of good high speed performance and excellent low speed handling qualities. The NACA 23000 shape is much more tolerant of ice accumulation than a laminar flow wing. The aircraft has a wingspan of 50'3" and incorporates a 7º wing dihedral. The total wing area is 294 sq. feet. The Beech King Air 90 series wing assembly consists of the center section and two outboard wing panels. The center section is attached to and becomes an integral part of the fuselage providing structural support for the engine nacelles and the outboard wing assemblies. On airplane serial number, LJ-1088 and after, the outboard wing assemblies are attached to the center section with six tension bolts located at the upper forward, aft upper and aft lower position and two shear bolts located at the lower forward wing attach point at the spar attach points on each wing. Shear between the outboard wings and the center section is transferred through soft aluminum washers between, and imbedded in, serrations on the upper spar fittings. The center section and outboard wing assemblies are of semimonocoque box construction. Both center section spars are I-beam sections built up from aluminum extruded tee caps, webbing and stiffeners. Similar construction is used in the outboard wing spars, except that a combination of aluminum extrusion and formed U-channel members comprise the main spar caps while those of the rear spar are composed of formed aluminum angles and cap strips. The leading edge assembly and the main outboard wing assembly are joined together at the main spar by continuous hinges. A subspar is installed at the forward end of the leading edge. The space forward of this subspar is utilized to route wiring and plumbing. Between the subspar and the main spar, bladder fuel tanks are installed the full span of the outboard leading edge.

Wing Center Section The center section main and rear spars are parallel and are continuous from one outboard wing attach joint the other outboard wing attach joint. A subspar is installed forward of the main spar between the fuselage and each nacelle to which a removable leading edge is attached. The area with the removable leading edge and forward of the subspar is used to route engine controls, plumbing and wiring, etc. A subspar located forward of the rear spar provides a tunnel for control cables and shafts and serves as a fuel wall for the bladder tanks from the root rib to the nacelle. Landing gear hinge point structural supports in the nacelles are made of machined alloy plate. Formed sheet metal formers and stringers establish the nacelle fairing and a cavity for a bladder fuel tank above and forward of the wheel well.




ATTACH FITTINGS

The major fittings in each wing and the center section are the supporting structures adjacent to the attachment points for the flap actuator, flap tracks and flap, the aileron hinge brackets and hinges, the main landing gear, drag legs and landing gear doors. Minor fittings include brackets to support cable pulleys, bellcranks and similar components. The main gear is bolted to heavy aluminum alloy fittings attached to the main rib assembly at the aft end of the wheel well. The main gear drag leg is bolted to an aluminum alloy forging attached to the main spar of the center section. Wing tips are fabricated from metal and include the nav light, strobe light, and recognition light. Compass sensors (flux valves) are located in the wing tips, away from electrical field interference. Two compass systems (#1-L.H. tip, #2R.H. tip) provide for redundancy in the cockpit.

PILOT TIP

Many pilots think the King Air 90 "flies like a big Bonanza” since they share a common airfoil.

POWER PLANT

T h e C9 0A aircraft is powered by two Pratt and Whitney PT6A series engines. The PT6 is a lightweight, free-turbine engine. It utilizes a three- stage axial compressor and a single stage centrifugal compressor. These compressors are driven by a single-stage reaction turbine. A reaction turbine, called the power turbine, drives the propeller shaft through a reduction gear box. The power turbine and the reaction turbine rotate independently of each other and there is no mechanical connection between the two. The engine is covered in detail in Chapter 5 of this workbook.

ELECTRICAL SYSTEM

The aircraft uses a 28 volt multiple bus electrical distribution system. D.C. power is provided by two 30 volt, 250 amp starter-generators. Either a NiCad or lead acid 24 volt battery supplies starting and backup electrical power. Alternating Current is supplied by two invertors which provide power at 26VAC (400Hz). More information on the electrical system is supplied in Chapter 2 of this workbook.

PROPELLER SYSTEM

Each engine is equipped with either a Hartzell or McCauley 3 or 4 blade propeller. They are full feathering, constant speed, reversing, variable pitch propellers mounted on the output shaft of the engine reduction gearbox. They are equipped with an auto-feathering system. More information on the propeller system is supplied in Chapter 6 of this workbook.




FUEL SYSTEM

The fuel system is a 384 usable gallon system with each wing divided into a main fuel tank and a nacelle tank. Fuel for each engine is supplied from a nacelle tank and four interconnected wing tanks for a total of 192 gallons of usable fuel for each side with all tanks full. The outboard wing tanks supply the center section wing tank by gravity flow. The nacelle tank draws its fuel supply from the center section tank. Since the center section tank is lower than the other wing tanks and the nacelle tank, the fuel is transferred to the nacelle tank by the fuel transfer pump in the low spot of the center section tank. Each system has two filler openings, one in the nacelle tank and one in the leading edge tank. To assure that the system is properly filled, service the nacelle tank first, then the wing tanks. A crossfeed valve in the left fuel system makes it possible to connect the two systems. The fuel system is covered in detail in Chapter 4 of this workbook.

ANTI-ICE/DE-ICE SYSTEMS

The King Air is fully equipped for flight into known icing. De-icing equipment includes wing and tail deice boots and the anti icing equipment includes pitot heat, stall vane/ fuel vent heat., windshield heat, prop heat and engine inlet heat. More information on the anti ice/de-ice system is supplied in Chapter 10 of this workbook.

ENVIRONMENTAL SYSTEM

The environmental system consists of the bleed air pressurization system, heating and cooling systems and their associated controls. The environmental system is covered in detail in Chapter 7 of this workbook.

LIMITATIONS

Airspeed Limitations The limitations included in this section have been approved by the Federal Aviation Administration and they must be observed in the operation of the BEECHCRAFT King Air C90A.

SPEED

KCAS

KIAS

REMARKS

Maximum Operating Speed

VM0

226

226

Do not exceed this

airspeed in any operation.

Maneuvering

VA

169

169

Do not make full or abrupt control movements

above this speed.




Maximum Flap Extension/Extended

VFE 35% (APPROACH)

100% (FULL DOWN)

184 140

184 148

Do not extend flaps or operate with flaps ex-

tended above this speed.

Maximum Landing Gear

Operating Extension Retraction

182 164

182 163

Do not extend or retract the landing gear above

this speed.

Maximum Landing Gear Extended

VLE

182

182

Do not exceed this speed with the landing

gear extended.

Air Minimum Control VMCA

92

90

This is the lowest speed at which the airplane

is directionally controllable after sudden

loss of engine when the remaining engine is at take-off power.

AIRSPEED INDICATOR MARKINGS**

AIRSPEED INDICATOR MARKINGS**

MARKING

KCAS VALUE OR RANGE

KIAS VALUE OR RANGE

SIGNIFICANCE

Red Line

226

226

Maximum Operating Limit Speed

White Arc 74 to 140 76 to 148 Full Flap Operating Range

White Triangle

182

184

Maximum Speed For Approach Flaps

"The Airspeed Indicator is marked in CAS Values.

WEIGHT LIMITS Maximum Ramp Weight: 9710 pounds Maximum Take-off Weight: 9650 pounds Maximum Landing Weight: 9168 pounds Maximum Zero Fuel Weight: No Structural Limit Maximum Weight in Rear Baggage Compartment: 350 pounds Maximum Weight in Nose Baggage Compartment: 350 pounds

CENTER OF GRAVITY LIMITS AFT LIMIT: 160.0 inches aft of datum at all weights FORWARD LIMITS: 153.2 inches aft of datum at 9650 lbs. 151.4 inches aft of datum at 9168 lbs. 144.7 inches aft of datum at 7400 lbs or less. DATUM is 83.5 inches forward of the center of front jack point.




MAC leading edge is 135.9 inches aft of datum. MAC length is 75.9 inches.

MANEUVER LIMITS This is a normal category airplane. Acrobatic maneuvers, including spins, are prohibited.

FLIGHT LOAD FACTOR LIMITS (9650 POUNDS) Flaps Up: 3.70 positive G's 1.68 negative G's Flaps Down: 2.00 positive G's

MINIMUM FLIGHT CREW: One Pilot MAXIMUM OCCUPANCY: Ten People

EMERGENCY PROCEDURES BOLD TYPE INDICATES MEMORY ITEMS!

Illumination of Cabin Door Warning Annunciator

WARNING! Do not attempt to check the security of the cabin door. Remain as far from the door as possible with seat belts securely fastened until the airplane has landed.

1. If the CABIN DOOR warning annunciator illuminates, depressurize cabin (consider altitude first) by activating cabin pressurization dump switch on pedestal.

2. Do not attempt to check cabin door for security until cabin is depressurized and the airplane is on the ground. Check security of cabin door (on the ground) by lifting cabin door step and checking position of arm and plunger. If unlocked position of arm is indicated, turn door handle toward locked position until arm and plunger are in position.

Emergency Exit The third right cabin window is the EMERGENCY EXIT hatch.

CAUTION Do not open Emergency Exit Hatch when cabin is pressurized.

1. Emergency Release Hatch Cover - OPEN




2. Release Button - PUSH (if release button will not push PULL hooks to overcome residual friction and then PUSH the release button) PULL handle and PUSH out hatch.

Cracked Windshield If it is positively determined that the crack is on the outer panel, no action is required.

CAUTION

Windshield wipers may be damaged if used on cracked outer panel. Heating elements may be inoperative in area of crack.

If it is determined that the crack is on the inner panel, descend or reset the pressurization controller to achieve 3 psi or less differential pressure within ten minutes. Visibility through the windshield may be significantly impaired.

Spins If a Spin is entered inadvertently: 1. Control Column - FULL FORWARD 2. Full Rudder - OPPOSITE DIRECTION OF SPIN 4. Power Levers – IDLE 5. Controls - NEUTRALIZE WHEN ROTATION STOPS 6. Execute a smooth pull out.

NOTE

Federal Aviation Administration Regulations do not require spin demonstration of airplanes of this weight; therefore no spin tests have been conducted. The recovery technique is based on the best available information.

Simulating One-Engine-Inoperative (Zero Thrust)

When establishing zero thrust operation, use the power setting listed below. By using this power setting to establish zero thrust, one avoids the inherent delays of restarting a shut down engine and preserves almost instant power to counter any attendant hazard. 1. Propeller - 1800 RPM 2. Power Lever - SET 100 ft-lbs torque

NOTE

This setting will approximate Zero Thrust at low altitudes using recommended One-Engine-Inoperative Climb speeds.

GENERAL CHAPTER QUESTIONS

1. To open the emergency exit: A. Turn the release handle clockwise and pull the door down and in. B. Release Button – PUSH. PULL handle and PUSH out hatch. C. Turn the release handle counterclockwise and push the door out. D. Pull the door release handle downward and inward.




2. T or F: The nose section is pressurized.

3. The airplane can accommodate up to people.

4. Hand held fire extinguishers are located and .

5. Proper latching of the airstair door can be verified by: A. Observing the annunciator light in the cockpit B. Confirmation of green position marks on the pins in the inspection ports. C. Observe the arm and shaft position in the observation window. D. All of the above

6. T or F: On the ground, the polarized window shades should be left in the polarized position. . 7. The oxygen bottle is located: A. In the nose section B. In the aft fuselage area C. In the baggage compartment D. The airplane uses oxygen generators.

8. The maximum take off weight is . 9. List: A. Va B. Vne C. Vlo D. Vle E. Vmc

10. The maximum landing weight is_ .

11. The maximum weight in the aft baggage compartment is:

.

12. What does the red line on the airspeed indicator represent?

13. What are the emergency procedures for an illuminated Door Light annunciator warning?

14. Maximum content weight in each cabinet drawer is pounds. 15. What does the white triangle on the airspeed indicator represent? ____________________________________________________________________________________________________________________________________




Chapter 2

Electrical System OBJECTIVES

After completion of this chapter, the student should be able to:

1. Locate the control switches for:

A. Battery B. Generators C. Inverters

2. Locate the following indicators:

A. DC load/volt meters B. AC frequency/volt meters

3. On the annunciator panel state the color, probable cause f or illumination and corrective action for the following:

A. Generator B. Inverter (if installed) C. Battery charge D. Ignition

4. Using the aircraft electrical schematic locate:

A. Battery B. Hot-wired bus C. Generators D. Current limiters E. Generator busses F. Triple fed busses G. Ground power plug H. Inverters I. H.E.D.

5. Trace the DC power distribution from:

A. Battery only B. Single generator only C. Two generators D. External power unit

6. State the procedures for conducting a:

A. Bus sense check B. Normal engine start

7. State procedure for detecting:

A. A failed current limiter, failed gen tie. B. Failed gen tie combined with loss of DC generator.




8 . List acceptable voltage, amperage and polarity for external power unit.

9. Trace AC power distribution.

ELECTRICAL POWER - DESCRIPTION AND OPERATION

The Beech King Air C90 electrical system is a 28-volt DC split bus system with a negative ground. During normal operation, primary electrical power is supplied by two 30-volt, 250-ampere DC starter-generators. The secondary source of power is a 24-volt nickel-cadmium battery or a 24-volt lead-acid battery. Volt/load meters on the overhead panel indicate the load on each generator. The electrical system is designed to provide maximum protection against loss of electrical power due to ground fault. High current sensors, bus-tie relays and current limiters are provided to isolate a ground fault from a power source. The arrangement of the electrical system buses are designed to afford multiple power sources for all circuits. The King Air C-90A utilizes a three bus system. The three main buses are the left generator bus, right generator bus and the triple fed bus. All circuit breakers which receive power from the triple fed bus are identified by a white ring around the circuit breaker in the edgelighted panel. The hot battery bus is connected directly to the battery to provide power for the operation of certain systems essential to flight without generator operation such as fire extinguish, firewall shutoff, etc. The battery bus, located in the lower forward cabin under the copilot’s floor, is triple fed from the battery and from each generator bus through 250-amp limiters diodes which provide fault isolation between the power sources. Each generator bus is located aft of the firewall in the inboard side of its respective nacelle. The center bus is located under the crew compartment floor through a 250-ampere limiter and generator bus-tie relay. The generator buses, battery bus and battery are all tied together by the center bus, which in turn supplies power to the landing gear and environmental system In normal operation, the buses are automatically tied into a single loop system in which all sources collectively supply power through individual protective devices. When the battery switch is closed, the battery relay and the battery bus-tie relays will close. Battery power is routed through the battery bus-tie relay to the center bus and both starter relays. Battery power is then available to permit starting either engine. After either engine has been started and the generator system has been activated, the generator control panel (voltage regulator) will bring the generator up to voltage, then close the generator line contactor and the generator bus-tie relay. The generator output will then be routed through the center bus and the battery bus tie relay to permit battery charging and to supply power to all airplane systems. As each generator bus is energized, power is routed to the opposite cross-start system, ready for use in starting the opposite engine. The current supplied by the operating generator during a generator-assisted start thus bypasses the 300-ampere current limiter of the operating generator to prevent opening the limiter.




BATTERY SYSTEM A good battery should be able to provide sufficient stored energy for reserve or emergency power requirements in the event of a dual generator failure. As the sole source of electrical power, a fully charged battery should provide adequate power for approximately 30 minutes. The battery’s voltage can be checked by using the volt/load meter located on the pilot’s left subpanel. Adequate starting performance is not always indicative of a good battery. Normally, a periodic capacity check of the battery is required at 18 month intervals. The airplane is equipped with a 24-volt, 36-ampere-hour nickel- cadmium battery or a 24-volt, 42-ampere-hour capacity sealed lead-acid battery. Many King Air operators have elected to remove the NiCad battery and replace it with the 24 volt, 42 ampere-hour lead acid battery. Since lead acid batteries have a straight line voltage drop as the battery discharges, the aircraft manufacturer was concerned with high ITT temperatures during engine start. This concern has proven to be unfounded and the lower costs and ease of operation of lead acid batteries have outweighed any advantages of the NiCad batteries. Normally, converting a King Air from a NiCad battery to a lead - acid battery also involves removal or disconnection of the BATTERY CHARGE annunciator light.

If the airplane is equipped with the NiCad battery, a battery charge light is installed on the annunciator panel to warn the pilot of an abnormally high battery charge rate. This condition can lead to a thermal runaway of the nickel-cadmium battery. If this occurs, the pilot should follow the checklist procedure which will isolate the battery from the charging system before further battery damage occurs. The most common cause of the thermal runaway is damage to the gas barrier between the plates resulting from overcharging the battery at a high rate and high temperatures. During normal operation, the idle current of the battery is less than one amp. It increases significantly above the normal level when the battery is charged at an elevated temperature or from a high charge voltage. For this reason, the battery case incorporates a thermostatically controlled air vent to provide cooling air flow around the battery. The vent is located on the underside of the battery box. The battery monitor system provides an indication of the high charge current resulting from high battery temperature, high charging voltage or gas barrier damage. The system will illuminate the BATTERY CHG annunciator during battery recharge to provide a self-test of the system. Following an engine start, the BATTERY CHG annunciator illuminates and remains on for approximately five minutes until the battery approaches full charge. If the annunciator light remains on longer than five minutes, the battery was in a low state of charge or has gas barrier damage. After the BATTERY CHG annunciator light extinguishes, it should remain off.




PILOT TIP The battery may be damaged if exposed to voltages higher than 30V for

extended periods of time. DC GENERATION - DESCRIPTION AND OPERATION

Direct current for the electrical system is supplied by a 24-volt, 40- or 45-ampere- hour battery and two 30-volt, 250-ampere starter-generators connected in parallel. These three power sources are controlled by the generator and battery switches which are located under the MASTER SWITCH gang bar on the pilot's outboard subpanel. The three switches are located under the MASTER SWITCH gang bar for simultaneous cut-off.

The generator switches have a third (RESET) position for putting the generator back on the line after each engine start. The generator switch is spring-loaded to return from the RESET position to the ON position for generator operation. In order to turn the generator ON, the generator switch must be held upward in the reset position for one second. It is then released to the ON position. Whenever the generator control switch is in the OFF position, battery voltage is routed from the generator control circuit breaker through the generator control switch and the normally closed contacts of the field disconnect relay to the coil of the field grounding relay. This energizes the field grounding relay which grounds the field of the respective starter-generator to the airframe structure. Regulator power is interrupted and, consequently, generator operation is disabled whenever the generator control switch is OFF or when the respective engine is being started. STARTER-GENERATORS

The starter-generators are dual purpose, 30-volt, 250-ampere DC units which produce torque for engine starts or generate electrical current to meet the airplane electrical loads. The generator buses are interconnected by two 325- ampere current limiters. During an engine start, the starter generator acts as a starter and drives the engine compressor section through the accessory gearing. As the compressor turns, the starter generator can draw up to 1,100 amperes initially before dropping off to 300 amperes as the engine accelerates to approximately 20% N1. Once on line, generator voltage and load can be monitored by checking the volt/load meter on the overhead panel.




GENERATOR CONTROL UNIT

During normal operation, each generator control unit (GCU) monitors starter-generator output voltage and controls the field excitation to maintain a constant load under varying operating conditions such as speed, load and temperature. Each starter- generator has its own GCU to provide voltage regulation, generator paralleling, reverse current sensing, and overvoltage and overexcitation protection. Before the GCU can regulate starter-generator output, it must use residual voltage to build starter-generator output to a level that the regulation circuit can control. When residual voltage is applied, the starter- generator field is excited and output is increased to a level sufficient for the regulator circuit to control. Starter-generator output is adjusted by the regulator circuit to maintain 28.25 ±0.25 vdc. If no overvoltage is present and the starter- generator output is at least 0.6 vdc greater than bus voltage, the reverse current relay is energized and starter-generator output is connected to the generator bus. The applicable yellow DC GEN caution annunciator is illuminated anytime the reverse current relay is open. When the reverse current relay is closed, the annunciator will extinguish and the volt/loadmeters should indicate starter- generator output.

PILOT TIP During an engine start, ensure that the generator control switch that controls the starter-generator for the engine being started is in the OFF position. This prevents the generation of field current during engine start. The presence of field current during engine start will reduce the torque available from the starter and may lead to a hotter start.

STARTER-GENERATOR PARALLELING

The generator system is designed so that the starter-generators loads are within 10% of each other when the starter-generators are operating above 25% of the rated output. The starter-generators must both be operating at equal speeds of 57% N1 or greater for dependable paralleling. The starter-generators should share the system load with 25 amperes (a difference of 0.1 on the loadmeters) with both engines at equal speeds of 57% N1 or greater. The starter-generators will not parallel below 0.25 electrical load per starter-generator, at unequal engine speeds or at speeds below 57% N1. Adjustments in regulator voltage are automatically performed by the GCU to ensure proper paralleling. Normally, the field power of the starter-generator carrying the greater load is reduced, while the field power of the unit carrying the smaller load is increased, until both units are carrying approximately the same load. Both generators should share the electrical load equally to prevent different wear rates between generator systems.

PILOT TIP

Due to the tolerance of the loadmeters, a difference of 10 percent of the rated output of one generator may be observed and is acceptable.




OVERVOLTAGE PROTECTION

The generator control units monitor starter-generator output voltage for excessive voltage that could potentially damage the airplane electrical system . The overvoltage relay is set to trip at 32 to 34 volts. If an overvoltage condition occurs, the overvoltage relay will trip and remove the affected starter-generator from the bus. This will leave the remaining starter generator carrying the entire aircraft’s electrical load. The resultant load read on the volt load meter will depend upon starter-generator speed, electrical load and the nature of the fault. Normally, one generator is capable of handling the entire aircraft’s electrical load. This overvoltage protection circuit requires a manual reset of the starter generator to bring the starter-generator back on-line.

REVERSE CURRENT PROTECTION

If the generator field becomes under excited for any reason, or the starter-generator slows down to the point where it can no longer maintain a positive load, (such as during an engine shutdown) the starter-generator will begin to draw current from the airplane bus. This is defined as reverse current. The reverse current protection function senses starter-generator reverse current passing through the windings of the starter-generator and determines if the starter-generator has become a load rather than a power source. If reverse current is present, the generator control unit will open the line contactor relay and remove the starter -generator from the bus.

OVER EXCITATION PROTECTION

Over excitation protection is provided by the generator control unit. The generator control unit over excitation protection circuit will activate in the event that starter- generator voltages begins to increase without control, but does not go into overvoltage. If the generator field reaches its design limit; the generator will drop of line. When a failure causes excessive field excitation, the affected starter- generator will attempt to carry the airplane’s entire electrical load. During normal operation, this is sensed at the generator control unit by comparing voltages of the starter-generators. A starter-generator will be de-energized if generator bus voltage is greater than 28.5 vdc and the current output differs between starter- generators by more than 15 percent for 5 seconds. This circuit functions during parallel operation only and does not require an overvoltage fault to trip the generator off line. FIELD FLASH CIRCUIT When the generator switch is placed in reset, the generator residual voltage from terminal “B+” of the starter-generator is applied to the generator field at terminal




“A+” through a low resistance circuit, bypassing the regulator until the generator voltage builds up high enough for the voltage regulator to effectively control the generator. Any time the generator control panel has been tripped for overvoltage or the generator has a low residual voltage, reset must be used in order to bring the generator on the line.

COMPONENT LOCATION

The generator control units, current limiters, paralleling rheostats, overvoltage relays, reverse current relays, volt/loadmeter shunts, and generator bus feeder limiters, are all located beneath the floor panels in the center aisle forward of the main spar. BUS TIE SYSTEM A system of current sensors, bus tie relays and the bus tie printed circuit board is utilized to provide protection in the event of a ground fault condition on one of the buses. Three sensors, one on the center/battery bus and one in each power panel, monitor current flow between the buses. Any time excess current is detected (approximately 275 amps) flowing through the current sensor toward one of the buses, the sensor will open the coil of bus tie relay for that bus and the faulted bus will be isolated from the rest of the system. The generator bus tie printed circuit board located in the battery box initiates the closed mode activity of the bus tie system by supplying energizing current to the coils of the generator bus tie relays, located in the battery box. The generator PCB will also be energized when an external power supply is connected and the external power switch and battery switch are on. The generator bus tie relays can also be closed manually through the bus tie switch. The battery bus tie relay closes automatically when the battery switch is turned on. When a sensor detects high current on the bus it is controlling, it supplies a ground signal to its respective bus tie de-activate circuit of the bus tie PCB and opens the coil circuit of its respective bus tie relay. The bus tie relay will remain open until reset using the reset sense test switch. The bus tie system can be functionally checked by placing the test switch, located on the left outboard subpanel, in the test mode which provides a 28 vdc signal to the test circuit of the sensors and simulates a high current condition. The bus tie relays can be reset by placing the test switch in the reset mode. The gen tie switch, located on the left outboard subpanel, makes it possible to manually open the generator bus ties when they are in the closed mode by opening the grounding circuits are restored to their closed state when the switch is placed in the center position.




Annunciators for L GEN TIE OPEN, R GEN TIE OPEN and BATT TIE OPEN, are activated through the annunciation circuits of the bus tie PCB. Additionally, the MAN TIES CLOSE annunciator is activated when the manual bus tie switch is placed in the CLOSED position.

AC GENERATION

Power for the avionics equipment and the AC powered engine instruments is supplied by either or two inverters installed on the wing center section outboard of each nacelle. Inverter operation is controlled by an inverter select switch on the left subpanel. Selection of either inverter activates a relay installed near that inverter to supply DC power. An inverter to supply the 26 VAC instrument power and the 115 VAC avionics power to the using systems. Dual sources of DC power are provided for each inverter. The power select relay for each inverter is automatically selected to provide the inverter power from the adjacent operating generator, or from the center bus if the generator is not operating. When the battery power is applied to the center bus prior to engine start, inverter power is routed through a limiter and the normally closed contacts of an inverter power select relay to the power relay of each inverter. As each generator is brought up to the voltage energizing the generator bus, voltage is also routed through a circuit breaker on the RH circuit breaker panel to the coil of each inverter power select relay located in the wing center section outboard of the nacelle.

EXTERNAL POWER

The external power receptacle is located just outboard of the nacelle in the right center section. The receptacle is designed for use with an auxiliary ground power unit having a standard AN plug. An external power control printed circuit board, installed in the card rack located in the left battery box, protects the airplane electrical system from an auxiliary ground power unit with reverse polarity and/or overvoltage. The external power control printed circuit board utililizes voltage from the auxiliary power unit through two circuit breakers, the external power control (7.5 amps) and external power sense (5 amps). The circuit breakers are located adjacent to the external power receptacle in the RH center section. A voltage from the small polarizing pin of the external power receptacle is routed through the external power control circuit breaker to the overhead




meter panel voltmeter select switch via the common terminal of the external power switch. This permits monitoring of the external power voltage at the output of the auxiliary power unit. A yellow EXT PWR annunciator light located in the caution annunciator panel will illuminate when an auxiliary power unit is plugged into the external power receptacle. If the APU output is of correct polarity and voltage, between 25 to 32 vdc, the external power control card will close the external power relay when the EXT PWR switch located on the left outboard subpanel is switched to the “ON” position. When the external power relay is energized by the external power control card, voltage is applied through the external power relay to the center bus and tie control card. Closing the bus tie relays with the bus tie control switch will apply power the generator buses, battery and the triple fed bus activating the entire airplane electrical system, without the battery installed, for ground maintenance. The EXT PWR annunciator light will illuminate and stay illuminated if the auxiliary power unit voltage is between 25 to 32 vdc. The EXTPWR annunciator light will flash in the following condition:

1. Battery power applied, external power connected, but not turned ON.

2. The voltage of the auxiliary power unit is below 25 vdc. 3. The voltage of the auxiliary power unit is above 32 vdc.

The external power control card contains an overvoltage circuit that will lock the auxiliary power unit off the line, when the voltage is above 32 vdc, and the auxiliary power unit cannot be reconnected until:

1. The auxiliary power unit voltage drops to near 0 vdc, or 2. The auxiliary power unit is disconnected, or 3. The external power switch in the cockpit is momentarily switched

off.

Without battery power applied, the EXT PWR annunciator and the power relay will be powered by the auxiliary power unit.

PILOT TIP The output setting must not exceed 1000 amperes on external power sources with a higher current-carrying capability. Any current in excess of 1000 amperes may overtorque the drive shaft of the starter-generator or produce heat sufficient to shorten the life of the unit. Observe the following precautions when using an external power source:

a. Use only an auxiliary power source that is negatively grounded if the

polarity of the power source is unknown, determine the polarity with a voltmeter before connecting the unit to the airplane.




b. Before connecting an external power unit, turn all radio equipment, the

generator switches, the battery switch, and the external power switch OFF. Place the bus tie switch in the OPEN position and the overhead voltmeter select switch in the EXT PWR position.

c. Regulate the voltage of the external power unit before plugging it into the

external power receptacle. d. Turn the external power unit ON> e. Monitor the voltage of the external power unit on the overhead meter panel

voltmeter.

CAUTION The battery may be damaged if exposed to voltages higher than 30 volts

for extended periods of time.

f. Turn the battery switch ON. g. Turn the external power switch ON. h. Select CLOSE position of the bus tie switch. Observe that the BAT TIE OPEN and GEN TIES OPEN annunciators are NOT illuminated.

AVIONIC MASTER SWITCH

The avionics systems installed on each airplane usually consist of individual nav/com units, each having its own ON–OFF switch. Avionics packages will vary on different airplane installations. Due to the large number of individual receivers and transmitters, a Beech avionics master switch placarded AVIONICS MASTER POWER is installed on the pilot's panel.

PILOT TIP Voltage is required to energize the Avionics Master Power relays to remove

the power from the avionics equipment. Therefore, never apply external power to the airplane without first applying batter voltage.

CIRCUIT BREAKERS

Both AC and DC power are distributed to the various aircraft systems via two separate circuit breaker panels which protect most of the components in the airplane. The smaller one is located below the fuel management panel, to the left of the pilot. The large panel is located to the right of the copilot's position. Each of the circuit breakers has its amperage rating printed on it. Procedures for tripped




circuit breakers, and other related electrical system warnings, can be found in the "Emergency" section of the Pilot's Operating Handbook. If a non-essential circuit breaker on either of the two circuit breaker panel’s trips while in flight, do not reset it. Resetting a tripped breaker can cause further damage to the component or system. If an essential system circuit breaker trips, wait 30 seconds and then reset it. If it fails to reset, DO NOT attempt to reset it again. Take corrective action according to the procedures in the "Abnormal" section of your POH.

STATIC DISCHARGING - DESCRIPTION AND OPERATION

A static electrical charge may build up on the surface of the airplane while it is in flight. This electrical charge, if retained, can cause interference in radio, avionics and electrical equipment operation. Static buildup can also affect an uncomfortable discharge through personnel disembarking from the airplane after landing; therefore, static dischargers are installed on the trailing edges of the flight surfaces on all airplane serials and on the wing tips.

ELECTRICAL SYSTEM LIMITATIONS

External Power Limits External power carts will be set to 28.0 - 28.4 volts and be capable of generating a minimum of 1000 amps momentarily and 300 amps continuously.

Generator Limits The In-Flight Limits are: 100% GENERATOR LOAD and a MINIMUM N1 of 85% During ground operation, observe the following limitations:

GENERATOR LOAD MINIMUM N1 0 to 50% 59%

50 to 80% 61% 80 to 85% 70%

Starter Limits Use of the starter is limited to:

40 seconds ON, 60 seconds OFF. 40 seconds ON, 60 seconds OFF. 40 seconds ON, then 30 minutes OFF.




EMERGENCY ELECTRICAL PROCEDURES

BOLD TYPE INDICATES MEMORY ITEMS!

Smoke and Fume Elimination Attempt to identify the source of smoke or fumes. Smoke associated with electrical failures is usually gray or tan in color, and irritating to the nose and eyes. Smoke produced by environmental system failures is generally white in color, and much less irritating to the nose and eyes. If smoke is prevalent in the cabin, cabin oxygen masks should not be intentionally deployed. If masks are automatically deployed due to an increase in cabin altitude, passengers should be instructed not to use them unless the cabin altitude exceeds 15,000 feet.

Electrical Smoke or Fire 1. Oxygen a. Oxygen System - PULL ON b. Crew (Diluter Demand Masks) - DON MASKS (100% position) c. Mic Selector - OXYGEN MASK d. Audio Speaker - ON 2. Cabin Temp Mode - OFF 3. Vent Blower – AUTO 4. Avionics Master - OFF 5. Nonessential Electrical Equipment - OFF If Fire or Smoke Ceases: a. Individually restore avionics and equipment previously turned off. b. Isolate defective equipment. c. Cabin Pressure - DUMP

WARNING! Dissipation of smoke is not sufficient evidence that a fire has been extinguished. If it cannot be visually confirmed that no fire exists, land at the nearest suitable airport.

If Smoke Persists or if Extinguishing of Fire is Not Confirmed:

a. Land at the nearest suitable airport.

NOTE Opening a storm window (after depressurizing) will facilitate smoke and fume removal.

INVERTER FAILURE 1. Select other inverter.




ABNORMAL ELECTRICAL PROCEDURES

GENERATOR INOPERATIVE [L DC GEN] or [R DC GEN] 1. Generator - RESET, THEN ON If generator will not reset:

1. Generator – OFF 2. Operating Generator - DO NOT EXCEED 100% LOAD

GENERATOR TIE OPEN [L GEN TIE OPEN] OR [R GEN TIE OPEN]

1. Appropriate Load Meter- MONITOR a. If Less Than 100%- BUS SENSE SITCH TO RESET b. If Greater Than 100% - TURN APPROPRIATE GENERATOR OFF

(monitor opposite loadmeter; not to exceed 100%) 2. If Gen Tie Will Not Reset – MONITOR LOADMETERS

BOTH GENERATOR TIES OPEN [L GEN TIE OPEN] and [R GEN TIE OPEN] 1. GEN TIES – MAN CLOSE 2. If Gen Ties Will Not Close- MONITOR LOAMETERS

a. Battery will not charge b. Battery will be depleted by equipment on center bus

BATTERY TIE OPEN [BAT TIE OPEN]

1. Center Bus Voltage – MONITOR If Center Bus Voltage is Normal (27.5 – 29.0 vdc)

2. BUS SENSE Switch – RESET [BAT TIE OPEN]- EXTINGUISHED

If Center Bus Voltage is Zero: 3. GEN TIES – OPEN ∙ Battery will not charge ∙ Systems powered by the center bus will not be operational ∙ Landing gear will have to be manually extended 4. LANDING GEAR RELAY Circuit Breaker (Pilot’s Subpanel) –

PULL CIRCUIT BREAKER TRIPPED

1. Nonessential Circuit – DO NOT RESET IN FLIGHT 2. Essential Circuit

a. Circuit Breaker (after allowing to cool for a minimum of 10 seconds) – PULL TO RESET

b. If Circuit Breaker Trips Again – DO NOT RESET

EXPANDED ELECTRICAL PROCEDURES MULTIBUS SYSTEM CHECK

1. Generator Tie Switch – OPEN L and R GEN TIE OPEN lights illuminate.

2. Generator Loadmeters – SPLIT (COMMENSURATE WITH LOAD) 3. Voltmeter Bus Switch – LEFT GEN THEN RIGHT GEN

(27.5 – 29.0) WITHIN 1.0 VOLT




4. Generator Tie Switch – NORM GEN TIE OPEN lights extinguish LOADS PARALLEL WITHIN 10%

5. Bus Sense Switch – TEST L and R GEN TIE and BAT TIE OPEN lights illuminate

6. Bus Sense Switch – RESET Annunciators extinguish

ELECTRICAL SYSTEM QUESTIONS

1. List the items on the hot battery bus (hot wired items).

.

2 . What is the primary source of electrical power for the BE-C90? A. The NiCad or lead-acid battery. B. Ground power. C. The two 250 amp starter-generators. D. Both a & b above.

3. Typical avionics that uses AC power include the :

.

4. The purpose of the inverter is to: A. Provide alternating current to all avionics. B. Convert AC current into DC current. C. Convert direct current into alternating current. D. Provide DC power to certain aircraft systems.

5. The King Air C90 has two _____ volt and _______AMP D.C. starter - generators that are regulated to ______ volts ± .25 volts.

6. T or F: Certain engine instrument gages use AC power.

7. What is the minimum the battery voltage for a battery start? _______ A G.P.U. start? .

8. T or F: The generators may be used for 100% of their rated load continuously.

9. List the GPU setting for starting: amps volts.




10. What is the function of the HED’s?

.

11. What are the primary functions of the generator control unit? A. . B. C. D. .

12. What does the reverse current relay do?

.

13. How many amps can the lead acid battery provide for 1 hour? A. 34 B. 42 C. 24 D. 12

14. T or F: While utilizing external power, the battery switch should be on.

15. Where is the battery located? A. In the left wing center section B. In the aft compartment C. In the right wing center section D. In the nose compartment

16. When a generator is off the line, what indication is present? A. A yellow DC GEN light is illuminated. B. The Generator switch is in the OFF position. C. A green DC GEN light is illuminated. D. A red DC GEN light is illuminated.

17. Where is the external power connector located? A. Under the left wing B. On the left aft fuselage C. Under the right wing, outboard of the engine nacelle D. On the right forward fuselage




18. When an engine is being started, in what position should the GEN switch be? A. RESET B. ON C. OFF

19. What indication is provided to alert the operator that an external power plug is connected to the airplane? A. An audible tone B. An EXT PWR light C. A master warning light D. Fluctuating generator meters

20. How many inverters are there? A. 1 B. 2 C. 3

D. 4 21. What is the rating of each inverter? A. 28-volt and 26-volt, 400 Hz B. 24-volt and 130-volt, 60 Hz C. 115-volt and 26-volt, 400Hz D. 30-volt and 115-volt, 120 Hz

22. What are the starter limits? A. 40 seconds ON, 60 seconds OFF, 40 seconds ON, 60 seconds OFF, 40 sec-

onds ON, 30 minutes OFF B. 10 seconds ON, 30 seconds OFF, 40 seconds ON, 60 seconds OFF, 60 sec-

onds ON, 90 seconds OFF C. 20 seconds ON, 60 seconds OFF, 20 seconds ON, 60 seconds OFF, 20 sec-

onds ON, 90 minutes OFF D. 15 seconds ON, 50 seconds OFF, 15 seconds ON, 60 seconds OFF, 10 sec-

onds ON, 5 minutes OFF

23. What is the purpose of static wicks?

.

24. Explain how to perform a multibus system check:

25. What does the GEN TIE switch do?

26. What is the Bus Sense switch used for?




Chapter 3

Annunciator System

OBJECTIVES

After completing this chapter, the pilot will be able to:

1. Identify the components of the annunciator system. 2. Describe the dimming procedure. 3. Describe the Master Warning and Master Caution features. 4. Explain the significance of the light colors used in the annunciator panel.

ANNUNCIATOR SYSTEM

The annunciator system consists of warning annunciator panel centrally located in the glareshield, and an annunciator panel dimming control, a press-to-test switch, and a fault warning light. A Red MASTER WARNING flasher, an amber MASTER CAUTION flasher, and a PRESS TO TEST button are also part of the system. These are located immediately to the left of the warning/caution/advisory annunciator panel. The illumination of a green annunciator light will not trigger the fault warning system

Whenever an annunciator-covered fault occurs that requires the pilot’s attention but not his immediate reaction, the appropriate amber caution annunciator illuminates, and the MASTER CAUTION flasher begins flashing. The MASTER CAUTION flasher can be extinguished by depressing the face of the MASTER CAUTION flasher to reset the circuit. Subsequently, when any additional caution annunciator illuminates, the MASTER CAUTION flasher will be activated again. An illuminated caution annunciator on the warning/caution/advisory annunciator panel will remain on until the fault condition is corrected, at which time it will extinguish. The MASTER CAUTION flasher will continue flashing until depressed. The warning/caution/advisory annunciator panel also contains the green advisory annunciators. There is no master flasher associated with these annunciators, since they are only advisory in nature, indicating functional situations which do not demand the immediate attention or reaction of the pilot. An advisory annunciator can be extinguished only by disengaging the condition/system indicated on the illuminated lens. The warning annunciators, caution annunciators, advisory annunciators, MASTER WARNING flasher, and MASTER CAUTION flasher feature both a “bright” and “dim” mode of illumination intensity. The “dim” mode will be selected




automatically whenever all of the following conditions are met: a generator is on the line; the OVERHEAD FLOOD LIGHTS are OFF; the PILOT FLIGHT LIGHTS are ON; and the ambient light level in the cockpit (as sensed by a photoelectric cell located in the overhead light control panel) is below a preset value. Unless all of these conditions are met the “bright” mode will be selected automatically.

The lamps in the annunciator system should be tested before every flight, and anytime the integrity of a lamp is in question. Depressing the PRESS TO TEST button, located to the left of the warning annunciator panel in the glare -shield, illuminates all the annunciator lights, MASTER WARNING flashers, and MASTER CAUTION flashers. Any lamp that fails to illuminate when tested should be replaced.




ANNUNCIATOR EMERGENCY PROCEDURES None.

ANNUNCIATOR ABNORMAL PROCEDURES None.




ANNUNCIATOR SYSTEM QUESTIONS 1. Name the three annunciator panels and the color of the lights associated with these

panels.

. 2. Where is the master warning flasher located?

.

3. What would make it illuminate?

.

4. What would cause the MASTER CAUTION to illuminate?

. 5. How do you dim the annunciator panel lights?

.




Chapter 4

Fuel System

OBJECTIVES

After completion of this chapter, the student will be able to:

1. Identify fuel system controls, components, functions and gauges. 2. Explain fuel annunciator lights, probable cause for il lumination and corrective action. 3. Describe fuel tanks, location and capacities 4. Identify approved fuels. 5. State sequence of filling tanks. 6. Locate all preflight fuel drains. 7. Describe fuel vent system. 8. Describe flow of fuel from tanks to engine . 9. Describe operation of fuel transfer pumps. 10. Describe operation of fuel crossfeed system. 11. Explain fuel check procedures conducted before flight. 12. List fuel system limitations, normal and emergency procedures.




FUEL SYSTEM - DESCRIPTION AND OPERATION

The airplane fuel supply system consists of individual fuel systems for each engine. The systems are controlled from a fuel control panel located on the left side of the pilot's compartment. A crossfeed line and valve allows each system to supply fuel to either or both engines as necessary. Fuel for each system is contained in one nacelle tank of 61 gallons capacity and four interconnected wing cells with a combined capacity of 134 gallons. Each fuel cell cavity is lined with a rubber bladder-type cell.

PILOT TIP Do not allow the fuel cells to dry out and crack. .

A 44 gallon cell is located in the wing center section. The outboard wing panel contains two 25 gallon tanks. A 40 gallon tank in the wing leading edge brings the total fuel capacity to 195 gallons per side. The total usable fuel capacity of both tanks is 384 gallons. The filler cap for this system of tanks is located on the leading edge near the wing tip. An anti-siphon valve is installed in each filler port which prevents loss of fuel or collapse of a fuel cell bladder in the event of improper securing or loss of the filler cap. The fuel system also incorporates electrical boost and transfer pumps and an electrically operated crossfeed valve. Three modes of operation are available, each of which is described briefly.

1. NORMAL OPERATION. Each engine receives fuel from its corresponding

fuel cells and boost pump. The cross-feed valve control switch is in the AUTO position. The cross-feed valve is closed but is armed for automatic operation.

2. AUTOMATIC CROSSFEED OPERATION. In the event of a boost pump failure, standby boost pressure is obtained by supplying fuel to both engines, through the crossfeed valve, from one boost pump. A drop in output pressure from the failed pump is sensed by a pressure switch which automatically opens the crossfeed valve when the pressure drops below 5 psi.

3. SUCTION FEED. This mode of operation may be employed after a boost pump has failed and allows the use of fuel from tanks on the side with the failed pumps. Suction feed operation is obtained by moving the crossfeed valve control switch from the AUTO position to the OFF (valve closed) position. Vacuum created by the engine driven fuel pump lifts fuel from the nacelle fuel tank. Suction feed operation is restricted to 10 hours total time between engine overhaul periods. If the engine driven fuel pump is operated on suction feed beyond the 10 hour limit, overhaul or replacement of the pump is necessary.

PILOT TIP Suction feed should only be used after cruise altitude has been attained.




Fuel level in the nacelle tank is automatically maintained at near full capacity during normal operation by a fuel transfer system whenever the fuel level in the nacelle tank drops by approximately 10 gallons. A transfer pump, located in each center section wing cell, pumps fuel from the wing tanks to the nacelle tank. The transfer pumps are controlled by float operated switches on the nacelle tank fuel quantity transmitters. A pressure switch, located in the fuel transfer line, will automatically turn off the transfer pump if a pressure of approximately 3.0 psi is not obtained within approximately 30 seconds after the pump is turned on or if the transfer pump pressure drops below 1 psi due to empty wing tanks or pump failure. A NO FUEL XFR warning light illuminates when the pump is automatically turned off. The NO FUEL XFR light is also illuminated when the transfer pump function switch is placed in the TEST position and will stay illuminated until sufficient pressure is created in the fuel transfer lines to open the pressure sensing switch. If the transfer pump fails, 28 gallons of fuel remains trapped and unusable in the wing because of wing dihedral and the location of the gravity feed line in the tank wall.

FUEL GAUGES

The fuel quantity indicating system is a capacitance type system that is compensated for specific gravity and reads in pounds on two fuel gages on the fuel control panel. Fuel quantity control monitors operate in conjunction with the fuel quantity capacitance probes in the various fuel cells to measure the quantity of fuel in the fuel system of each wing. A selector switch located between the fuel quantity indicators in the fuel panel beside the pilot may be set in either the NACELLE or TOTAL position. Each side of the airplane has an independent gauging system consisting of a fuel quantity transmitter unit in the nacelle tank, one in the center section tank, one in the inboard wing tank and two in the leading edge tank. A maximum indication error of 3% may be encountered in the system. The system is designed for the use of Jet A, Jet A1, JP-5 and JP-8 aviation kerosene, and compensates for changes in fuel density due to temperature changes. If other fuels are used, the system will not indicate correctly. The gages are marked in pounds.

PILOT TIP

Fill the nacelle tanks first. Filling the nacelle tanks first prevents fuel transfer through the gravity feed interconnect lines from the wing tanks into the nacelle tanks during fueling. If wing tanks are filled first, fuel will transfer from them into the nacelle tank leaving the wing tanks only partially filled.




FUEL DRAIN VALVES

The drain ring for the firewall fuel strainer drain is located on the firewall under the cowling on the right side of the engine. The leading-edge tank sump is on underside of outboard wing just forward of main spar. The boost pump sump is at the bottom center of nacelle forward of wheel well. The transfer pump sump drain is just outboard of the wing root and forward of flaps. The wheel well sump is located inside of the wheel well. The drains should be checked for fuel contamination during each preflight. To permit purging the fuel supply line during engine start, a purge solenoid valve is installed in the fuel return line. This valve is connected to the starter switch and is opened only when the switch is in the START position.

PILOT

TIP Check fuel at each drain point for contamination and allow a three-hour settle

period whenever possible.

FIREWALL SHUTOFF VALVES

Electrically operated, gate-type shutoff valves are mounted behind the firewall on the outboard side of each nacelle. Relief valves are incorporated in the valves to relieve thermal expansion downstream of the valve. The firewall shutoff valves receive electrical power from the triple-fed bus. When the FUEL FIREWALL VALVE switch is closed, its respective firewall shutoff valve shuts off the flow of fuel to the engine. Only fuel is cut off to the engine with this switch

FUEL VENTS

The main and auxiliary fuel systems are vented through a recessed vent coupled to a static vent on the underside of the wing adjacent to the nacelle. One vent (NACA) is recessed to prevent icing. The second vent is heated to prevent icing and serves as a backup should the NACA vent become plugged. The outer wing tanks are cross vented with one another.

FUEL PUMPS

Fuel is pumped to the engine by an electrically powered low pressure boost pump submerged in the nacelle tank. The purpose of this pump is to provide pressurized fuel to the high pressure engine driven fuel pump. The low pressure boost pump




provides lubrication and prevents cavitation of the high pressure fuel pump. It is not an emergency back up pump to the high pressure pump. The high pressure pump is engine driven and operates at approximately 800psi. The high pressure engine-driven fuel pump is mounted on the accessory case in conjunction with the fuel-control unit. This pump is protected against fuel contamination by an internal, 90-mesh strainer. This pump provides sufficient fuel pressure to insure a proper spray pattern of fuel in the combustion chamber. Failure of this pump results in an immediate engine flameout. The high pressure pump is not designed to suction feed fuel from the nacelle tank. Its function is to push fuel into the engine. If an engine driven high pressure pump is required to suction feed from the nacelle tank, severe pump damage will result. For this reason, operation with the FUEL PRESSURE annunciator on is limited to 10 hours between engine driven high pressure pump overhaul or replacement. Failure of the electric boost pump would illuminate the FUEL PRESSURE annunciator light. A pressure switch senses boost pump fuel pressure at the fuel filter. At less than 10 psi of pressure, a switch closes and actuates the red FUEL PRESSURE warning light in the annunciator panel. At this time, the system will begin to crossfeed automatically. The pilot may elect to close the crossfeed switch and continue the flight using the high pressure engine driven fuel pump or continue with the crossfeed operation.

CAUTION OPERATION WITH THE FUEL PRESSURE LIGHT ON IS LIMITED TO 10 HOURS BETWEEN OVERHAUL OR REPLACEMENT OF THE ENGINE -

DRIVEN FUEL PUMP.

The boost pumps are controlled by toggle switches on the fuel-control panel. The power source for the boost pumps is supplied from the left or right Generator Bus. The alternative source of power to the boost pumps is directly from the battery through the Hot Battery Bus. To prevent electrical interference with the avionics equipment of the aircraft, a noise filter for the standby boost pump is installed on the airplane. After shutdown, both boost pump switches must be in the off position to prevent discharge of the battery.

FUEL TRANSFER PUMPS

A submerged fuel transfer pump, located in each center wing section cell, automatically maintains the fuel level in the nacelle fuel cells at or near full capacity. (61 Gals) Fuel is transferred automatically when the TRANSFER PUMP – OVERRIDE – AUTO – OFF switches are placed in AUTO, unless the nacelle tanks are full. Magnetic switches, incorporated in the nacelle fuel cell fuel quantity transmitter, control the operation of the transfer pumps. A pressure switch connected to the transfer line automatically turns the transfer pump off when the wing cells are empty or the pump fails. To allow time for the pressure to buildup when the pump is first turned on, a time delay relay keeps the pump energized for approximately 30 seconds. If the pressure does not build up within this period, the pump is automatically turned off. When the fuel transfer switches are turned on, the transfer pumps begin operation and continue




operating until the nacelle tank is filled and the high level switches of the fuel level transmitter close. This energizes the transfer pump relay and opens the transfer pump circuit, stopping the pump. When the fuel level in the nacelle tank decreases by approximately 10 gallons, a relay is energized starting the transfer pump and the cycle is repeated. If within 30 seconds the transfer pump fails to produce sufficient pressure to open the fuel pressure switch, the contacts of the time delay relay close. This provides a NO FUEL TRANSFER signal for the annunciator panel and energizes the transfer pump relay to stop the pump. The time delay relay is latched by a diode to prevent the transfer pump from being turned on. The time delay relay may be reset by placing the transfer pump in the OFF position for a time (normally about 60 seconds) sufficient for the relay to cool and the points to open. The function switch provides a means for testing either the left or right transfer system. Placing the transfer pump switch in the ON position starts the pump, except when the nacelle tank is full. Should a no-transfer condition exist, the LH TEST or RH TEST position of the function switch bypasses the time delay relay to give an immediate NO FUEL XFR indication on the annunciator panel. If the pump is not running, due to normal cycling, selection of the TEST position biases the transistor switch to start the pump. A momentary NO FUEL XFR indication denotes normal transfer. If the time delay relay has been actuated, the TEST selection will not start the pump. The OVERRIDE positions of the transfer pump switches may be used in the event that either or both nacelle tanks’ float switches fail to function. When in the OVERRIDE position, the transfer pumps run continuously. If the nacelle tanks become full, the excess will be returned to the wing center section tanks through the fuel vent lines. If the transfer pump fails to operate during flight, gravity feed will perform the transfer. When the nacelle tank level drops to approximately 150 pounds, or approximately 22 gallons (83.3 liters), the gravity port in the nacelle tank opens and gravity flow from the wing tank starts. All wing fuel, except approximately 188 pounds, (28 gallons, 106 liters) from each wing, will transfer during gravity feed.

FUEL FILTERS

From the firewall shutoff valve, fuel is routed to the engine-driven fuel pump through the main fuel filter on the lower center of the engine firewall. This 20- micron filter incorporates an internal bypass valve to permit fuel flow in the event of a blockage. In addition to the main fuel filter, a screen strainer filter is located at each tank outlet before the fuel reaches the boost or transfer pumps. The high pressure engine driven pump incorporates an integral strainer to protect the pump. A red button on the top of the fuel filter is a contamination indicator. Fuel pressure differential of 1.0 to 1.4psi, due to contamination, will cause the red button to pop up. This is an indication that the filter needs servicing. Cleaning the filter should be accomplished as soon as practicable after the button has




popped up, whether or not the regular servicing interval has been reached. Blockage of the fuel filter will cause fuel to bypass the filter and flow to the engine. Internal passages and relief valves in the fuel filter all this.

ENGINE FUEL CONTROL LINE HEATER

A heating element is wrapped around t he engine air pressure sense line immediately before entering the engine fuel control unit. The line between the fuel control unit and primary governor is similarly heated. Each heating element is controlled by a switch in the condition lever.

FUEL HEATER

From the main filter, fuel is routed through the fuel flow transmitter and then to the fuel heater. The fuel heater utilizes heat from the engine oil to warm the fuel prior to sending it to the fuel control unit. The fuel heater is thermostatically controlled to maintain a temperature range of 70º to 90ºF. This action prevents water from freezing in the fuel lines. The fuel is then routed to the fuel-control unit that monitors the flow of fuel to the engine fuel nozzles .

CROSSFEED

Crossfeed is only to be conducted during single engine or boost pump failure operations. Each nacelle tank is connected to the opposite engine by a crossfeed line. Crossfeed operation is controlled by a three position crossfeed switch labeled OPEN, CLOSED or AUTO. In the event of a boost pump failure, standby boost pressure is obtained by supplying fuel to both engines, through the crossfeed valve, from one boost pump. A drop in output pressure from the failed pump is sensed by a pressure switch which automatically opens the crossfeed valve when the pressure drops below 5 psi. When the crossfeed valve is open, the FUEL CROSSFEED light on the annunciator panel will illuminate. The crossfeed will not transfer fuel from one tank to another; its primary function is to supply fuel from one side to the opposite engine during an engine -out condition or a boost pump failure. In the event of a boost pump failure during takeoff, the system will begin to crossfeed automatically allowing the pilot to complete the takeoff without an increase in work load at a crucial time. After the takeoff is completed, or if the boost pump fails after takeoff, the crossfeed switch may be closed and the flight continued, relying on the engine-driven high pressure pump. In some instances, the pilot may elect to continue the flight with the remaining boost pump and the crossfeed system in operation.




FUEL PURGE SYSTEM

Engine compressor discharge air (P3 air) pressurizes a purge tank. On engine shutdown, the fuel manifold pressure subsides, allowing the engine fuel manifold poppet valve to open. The purge tank pressure then forces fuel out of the engine fuel manifold lines through the fuel nozzles and into the combustion chamber. As the fuel is burned, a momentary surge in gas generator rpm (N1) should be observed. The entire operation is automatic and requires no input from the crew. On engine start-up, fuel manifold pressure closes the fuel manifold poppet valve allowing P3 air to pressurize the purge tank.

FUEL LIMITATIONS

Approved Engine Fuels COMMERCIAL GRADES: Jet A, Jet A-1, Jet B MILITARY GRADES JP-4, JP-5, JP-8

Emergency Engine Fuels COMMERCIAL AVIATION GASOLINE GRADES: 80 Red (Formerly 80/87) 91/98

10OLL Blue 100 Green (Formerly 100/130) 115/145 Purple

Limitations on the use of aviation gasoline 1. Operation is limited to 150 hours between engine overhauls. 2. Operation is limited to 8,000 feet pressure altitude (FL 80) or below with boost

pumps inoperative. 3. Crossfeed capability is required for climbs above 8,000 feet pressure altitude

(FL 80).

Approved Fuel Additives/Anti-Icing Additives Engine oil is used to heat the fuel on entering the fuel control. Since no temperature measurement is available for the fuel at this point, it must be assumed to be the same as the OAT. The graph below is used to determine the minimum oil temperature required to maintain the fuel temperature above the freezing point of water, and thus prevent ice accumulations in the fuel control unit. Enter the graph at the known or forecast OAT and determine the minimum oil temperature required for each phase of flight. If the anticipated actual oil temperature is not equal to, or above this minimum temperature, anti-icing additive conforming to MIL-1-27686 or MIL-1-85470 must be added to the fuel.




CAUTION Before refueling, check with the fuel supplier to determine whether or not anti-icing additive has already been added to the fuel. If anti -icing additive is required, it must be properly blended with the fuel to avoid deterioration of the fuel cell sealant. The additive concentration shall be a minimum of 0.10% and a maximum of 0.15% by volume. To assure proper concentration by volume of fuel on board, blend only enough additive for the unblended fuel.

Fuel Biocide Additive Fuel biocide-fungicide BIOBOR JF in concentrations of 135 ppm or 270 ppm may be used in the fuel. BIOBOR JF may be used as the only fuel additive, or it may be used with the anti-icing additive conforming to MIL-1-27686 or MIL- 1-85470 specification. Used together, the additives have no detrimental effect on the fuel system components. Refer to the Beech Super King Air 90 Series Maintenance Manual and to the latest revision of Pratt and Whitney Canada Engine Service Bulletin No. 3044 for concentrations to use and for procedures, recommendations and limitations pertaining to the use of biocidal/fungicidal additives in turbine fuels.

FUEL MANAGEMENT

USABLE FUEL (GALLONS X 6.7 = POUNDS) Total Fuel Quantity 387 gallons Total Usable Fuel Quantity 384 gallons (2573 pounds)

Fuel Management Limitations 1. Do not take off if fuel quantity indicator is in yellow arc or if fuel quantity is

less than 265 pounds in each wing system.




2. Operation on aviation gasoline is limited to 150 hours during any one engine overhaul period.

3. Operation is limited to 8000 feet when operating on aviation gasoline with boost pumps inoperative.

4. Both boost pumps must be operable prior to takeoff. 5. Operation with the FUEL PRESS Annunciator on is limited to 10 hours

between main engine driven fuel pump overhaul or replacement period, (See FUEL PRESSURE this section).

6. Crossfeeding of fuel is permitted only in the event of: a) Electric boost pump failure or engine failure.

Fuel Crossfeed Crossfeeding of fuel is permitted only when one engine is inoperative or Boost Pump failure.

Fuel Gages in the Yellow Arc Do not take off if fuel quantity gages indicate in the yellow arc or indicate less than 265 pounds of fuel in each main tank system.

EMERGENCY FUEL SYSTEM PROCEDURES

BOLD TYPE INDICATES MEMORY ITEMS!

Boost Pump Failure

NOTE With crossfeed in AUTO, a boost pump failure will be denoted only by the illumination of the FUEL CROSSFEED Annunciator. To identify the failed boost pump, momentarily place the crossfeed in the CLOSED position. The FUEL PRESS Annunciator on the side of the failed boost pump will illuminate. Then place the crossfeed switch in the OPEN position. The FUEL PRESS Annunciator will then extinguish.

1. Inoperative Fuel Boost Pump - OFF 2. Determine whether continuation of flight with crossfeed open is possible.

CAUTION

If crossfeed is discontinued, excessive power fluctuations may be experienced; open crossfeed immediately.

3. To continue flight with crossfeed closed, satisfactory operation may be

obtained by: A. Reducing power B. Descending to a lower altitude C. Waiting for fuel to cool




NOTE

Accumulated time of operation with FUEL PRESS Annunciator illuminated is limited to ten hours.

Crossfeed (Emergency One-Engine-Inoperative Operation) 1. Fuel Boost Pumps - ON 2. Transfer Pumps - ON 3. Crossfeed - OPEN, Check FUEL CROSSFEED Annunciator - ON 4. Fuel Boost Pump (non-feeding tank) - OFF (Check respective FUEL PRESS

Annunciator out)

To Discontinue Crossfeed: 1. Both Fuel Boost Pumps - ON 2. Crossfeed Switch – CLOSED 3. Fuel Boost Pump (inoperative engine) - OFF

ABNORMAL FUEL PROCEDURES

Transfer Pump Failure When the L or R NO FUEL XFR annunciator illuminates and there is fuel in the wing tanks, the nacelle fuel quantity will decrease to approximately 150 pounds indicating a failure of that transfer pump and gravity-feeding will begin.

1. Transfer Pump OFF

CAUTION If a transfer pump fails during flight, all but 28 gallons (190 pounds) will gravity

feed into the nacelle tank.

NOTE When wing fuel is depleted, the L or R NO FUEL XFR annunciator will

illuminate as a result of normal system operational logic.

Failure of Nacelle Tank Switch If the nacelle fuel quantity drops to approximately 150 pounds and there is fuel in the wing tanks, a failure of the nacelle tank switch is indicated. Proceed as follows: 1. Transfer Pump Switch – OVERRIDE

In this mode the transfer pump will run continuously until the transfer pump switch is returned to OFF position.




FUEL SYSTEM EXPANDED PROCEDURES

Fuel Panel Check 1. Circuit Breakers – IN 2. Battery Switch - ON 3. Fuel Firewall Valves – CLOSED 4. Battery Switch - OFF 5. Crossfeed - OPEN (Check FUEL CROSSFEED Annunciator on), then

CLOSED 6. Boost Pumps - ON (listen for operation) 7. Battery Switch - ON (Check left and right FUEL PRESS Annunciators on.) 8. Fuel Firewall Valves - OPEN (Check left and right FUEL PRESS

Annunciators off.) 9. Fuel Quantity – CHECK

10. Transfer Pumps - ON (listen for operation), then OFF.

If either of both pumps fails to operate, press the Transfer Test Switch and monitor the respective NO FUEL XFR Annunciator.

Boost Pump/Auto Crossfeed Test 1. Left Boost Pump OFF L FUEL PRESS light flashes and extinguishes; FUEL CROSSFEED light illuminates. 2. Left Boost Pump ON 3. Crossfeed Switch CLOSED THEN AUTO 4. Right Boost Pump OFF R FUEL PRESS light flashes and extinguishes; FUEL CROSSFEED light illuminates. 5. Right Boost Pump ON 6. Crossfeed Switch CLOSED THEN AUTO

FUEL SYSTEM QUESTIONS

1. List the items on the fuel panel that receive power from the Hot Battery Bus.

_________________________________________________________.

2. T or F: The engine will continue to operate at reduced power with boost pump pressure after the failure of the high pressure fuel pump.

3. T or F: Prist always has to be added to the fuel.

4. Maximum useable fuel capacity is: lbs.




5. Fuel is heated prior to entering the fuel control unit by: A . Bleed air from the engine’s compressor B . Engine oil, through an oil-to-fuel heat exchanger C . The friction heating caused by the boost pump D . An air-to-fuel heat exchanger prior to the fuel control unit.

6. Which of the following is a function of the electric boost pump?

A. It functions as a backup pump in the event of a primary fuel pump failure. B. It is used with aviation gas in climbs above 8,000 feet C. It is used in crossfeed operation D. B and C

7. Total fuel capacity gallons lbs.

Main Tanks gallons lbs. Nacelle Tanks gallons lbs.

8. When is crossfeed use authorized? A. For single-engine operation B. For climbs above 8,000 feet when aviation gas is used C. When fuel imbalance is 150 pounds. D. When fuel pressure decreases below 10 ± psi.

9. A hour settle period is recommended before sampling the fuel.

10. Which of the following limitations applies to operation with aviation gas?

A. A maximum altitude of 20,000 feet with both boost pumps operative and 150 hours between overhauls

B. A maximum altitude of 8,000 feet with boost pump inoperative and 150 hours between overhauls.

C. A maximum altitude of 20,000 feet with one boost pump inoperative and 150 hours between overhauls

D. A maximum of 150 hours between overhauls only

11. Is a fuel biocide additive required for this aircraft? .

12. Illumination of the fuel pressure warning light indicates:

13. T or F: Take off with the fuel quantity indicator in the yellow is approved.

14. T o r F : The “NO FUEL XFR” light will come on after the wing fuel is completely transferred to the nacelle tank.

15. You fuel the airplane with jet fuel and mix in 100 gallons of AVGAS. Each engine must be charged______________ hour(s) against its 1 5 0 hour AVGAS limitation.

16. If a transfer pump fails during flight, all but ___ gallons will gravity feed into the nacelle tank.




17.List the Cross -Feed Procedure: __________________________________________________________ __________________________________________________________ _________________________________________________ _________ _______________________________________________.

18. T or F: All fuel vents are electrically heated.




Chapter 5

Engine System OBJECTIVES

After completing this unit, the student should be able to: 1. Trace the internal airflow pattern of the engine.

2. State the basic design type of the engine.

3. State the power source for each engine gauge. 4. List pertinent engine limitations and restrictions.

5. Place in correct order the procedural steps of a normal engine start.

6. Place in correct order the procedural steps for the engine clearing procedure. 7. List the starter time limitations.

8. State the correct procedure for normal engine shutdown.

GENERAL ENGINE DESCRIPTION

The King Air C90 was introduced with Pratt & Whitney PT6A-21 engines. The -21 is rated at 550 SHP. The Pratt & Whitney PT6A engine is a light weight, reverse flow, free turbine engine driving a propeller through a two-stage reduction gearbox. Two major rotating assemblies compose the heart of the engine. One assembly consists of the compressor and the compressor turbine. The other includes one power turbine and the power turbine shaft. The two shafts are not connected together and rotate at different speeds and in opposite directions. This configuration allows the pilot to vary the propeller speed independently of the compressor speed. Starter cranking torque is low since only the compressor is




initially rotated on start. Activating the starter mounted on the accessory gearbox starts the engine. The compressor draws air into the engine through a screened annular air inlet and increases the air pressure across the 3 axial stages and one centrifugal impeller. From there it is delivered to the combustion chamber. Air enters the combustion chamber through small holes and at approximately 17% N1 fuel is introduced into the combustion chamber. Two spark igniters located in the combustion chamber ignite the mixture and the hot gases are directed to the turbine area. At this point, the ignition and starter are turned off since a continuous flame now exists in the combustion chamber. The hot expanding gases accelerate through the compressor turbine and drive the compressor. The expanding gases continue to the power turbine and provide rotational energy to drive the propeller shaft. The reduction gearbox reduces the power turbines speed (approximately 33,000 RPM) to one suitable for propeller operation (1800 to 2200 RPM). This is done through a reduction gearbox which converts the high speed, low torque of the power turbine to low speed, high torque required of the propeller. Gases leaving the power turbines are expelled out to the atmosphere by the exhaust duct. Engine shutdown is accomplished by cutting fuel going to the combustion chamber. An integral oil tank located between the inlet case and the accessory gearbox provides oil to bearings and other various systems, such as propeller and torque systems. A fuel control unit mounted on the accessory gearbox regulates fuel flow to the fuel nozzles in response to power requirements and flight conditions. The propeller governor, mounted on the reduction gearbox, controls the speed of the propeller by varying the blade angle depending on power requirements, pilot RPM selection and flight conditions. To properly understand the operation of the PT6 series engine, there are several basic terms the pilot should become familiar with:




TURBOPROP ENGINE SYMBOLS AND THEIR MEANINGS N1 (or NG) - Gas generator speed (rpm or %) N2 (or Nf) - Power turbine speed (rpm or %) Np - Propeller speed (rpm or %) FCU - Fuel control unit Tq - Torque OAT - Outside air temperature PSIG - Pounds per square inch gage PSIA - Pounds per square inch absolute SHP - Shaft horsepower ESHP - Equivalent shaft horsepower FOD - Foreign object damage Beta - Propeller non-governing mode of operation P3 - Compressor discharge pressure Px - Acceleration and speed enrichment pressure Py - Governor pressure P1 - Fuel pump delivery pressure P2 - Metered fuel pressure Po - Bypass fuel pressure Wf - Fuel flow T5 - Interturbine temperature (ITT) BOV - Bleed off valve RGB - Reduction gearbox AGB- Accessory gearbox

N1, Np, Tq, and T5 are indicated on engine gauges along with oil temperature, oil pressure and fuel flow.

The engines used on the King Air C90 have seven major sections:

1) Air inlet section

2) Compressor section

3) Combustion section

4) Compressor Turbine

5) Power Turbine

6) Exhaust

7) Reduction Gear




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.

Air Intake Section

The air inlet system is designed to provide the maximum possible total pressure at the air inlet screen over a wide band of normal flight conditions. The compressor air intake consists of circular, screen-covered aluminum housing. The screen greatly reduces the possibility of foreign objects being ingested into the engine. Because the screen area is very large, the velocity through the screen is sufficiently low to permit a high degree of screen blockage from debris or ice without significant power losses. Air is directed to the air intake through air scoops located on the bottom of the engine. The function of the air intake section is to direct airflow to the compressor section.

Compressor Section

The compressor section consists of a four-stage compressor assembly comprised of three axial stages and one centrifugal stage. The function of the compressor is to compress and supply air for combustion, engine cooling, pressurization and pneumatics, compressor bleed valve operation, and bearing

sealing and cooling. Bleed air is taken off the engine after the compressor stage and prior to the air entering the combustion can. This air is referred to as P 3 air due to the station it is extracted from. It is used for airframe pressurization and pneumatic systems.





Compressor Bleed Valves

Below approximately 80% N1, the compressor axial stage produces more compressed air than the centrifugal stage can use. A compressor bleed valve compensates for this excess airflow at lower engine RPMs by bleeding axial stage air to reduce backpressure on the centrifugal stage. The pressure relief helps prevent compressor stalls in the centrifugal stage. The compressor bleed valve is located at the 6 o’clock position of the engine. It is a pneumatic piston which references the pressure differential between the axial and centrifugal stages. The function of this valve is to prevent compressor stalls and surges in the low N1 operating range. At low N1 RPM, the valve is in the open position. At takeoff and cruise the compressor bleed valve will be closed. If the compressor bleed valve was to stick in the closed position, a compressor stall would result from the attempt to accelerate the engine to takeoff power. If the valve was to stick in the open position, the ITT would increase, the torque decrease, while N1 RPM would remain the same.

PILOT TIP: - Throttle back if continuous compressor surge is encountered. - Accelerate slowly if an engine is prone to surge. - Surge may damage the compressor and hot section. Have the engine bleed valve checked.

Combustion Section

The function of the combustion section is to create and extract energy from the hot expanding gases to drive the compressor turbine, axial compressors and the items on the accessory gear box. At the same time, it drives the power turbine and propeller to provide thrust for the aircraft. The PT6 engine utilizes an annular combustion chamber. Fuel is injected into the combustion chamber through fourteen simplex fuel nozzles by a dual manifold. Ignition is provided by two high energy igniters. The ignition system consists of a series dual low tension capacitor discharge unit energized from a solid state D.C. power source. It is designed for duty at 9 to 30 volts D.C. with a spark rate of one per second. The system stores 4.5 joules of energy and the two igniters are fired simultaneously. Even though the engine has two igniter plugs, it will start with only one operating.

Turbine Section

The PT6A uses two reaction turbines. The single stage power turbine extracts energy from the combustion gases and drives the propeller and its accessories through a planetary reduction gearbox. This combination is defined as NP. The single-stage compressor turbine extracts energy from the combustion gases to




drive the gas generated compressor and the accessory gear section which is mounted on the rear of the engine. This combination is defined as N1. A 2.3 U.S. gallon integral oil tank is formed between the accessory gear-box and the compressor air inlet plenum. The oil tank filler cap is fitted with a calibrated dipstick.

Exhaust Section

The exhaust gas from the turbine is passed into a vaneless exhaust duct and exits from the engine and into the atmosphere through two ports on opposite sides of the engine. The two heat resistant exhaust outlets are located at the 9 o’clock and 3 o’clock position.

Reduction Gear Section The second stage turbine drives a two stage planetary reduction gearbox located at the front of the engine. The primary function of the reduction gear section is to reduce the high RPM of the power turbine to a speed required for propeller operation. The reduction gear section is also used for the torque meter operation and it includes a drive section for the propeller governor, the propeller overspeed governor, and the propeller tach generator.

The Accessory Section The accessory drive section forms the aft portion of the engine. The accessory section is driven by the compressor turbine through a shaft that extends through the oil tank to the accessory gearbox. The function of the accessory section is to drive the engine and accessories. The accessory section includes:

1. The fuel control unit

2. The high pressure fuel pump 3. Lubricating pumps and scavenge pumps

4. N1 tach generator

5. DC starter generator

Engine Lubrication System The engine integral lubrication system provides a constant supply of clean oil to the engine bearings, reduction gears, accessory drives, torquemeter and propeller governor. The oil lubricates and cools the bearings and carries any extraneous matter to the oil filter where it is precluded from further circulation. A chip detector is also located in the reduction gear-box of each engine to detect and transmit a signal to the annunciator panel to warn pilots of ferrous metal particles in the reduction gearbox.




Oil Tank

The 2.3 U.S. gallon oil tank is an integral part of the compressor inlet case and is located in front of the accessory gearbox. The oil filler neck protrudes through the accessory gearbox and is closed by a cap which incorporates a quantity measuring calibrated dipstick. The markings on the dipstick correspond to U.S. quarts and indicate the quantity of oil required to top the tank to the full mark. Servicing the engine oil system primarily involves maintaining the engine oil at the proper level. Do not mix different oil brands together. The dipstick is marked in U.S. quarts and indicates the last five quarts required to bring the system up full. Access to the dipstick cap is gained through an access door on the aft engine cowl. While the airplane is standing idle, engine oil could possibly seep into the scavenge pump reservoir, causing a low dipstick reading. Therefore, the oil should be check approximately 15 minutes after engine shut down.

CAUTION

Do not mix different brands of oil when adding oil between oil changes. Different brands or types of oil may be incompatible because of the difference in their

chemical structures.




NOTE

The dipstick indicates one quart below full when the oil level is normal. Overfilling may cause a discharge of oil through the breather until a satisfactory level is reached.

Pumps

A main pressure pump is located in the tank and driven by an accessory gear on the compressor shaft. It supplies oil directly to the engine bearings and the accessory drive gears. At maximum steady state gas generator speed (N1 = 37,500 rpm), the main pressure pump maintains an oil flow of up to 90 lb/min. Oil pressure is regulated within the range 80 – 100 Psi by a pressure relief valve in the engine. Actual range on each model is dependent upon the aircraft serial number.

Oil Cooler

The system is fully automatic and incorporates a thermal sensor to regulate the amount of air flow through the oil cooler. It is equipped with a bypass valve to insure oil flow in the event the oil cooler becomes blocked. The oil-fuel heat exchanger uses hot engine oil to heat fuel before it enters the engine fuel system. When gas generator speeds are above 72% N1, and oil temperatures are between 60 and 70 degrees Celsius, normal oil pressure is between 80 and 100 psi.

Oil Temperature

A DC powered oil temperature gauge uses a resistance bulb to sense oil temperature.

Oil Pressure

Oil pressure from the pressure pump outlet line is sensed by a transmitter and sent to a combination oil pressure/oil temperature gauge located on the panel. This gauge is also DC powered.

Chip Detection

A chip detector is installed at the 6 o'clock position on the front case of the reduction gearbox. The chip detector provides the pilot with an indication on the annunciator panel if the presence of ferrous particles in the lubrication system has been attracted to the magnetic poles in the chip detector. This detector will activate a yellow light on the annunciator panel, L CHIP DETECT OR R CHIP DETECT, to alert the pilot of oil contamination. C90A aircraft are equipped with red “CHIP DETECT” annunciator panel lights.




Fuel Heater

Oil that is returned from the accessory gearbox is directed to an oil to fuel heater prior to being returned to the oil tank. The oil-to-fuel heater, mounted below the fuel pump at the rear of the engine is essentially a heat exchanger which utilizes heat from the engine lubricating oil system to preheat the fuel in the fuel system. A fuel temperature-sensing oil bypass valve regulates the fuel temperature by either allowing oil to flow through the heater or bypass it to the engine oil tank. The temperature-sensing oil bypass (thermal element) valve consists of a highly expansive material sealed in a metallic chamber. The expansion force is transmitted through a diaphragm and plunger to a piston. Since the element only exerts an expansive force, it is counterbalanced by a return spring which provides a contracting force during decreases in temperature. The element senses the temperature of the outlet fuel and, at temperatures above 21°C (70°F), starts to close the valve and simultaneously opens the bypass valve. At 32°C (C90°F), the core valve is completely closed and oil bypasses the heater core.

ENGINE FUEL SYSTEM

The engine fuel system consists of the electric low pressure fuel pump, oil to fuel heater, the high pressure engine driven fuel pump, and the fuel control unit (FCU). A flow divider sends fuel to two fuel manifolds where it is sent to the 14 fuel nozzles. If the high pressure engine driven fuel pump fails, the engine will shut down. The low pressure pump’s pressure is insufficient to run the engine.




Fuel Control Unit

The PT6 fuel control unit is a hydro-pneumatic device whose function is to supply the proper amount of fuel to the fuel nozzles during all modes of each operation. In short, it’s a N1 governor. It is calibrated for starting flow rates, acceleration, and maximum power. The FCU compares gas generator speed (N 1) with the power lever setting and regulates fuel to the engine fuel nozzles. The FCU also senses compressor section discharge pressure, compares it to rpm, and establishes acceleration and deceleration fuel flow limits. The pneumatic section of the FCU determines the flow rate of fuel to the engine for all operations. It does this by modify the amount of air pushing on the N1 governor bellows. This bellows or diaphragm reacts to the increase or decrease in P3 air by moving in one direction or the other. P3 air is introduced into the bellows so that it sets up a differential pressure on each side of the diaphragm. Therefore, any change in P3 pressure will move the diaphragm. Attached to the diaphragm is a fuel metering valve which moves as the diaphragm moves. When pressure is increased, the fuel-metering valve attached to the bellows will move in an opening direction to increase fuel flow and increase N1 rpm. As P3 pressure decreases, fuel flow also decreases which reduces the N1 rpm. The N1 governor increases or decreases P3 pressure in the bellows by varying the opening of relief orifices in the bellows.




STARTING AND IGNITION SYSTEM

The engine is started by a three-position switch located on the pilot's left subpanel placarded, IGNITION AND ENGINE START - LEFT - RIGHT - ON - OFF - STARTER ONLY. The switch is moved downward to the STARTER ONLY position to motor the engine. This is used to clear residual fuel without the ignition circuit on. The switch is spring loaded and will return to the center position when released. Moving the switch upward to the ON position activates both the starter and ignition, and the appropriate green IGNITION ON light on the annunciator panel will illuminate. When engine speed has accelerated through 50% N1 on starting, the starter is deactivated by placing the switch in the center OFF position.

AUTO IGNITION

The auto ignition system provides automatic ignition to prevent engine loss due to combustion failure. This system ensures ignition during takeoff, landing, turbulence, in icing or precipitation conditions provided the system is armed. To arm the system, move the required ENG AUTO IGNITION switches, located on the pilot's subpanel, from OFF to ARM. If for any reason the engine torque falls below approximately 400 foot-pounds, the igniter will automatically energize and the IGNITION ON light on the caution/advisory annunciator panel will illuminate. For extended ground operation, the system should be turned off to prolong the life of the igniter units. FIRE DETECTION SYSTEM The optional fire detection system on these airplanes is designed to provide warning in the event of an engine compartment fire. The system consists of a set of three photoconductive cells in each engine compartment, a control amplifier mounted on a panel on the aft side of the forward pressure bulkhead, an annunciator warning light (placarded either FIRE L ENG and FIRE R ENG or L ENG FIRE and R ENG FIRE) for each engine compartment, a test switch on the inboard side of the copilot's subpanel and a circuit breaker placarded FIRE DET on the right circuit breaker panel. The test switch on the left subpanel has four positions; OFF, 1, 2, and 3. The system may be tested any time on the ground or in flight by rotating the switch from OFF to any of the positions to activate a corresponding set of flame detectors in each nacelle. The annunciator warning lights should illuminate as the selector is rotated through each of the three positions. Failure of a light to illuminate in any one position indicates trouble in that particular detector circuit. The photoconductive cells are sensitive to infrared rays and are positioned to receive direct and reflected rays, thus providing coverage for the entire engine compartment. The cell emits an electrical signal proportional to the infrared intensity and ratio of the radiation striking the cell.




Heat level and rate of heat increase are not contributing factors in the activation on the cells. To prevent stray light rays from signaling a false alarm, a relay in the control amplifier closes only when the signal strength reaches a preset alarm level. When the relay closes, the appropriate annunciator will illuminate. When the fire has been extinguished, the cell output voltage will drop below the alarm level and the control amplifier will automatically reset. No manual resetting is required to reset the detection system.

For fire detection/protection purposes, critical areas around the engine have been divided into three zones as follows: • Zone 1 - The accessory compartment. • Zone 2 - The plenum chamber area. • Zone 3 - The engine exhaust area (hot section).

FIRE EXTINGUISHING SYSTEM

The optional engine fire extinguishing system consists of a supply cylinder, mounted on brackets behind the main spar in each wheel well, and plumbing that carries the extinguishing agent to spray nozzles located in each of the engine compartments. Each supply cylinder is charged with 2 1/2-pounds of Bromotrifluoromethane (CBrF3) and pressurized with dry nitrogen to 450 psi at 70° F. Spray nozzles are positioned under the engine exhaust area and in the accessory area. These strategically positioned nozzles discharge the entire supply of the fire extinguishing agent into the engine compartment within approximately a half second. Each fire extinguisher is actuated by its respective control switch which is located on the glareshield left and right of the warning annunciator panel. Pressing the switch will cause a squib in the cartridge to fire. This releases the extinguishing agent into the plumbing and out the nozzles. The




power to the switches is derived from the hot battery bus. To actuate the system, raise the safety-wired clear plastic switch cover and press the face of the lens. Do not attempt to restart the engine after the extinguisher has been discharged.

POWERPLANT LIMITATIONS NUMBER OF ENGINES Two ENGINE MANUFACTURER Pratt & Whitney Canada (Longueuil, Quebec, Canada) ENGINE MODEL NUMBER PT6A-21 POWER LEVERS Do not lift power levers in flight. ENGINE OPERATING LIMITS The following limitations shall be observed. Each column presents limitations. The limits presented do not necessarily occur simultaneously. Refer to Pratt & Whitney Engine Maintenance Manual for specific actions required if limits are exceeded.

FOOTNOTES: (1) Maximum permissible sustained torque is 1315 ft-lbs. Propeller speeds (N 2 ) must be set so as not to exceed power limitation. (2) For every 10 OC below -30 OC ambient temperatures, reduce maximum allowable N1 by 2.2%.

OPERATING CONDITION

SHP

TORQUE FT-LBS

(1)

MAXIMUM

OBSERVED I T T º C

N1

RPM ( 2)

N 1 %

PROP RPM N2

OIL

PRESS. PSI(3)

OIL

TEMP °C (4)

STARTING

LOW IDLE

HIGH IDLE

TAKEOFF AND MAX CONT

CRUISE CLIMB/MAX CRUISE

MAX REVERSE (8)

TRANSIENT

---

---

---

550

538

---

---

---

---

---

1315

1315 (7)

---

1500 (5)

1090 (5)

660 (6)

---

695

680

695

825 (5) (9)

---

---

---

38,100

38,100

---

38,500

---

51 (min)

70(approx)

101.5

101.5

88

102.6

---

---

---

2200

2200

2100

2420

---

40(min)

---

80 to 100

80 to 100

80 to 100

---

-40(min)

-40 to 99

0 to 99

10 to 99

0 to 99

0 to 99

0 to 99




(3) When gas generator speeds are above 72 % Ni and oil temperatures are between 60 °C and 70 OC, normal oil pressure is between 80 and 100 psi. Oil pressure between 40 and 80 psi is undesirable; it should be tolerated only for the completion of the flight, and then only at a reduced power setting. Oil pressure below 40 psi is unsafe; it requires that either the engine be shut down, or that a landing be made as soon as possible, using the minimum power required to sustain flight.

(4) For increased service life of engine oil, and oil temperature of between 74 to 80OC is recommended. A minimum oil temperature of 55 OC is recommended for fuel heater operation at take-off power.

(5) These values are time-limited to two seconds. (6) High ITT at ground idle may be corrected by reducing accessory load and or increasing N, rpm. (7) Cruise torque values vary with altitude and temperature. (8) Reverse power operation is limited to one minute. (9) High generator loads at low N1 speeds may cause the ITT transient

temperature limit to be exceeded. Observe generator load limits.

STARTER OPERATING TIME LIMIT Use is limited to 40 seconds on, 60 seconds off, 40 seconds on, 60 seconds off, 40 seconds on, then 30 minutes off.

OIL SPECIFICATION Any oil specified by brand name in the latest revision of Pratt & Whitney Service Bulletin Number 1001 is approved for use in the PT6A-21 engine.

APPROVED ENGINE OILS 7.5 Centistoke Turbine Engine Oils 5 Centistoke Turbine Engine Oils

Do not attempt to restart the engine after the extinguisher has been actuated.

EMERGENCY ENGINE PROCEDURES

All airspeeds quoted in this section are indicated airspeeds (IAS) and assume zero instrument error.

EMERGENCY AIRSPEEDS

One-Engine inoperative Best Angle-of-Climb (VXSE) 100 kts. One-Engine inoperative Best Rate-of-Climb (VySE) 107 kts.

Air Minimum Control Speed (VmcA) 90 kts. Emergency Descent 182 kts

Maximum Range Glide 125 kts




Engine Failure

NOTE To obtain best performance with one engine inoperative, the airplane must be banked 3° to 5° into the operating engine while maintaining a constant heading.

Emergency Engine Shutdown

ENGINE FIRE ON GROUND Affected Engine: 1. Condition Lever - CUT-OFF 2. Fuel Firewall Valve - CLOSED 3. Starter Switch - STARTER ONLY 4. Boost Pump - OFF 5. Fuel Transfer Pump - OFF 6. Crossfeed - CLOSED 7. Fire Extinguisher - ACTUATE (as required)

CAUTION This fire extinguisher is a single-shot system, with one

cylinder for each engine.

ENGINE FAILURE DURING GROUND ROLL

1. Power Levers – IDLE 2. Brakes - AS REQUIRED

If Insufficient Runway Remains for Stopping:

3. Condition Levers - FUEL CUT OFF 4. Firewall Shutoff Valves - CLOSED 5. Master Switch - OFF (Gang bar down) 6. Boost Pumps-OFF

ENGINE FAILURE AFTER LIFT-OFF

1. Power - MAXIMUM ALLOWABLE 2. Propeller RPM (operative engine) - FULL INCREASE 3. Airspeed - MAINTAIN (take-off speed or above) 4. Landing Gear - UP 5. Power Lever (inoperative engine) – IDLE




NOTE If the autofeather system is being used, do not retard the failed engine power lever until the autofeather system has completely stopped propeller rotation. To do so will deactivate the autofeather circuit and prevent automatic feathering.

6. Propeller (inoperative engine) - FEATHER 7. Airspeed - BEST RATE-OF-CLIMB SPEED (after obstacle clearance altitude is reached) 8. Clean-up (inoperative engine):

a. Condition Lever - CUT-OFF b. Bleed Air Valve - AS REQUIRED c. Engine Auto Ignition - OFF d. Fuel Firewall Valve - CLOSED e. Boost Pump - OFF f. Fuel Transfer Pump - OFF g. Crossfeed - CLOSED h. Generator - OFF i. Fuel Control Heat - OFF j. Autofeather Switch - OFF k. Propeller Synchrophaser - OFF

9. Electrical Load - MONITOR

ENGINE FAILURE IN FLIGHT BELOW AIR MINIMUM CONTROL SPEED (VMCA)

1. Reduce power on operative engine as required to maintain control. 2. Lower nose to accelerate above VMCA- 3. Adjust power as required. Secure affected engine as in EMERGENCY ENGINE SHUTDOWN.

ENGINE FLAMEOUT (2nd Engine) 1. Power Lever - IDLE 2. Propeller Lever - DO NOT FEATHER

3. Condition Lever - FUEL CUT OFF 4. Conduct Air Start Procedures.

NOTE The propeller will not unfeather without engine operating.

ENGINE OUT GLIDE 1. Landing Gear – UP 2. Flaps - UP 3. Propellers - FEATHERED 4. Airspeed - 125 KNOTS




WARNING! Determine that procedures for re-starting first and second failed engines are

ineffective before feathering second engine propeller.

The Glide Ratio is 1.8 nm for each 1000 feet of altitude.

ABNORMAL ENGINE PROCEDURES

Low Oil Pressure

Oil pressure values between 40 and 80 psi are undesirable and should only be tolerated for the completion of the flight. In this situation, the engine should be operated at reduced power settings. Oil pressure values below 40 psi are unsafe and require that the engine be shut down, or that a landing be made at the nearest suitable airport, using the minimum power required to sustain flight.

Air Start

STARTER ASSIST CAUTION The pilot should determine the reason for engine failure before attempting an air start.

Above 20,000 feet, starts tend to be hotter. During engine acceleration to idle speed, it may become necessary to move the condition lever periodically into CUT-OFF in order to avoid an over-temperature condition.

All electrical loads that are not consistent with flight conditions should be reduced.

1. Cabin Temp Mode - OFF, Blower - AUTO 2. Radar - STANDBY or OFF 3. Windshield Heat - OFF 4. Power Lever - IDLE 5. Condition Lever - CUT-OFF 6. Fuel Panel - CHECK:

A. Fuel Firewall Valve - OPEN B. Boost Pump - ON C. Transfer Pump - ON D. Crossfeed - AUTO

NOTE If conditions permit, retard operative engine ITT to 50° below redline to reduce the possibility of exceeding ITT limit. Cross -generator air starts normally increase ITT about 50°C on operating engine.




7. Ignition and Engine Start Switch - ON (up); Check IGNITION Annunciator - ON 8. Condition Lever - LOW IDLE (8 seconds after start switch ON) 9. Ignition and Engine Start Switch - OFF (N1 above 51%) 10. Generator - RESET (hold for one second) then - ON 11. Propeller - AS REQUIRED 12. Power Lever - AS REQUIRED 13. Fuel Control Heat – ON 14. Electrical Equipment - AS REQUIRED

WINDMILLING ENGINE AND PROPELLER (No Starter Assist) 1. Cabin Temp Mode - OFF; Blower - AUTO 2. Radar - STANDBY or OFF 3. Windshield Heat - OFF 4. Power Lever - IDLE 5. Propeller - 2200 RPM 6. Condition Lever - CUT-OFF 7. Fuel Panel - CHECK

A. Fuel Firewall Valve - OPEN B. Boost Pump - ON C. Transfer Pump - ON D. Crossfeed - AUTO

8. Generator (inoperative engine) - OFF 9. Airspeed - 140 knots minimum 10. Altitude - BELOW 20,000 FEET 11. Auto-ignition Switch - ARM 12. Condition Lever - LOW IDLE (8 seconds after auto ignition is armed) 13. Power and Propeller Levers - AS REQUIRED (after ITT has peaked) 14. Generator - RESET (hold for one second) then ON 15. Auto Ignition Switch - OFF 16. Fuel Control Heat – ON 17. Electrical Equipment - AS REQUIRED

ONE-ENGINE-INOPERATIVE LANDING

When it is certain that the field can be reached: 1. Flaps - APPROACH 2. Landing Gear - DOWN 3. Propeller Control - FULL FORWARD

4. Airspeed - 110 KNOTS

When it is certain there is no possibility of go-around: 5. Flaps - DOWN 6. Airspeed - 100 KNOTS 7. Execute Normal Landing

NOTE

Single-engine reverse thrust may be used with caution after touchdown on smooth, dry, paved surfaces.




ONE-ENGINE-INOPERATIVE GO-AROUND

1. Power - MAXIMUM ALLOWABLE 2. Flaps - UP 3. Landing Gear - UP 4. Airspeed - 107 KNOTS

WARNING!

Level flight might not be possible for certain combinations of weight, temperature, and altitude. In any event, DO NOT attempt a one -engine go-

around after flaps have been fully extended. ENGINE SYSTEM QUESTIONS

1. What does the term "free-turbine" refer to?

.

2. N1 refers to RPM of? .

3. The PT6A engine power section consists of: A. One compression stage and four turbine stages.

B. A two-stage reaction turbine. C. A one-stage turbine and a centrifugal compressor. D. Twin-spool, two-stage turbines.

4. If a chip detector light illuminates, you must do one of the following:

A. Continue the flight and have the filter checked after landing. B. Reduce torque to 500 foot-pounds for the remainder of the flight. C. Check engine instruments and, if normal, no action is required. D. Shut the engine down and land as soon as practical.

5. What is another name for T5 temperature and what gauge can it be read on? .

6. Bleed Air comes from what station on the engine? .

7. When is the best time to check the oil? .

8. T or F: Circle the correct answer.

T F The N1 gauge is marked in percent of gas generator RPM. T F Temperature and torque are two separate limitations. T F Fuel control heat is used to warm P3 air going into the F.C.U. to keep ice particles from blocking the reference air line. T F Your hand should be on the ignition and start switch during a start. T F Although the engine has two igniter plugs, it will start with only one operating. T F ITT, N1, and prop RPM are all self-generating engine instruments.




9. The Pratt & Whitney PT6A-21 engine is rated at: A. 550 SHP B. 850 SHP C. 500 SHP D. 600 SHP

10. During a ground start of the right engine, the IGNITION ON light should illuminate: A. At 10% N1 rpm. B. When the condition lever is moved to LO IDLE. C. At a stabilized 16% N1 D. When the start switch is moved to the IGNITION and ENGINE START position.

11. T or F: The compressor bleed valve is designed to prevent compressor stalls at reduced power.

12. Another name for bleed air is? .

13. What is the approximate engine out glide speed? _

14. T or F: The power turbine and N1 shafts turns in opposite direction.

15. What speed is the compressor turning at 100% N1?

16. What are the following engine limits for the engine during takeoff? ITT TORQUE Np N1

17. The Low Idle ITT limit of the engine is °C.

18. On a hot day while awaiting take-off clearance, you see the ITT above the Low Idle limit, what should you do?

.

19. T or F: Illumination of a CHIP DETECT annunciator indicates a metal contamination in the engine oil supply.

20. Oil pressure values below psi are unsafe and require that the engine be shut down.

21. The fire detection system on these airplanes is designed to provide warning in the event of a fire in the: A. Engine compartment B. Nose compartment C. Wheel well D. All of the above.




22. What are the memory items for an emergency engine shutdown:

23. True or False. Circle the correct answer.

T F The N1 gauge is marked in percent of gas generator RPM. T F Temperature and torque are two separate limitations. T F The condition levers should be milked to keep ITT temperatures within limits on a normal ground start. T F It is more important to have your hand on the ignition and start switch during a start than to have your hand on the condition lever. T F Even though your engine has two ignition plugs, it will start with only one operating. T F ITT,N1 and prop RPM are all self-generating engine instruments.

24. What caution is there regarding the addition of oil to your engine?




Chapter 6

Propeller System OBJECTIVES


1. Identify the major components of the propeller system. 2. Describe the operation of the propeller governor, overspeed governor and the fuel topping governor. 3. Explain onspeed, overspeed and underspeed conditions. 4. Describe the feathering process. 5. Explain the use of "Beta". 6. Explain the autofeather system and describe its operation. 7. Understand the emergency procedures.

GENERAL

The King Air C-90A utilizes a three or four blade propeller. The propellers are constant speed, full feathering, and reversible. They are controlled by engine oil from a single acting, engine-driven governor backed by an overspeed governor. This hydraulic action controls the propeller governor which boosts engine oil pressure to move a piston in the propeller dome that regulates the blade angle for constant speed setting in all flight attitudes and speeds. Centrifugal counterweights and feathering springs drive the propeller blades into the feather or high pitch position. The centrifugal counterweights on each blade, in conjunction with a feathering spring, increase pitch (decrease rpm) to the feathered position as governor oil pressure is relieved. The feathering spring completes the feathering operation when centrifugal twisting moment is lost as the propeller stops rotating. The propeller automatically feathers on engine shutdown, preventing the free turbine from windmilling. However, if an engine fails in flight, the propeller will not feather because of the wind-milling effect and governor action. Feathering in flight should be manually selected by using the propeller control lever. An automatic feathering system is installed which will immediately dump oil from the propeller hub if the oil pressure drops below 6.5 psi.

PILOT TIP

Always tie down the propellers when parked. Unrestrained props tend to windmill and prolonged windmilling at zero oil pressure will result in bearing damage.




BASIC PRINCIPLES

Constant-speed propellers operate in three conditions which are controlled by a propeller governor. They are:

Onspeed Overspeed

Underspeed

Onspeed This is when the selected rpm and actual rpm are the same. Overspeed This is when the actual rpm is greater than the selected rpm. Underspeed This is when the actual rpm is less than the se lected rpm.

PROPELLER GOVERNOR

The King Air is equipped with three propeller governors. They are the primary governor, the overspeed governor and the fuel topping governor.

PRIMARY GOVERNOR

The normal RPM range of the primary governor is from 1800 RPM two 2200 RPM. The primary governor is needed to convert a variable pitch propeller into a




constant speed propeller. It does this by changing blade angle to maintain the propeller speed the pilot has selected. For example, if the propeller control is set at 1900 RPM in normal cruising flight and a descent is initiated without changing power, the airspeed will increase. This decreases the angle of attack of the propeller blades causing less drag on the propeller. As a result, the RPM’s begin to increase.

The governor will sense this "overspeed" condition and increase blade angle to a higher pitch. The higher pitch increases the blade's angle of attack, slowing it back to 1900 RPM, or "onspeed." If the airplane changes from cruise to climb airspeeds without a power change, the propeller RPM tends to decrease, but the governor responds to this "underspeed" condition by decreasing blade angle to a lower pitch, and the RPM returns to its original value. Thus the governor gives "constant speed" characteristics to the variable pitch propeller. Power changes, as well as airspeed changes, cause the propeller to momentarily experience overspeed or underspeed conditions, but once more the governor reacts to maintain the onspeed condition. There are times, however, when the primary governor is incapable of maintaining selected RPM. To help explain this situation, imagine an airplane approaching to land with its governor set at 1900 RPM. As power and airspeed are both reduced, underspeed conditions exist which cause the governor to decrease blade angle to restore the onspeed condition. If blade angle could decrease all the way to 0º or even reverse, the propeller would create so much drag on the airplane that aircraft control would be dramatically reduced. The propeller, acting as a large disc, would blank the airflow around the tail surfaces, and a rapid nose-down pitch change would result. To prevent these unwanted characteristics, a low pitch stop in installed. As the blade angle is decreased by the governor, eventually the low pitch stop is reached, and the blade angle becomes fixed and cannot continue to a lower pitch. The governor is therefore incapable of restoring the onspeed condition, and propeller RPM falls below the selected governor RPM setting.

LOW PITCH STOP

Whenever the propeller rpm is below the selected governor the propeller rpm, the propeller blade angle is at the low pitch stop. (Assuming the prop is not feathered) For example, if the propeller control is set at 1900 RPM but the propeller is turning at less than 1900 RPM, the blade angle is at the low pitch stop. Normally, the low pitch stop is simply at the low pitch limit of travel, determined by the propeller’s construction. But with a reversing propeller, the extreme travel in the low pitch direction is past 0 ° , or into reverse and negative blade angles. Consequently, the low pitch stop on this propeller must be designed in such a way that it can be removed or repositioned when reversing is desired. The low pitch stop is created by mechanical linkage sensing the blade angle. The linkage causes a valve to close to stop the flow of oil coming into the propeller dome. Since this oil causes low pitch and reversing, once it is blocked off a low pitch stop has been created. The low pitch stop




valve, commonly referred to as the "beta" valve, is quite positive in its mechanical operation. Furthermore, the valve is spring loaded to provide redundancy in the event of mechanical loss of beta valve control. The position of the low pitch stop is controlled from the cockpit by the power lever. Whenever the power lever is at idle or above, this stop is set at approximately 15º blade angle. But bringing the power lever aft of idle progressively repositions the stop to blade angles less than 15°. Keep in mind that just because the low pitch stop has been moved back to smaller angles than 1 5 ° , this only affects the actual blade angle when it is on the low pitch stop. If the propeller RPM is still on the selected governor setting bringing the power lever aft of IDLE will not cause the propeller to reverse. Only when the propeller RPM is below the selected governor RPM does reversing actually occur when the power lever is brought aft. This is because in this condition the blade angle is on the low pitch stop, which is being repositioned into the reverse range. The region between 15º and 5º blade angle is ref erred to as the “beta for taxi" range. In this range, the engine's compressor speed N1 remains at the value it had when the power lever was at IDLE (52% to 70%, based on condition lever position and propeller configuration). From +5° to -9º blade angle, the N1 speed progressively increases to a maximum value at -9° of approximately 85% N1. This region, designated by red and white stripe on the power lever gate, is referred to as the "beta plus power" ranger and ends at maximum reverse.

SECONDARY LOW PITCH STOP

The secondary low pitch stop acts as a backup to prevent the blade angle from decreasing below a minimum safe value if the primary low pitch stop fails

OVERSPEED GOVERNOR

The overspeed governor provides protection against excessive propeller speed in the event of a primary governor malfunction. Since the PT6's is driven by a free turbine (independent of the engine's compressor) overspeed can rapidly occur if the primary governor fails. The operating point of the overspeed governor is set 4% greater than the primary governor’s maximum speed. Since the maximum speed selected on the primary governor is 2200 RPM, then the overspeed governor is set at 2288 RPM. As a runaway propeller's speed reaches 2288 RPM, the overspeed governor will begin increasing blade angle to a higher pitch, to prevent the RPM from continuing its rise. A propeller tachometer that stabilized at approximately 2288 RPM would indicate failure of the primary governor and proper operation of the overspeed governor. A test switch will reset this point of the overspeed governor down to approximately 2000 RPM f or a preflight check.




FUEL TOPPING GOVERNOR

If the propeller sticks or moves too slowly during a transient condition causing the propeller governor to act too slowly to prevent an overspeed condition, the power turbine governor, contained within the constant speed governor housing, acts as a fuel topping governor. If the propeller overspeeds the fuel-topping governor will vent air pressure from the Fuel Control Unit reducing fuel flow to the engine. The FTG will reduce fuel flow when the propeller overspeed reaches approximately 106% of the selected propeller rpm. Since the FTG uses the same flyweights and pilot valve mechanism as the primary governor, the fuel-topping governor will not be operational if the primary governor fails. In this case, prop overspeed will be controlled by the backup overspeed governor. During operation in the reverse range, the fuel topping governor is reset to approximately 95% propeller rpm before the propeller reaches a negative pitch angle. This ensures that the engine power is limited to maintain a propeller rpm somewhat less than that of the constant speed governor setting. The constant speed governor therefore will always sense an underspeed condition and direct oil pressure to the propeller servo piston to permit operation in Beta and reverse ranges.

PROPELLER FEATHERING

The propellers installed on the King Air are full feathering props. Using normal oil pressure, the propellers can be feathered manually, or with the Autofeather system. By placing the propeller control lever aft into the feathered detent, the pilot valve is mechanically lifted and dumps oil from the propeller dome into the reduction gearbox. This loss of oil pressure allows the centrifugal flyweights and feathering springs to rapidly drive the propeller to feather. If the pilot fails to feather the propellers during shutdown, the oil pressure will decrease and the centrifugal force of the counterweights and springs will eventually feather the propeller. However, this is not the recommended procedure.

AUTOFEATHER

The automatic feathering system provides a means of immediately dumping oil from the propeller servo to enable the feathering spring and counterweights to start the feathering action of the blades in the event of an engine failure. Although the system is armed by a switch on the subpanel, placarded AUTOFEATHER – ARM – OFF – TEST, the completion of the arming phase does not occur until both power levers are advanced above 90% N1 at which time both the right and left indicator lights on the caution/advisory annunciator panel indicate a fully armed system. The annunciator panel lights are green, and placarded: L AUTOFEATHER and R AUTOFEATHER. The system will remain inoperative as long as either power lever is retarded below 90% N1 position. The system is designed for use only during take off and landing and should be turned off when establishing cruise climb. If an engine fails while the system is armed and engine torque begins to drop off below 400 foot- pounds, a switch on the failed




engine opens and disarms the autofeather system for the opposite engine. Disarming of the Autofeather portion of the operative engine is further indicated when the annunciator indicator light for that engine extinguishes. If the torque on the failed engine continues to drop below approximately 200 ft-lbs, the oil is dumped from the servo and the feathering spring rapidly starts the blades toward the feather position.

PROPELLER BETA AND REVERSING

When the power lever controls are lifted for placement in the reverse range, the power levers actuate the Beta valve to direct governor pressure to the propeller piston, decreasing blade angle through zero and into a negative range. The travel of the propeller servo piston is fed back to the Beta valve to null its position and, in effect, provide many negative blade angles all the way to f ull re verse. The opposite will occur when the power lever is moved from full reverse to any forward position up to idle, therefore providing the pilot with manual blade angle control for ground handling. As a precaution against overtorquing the engines or developing asymmetrical thrust, an RVS NOT READY light is located in the pedestal annunciator panel. Power to the warning light switches is supplied through the landing gear control switch when the landing gear is in the DOWN position. When both propeller levers are in the high rpm position, the switches are open and the warning light is out. When either propeller lever is moved from the high rpm position, its respective warning switch closes to illuminate the RVS NOT READY light in the pedestal annunciator panel. The prop levers must be in the high RPM position to ensure constant reversing characteristics.




PILOT TIP

Propellers should be moved out of reverse by 40 knots to minimize blade erosion.

PROPELLER SYNCHROPHASER

The Type I propeller synchrophaser automatically matches the right slave propeller and maintains the blades of one propeller at a predetermined position relative to the blades of the other propeller. To prevent the right propeller from losing excessive rpm if the left propeller is feathered while the synchrophaser is on, the synchrophaser is limited to approximately ±30 rpm from the manual prop control setting. Normal governor operation is unchanged but the synchrophaser will continuously monitor propeller rpm and reset the governor as required. A magnetic pickup mounted in each propeller overspeed governor transmits electric pulses to a transistorized control box. The control box converts any pulse rate differences into correction commands, which are transmitted to an actuator motor. The motor then trims the right propeller governor through a flexible shaft to exactly match the left propeller. A toggle switch, installed on the instrument panel, turns the system on. With the switch off, the actuator automatically runs to the center of its range of travel before stopping to assure that when next turned on the control will function normally. To operate the system, synchronize the propeller in the normal manner and turn the synchrophaser on. The right propeller rpm and phase will automatically be adjusted to correspond with the left. To change rpm, adjust both propeller controls at the same time. This will keep the right governor setting within the limiting range of the left propeller. If the synchrophaser is on but is unable to adjust the right propeller to match the left, the actuator has reached the end of its travel. Turn the synchrophaser switch off (allowing the actuator to run to the center of its range and the right propeller to be governed by the propeller lever), synchronize the propellers manually and turn the synchrophaser switch on.

The Type II propeller synchrophaser system automatically matches the rpm of both propellers as a result of maintaining a specific phase relationship between the blades of the left and right propellers. The control box senses pulses which are generated by pickups mounted at identical locations on both engines. Ferrous metal targets, mounted on the propeller spinner bulkheads, provide the pulse reference for the pickups. Adjusting the RPM’s of the propellers is accomplished by the control box with correction commands to each propeller governor. The governor servo can increase but never decrease the speed set by the propeller control lever. The rpm of one propeller will follow the changes in rpm of the other propeller over the predetermined holding range of the governor. (Approximately 25 rpm) This limited holding range prevents either propeller from losing more than a limited rpm if the rpm of the other propeller is manually reduced, such as in power changes or propeller feathering, while the synchrophaser is on. The synchrophaser system is controlled through a toggle switch placarded PROP SYNCH-ON-OFF. To operate the system, synchronize the propellers in the normal manner and turn the synchrophaser on. To change




rpm, adjust both propellers at the same time. This will keep the setting within the holding range of the system. If the synchrophaser is on, but will not synchronize propellers, the propeller speeds are not within the limits required for the system to assume control. Turn the synchrophaser off, synchronize the propellers manually, and then turn the synchrophaser on.

PROPELLER CARE

Avoid operating the airplane on loose stones or gravel surfaces which can be disturbed by propeller blast during a full power takeoff. This type operation can damage the propeller blades and may produce fatigue cracks which can result in blade failure. When taking off on a loose surface, minimize blade damage by allowing the airplane to start the takeoff roll before applying full power. Always remove nicks, gouges and scratches on the propeller leading or trailing edges or on the blade surfaces. Even a small nick is detrimental, especially if it is located in the outer 18 inches of the propeller diameter. This is the blade area subject to the highest vibration and stress.

PILOT TIP Do not move the airplane by pulling or pushing on the propellers.

LIMITATIONS

Propeller Rotational Speed Limits Transients not exceeding 5 seconds-2420 rpm Reverse-1900 rpm All other conditions- 2200 rpm

Propeller Rotational Overspeed Limits The maximum propeller overspeed limit is 2420 rpm and is time-limited to five seconds. Sustained propeller overspeeds faster than 2200 rpm indicate failure of the primary governor. Sustained propeller over-speeds faster than 2288 rpm indicate failure of both the primary governor and the secondary governor, and such overspeeds are unapproved.

PROPELLER EMERGENCY PROCEDURES

Primary Governor Failure PROPELLER 2,288 RPM 1. Power Lever ADJUST AS NECESSARY 2. Prop Sync OFF

Propeller Blade Angle Stuck (FIXED PITCH PROP)

CAUTION Do not conduct emergency engine shutdown.




1. Airspeed and Power - ADJUST TO MAINTAIN POSITIVE THRUST 2. Propeller Lever (Affected Engine) - FULL FORWARD

PROPELLER ABNORMAL PROCEDURES

Propeller will not manually feather 1. Autofeather Switch - HOLD IN TEST

PROPELLER EXPANDED PROCEDURES

Overspeed Governor/Vacuum and Pneumatic Check 1. Propeller Levers - FULL FORWARD 2. Propeller Test Switch - HOLD TO TEST 3. Left Power Lever - 2000 RPM 4. Left Overspeed Governor/Vacuum and Pneumatic - CHECK (2000 ± 40) VAC 4.3-5.9 PNEU 12-20 5. Left Power Lever - IDLE 6. Right Power Lever - 20000 RPM 7. Right Overspeed Governor/Vacuum and Pneumatic - CHECK (2000 ± 40) VAC 4.3-5.9 PNEU 12-20 8. Propeller Test Switch - RELEASED

Autofeather Test 1. Power Levers - 500 ft-lb torque. 2. Autofeather Switch - Hold to test position. 3. Power Levers - Retard individually.

A. 400 ft.-lb - Opposite annunciator extinguished. B. 200ft.-lb - Autofeather annunciator light will cycle on and off.

4. Power Levers - Both idle. 5. Autofeather Switch - Armed.

PROPELLER SYSTEM QUESTIONS

1. The primary propeller governor has a governing range of RPM to RPM.

2. The overspeed governor is set to RPM.

3. T or F: The prop control levers should be full forward prior to selecting

reverse.

4. The overspeed governor is reset to what RPM for testing?

5. T or F: Moving the propeller lever into reverse without the engine running will damage the reversing linkage.




6. With the auto feather system armed during an engine failure, the propeller of the failed engine will feather at lbs of torque.

7. If the actual propeller RPM is lower than the selected RPM, what speed

condition is the prop governor in? A. Underspeed B. On Speed C. Overspeed

8. When will the prop reverse not ready annunciator light illuminate?

9. The type I synchronizer/synchrophaser system maintains both props at the same RPM by adjusting RPM of the: A. RIGHT PROP B. LEFT PROP

10. W hen using maximum reverse power at HI IDLE and full increase rpm, you would expect a maximum propeller rpm of:

A. 2000RPM B. 1900RPM C. 2420RPM D. 2288RPM




Chapter 7

Pressurization and Environmental Systems

OBJECTIVES


1 . Identify the components in the pressurization system. 2 . Explain the operation of the pressurization system. 3 . Recognize pressurization system emergencies. 4 . Identify the components in the environmental system. 5 . Explain the operation of the heating and air conditioning system. 6 . Explain the operation of the emergency oxygen system.

INTRODUCTION

This chapter describes the operation of the pressurization and environmental systems of the C-90A. Pressurization allows the altitude of the cabin to be lower than the altitude of the aircraft without the need for supplemental oxygen. Whenever cabin altitude and aircraft altitude are identical, there is no pressure differential . Pressure differential is measured in "pounds per square inch differential" (psid). This is the difference between inside cabin pressure, and outside ambient pressure. Whenever the inside cabin pressure is the greater than the outside ambient pressure, then the differential is a positive number. If cabin pressure is less than ambient pressure, then the differential is a negative number. So at 5.1 psid the cabin can be at sea level with the aircraft at approximately 11,000 feet. With the cabin at 12,000 feet, the aircraft can climb to nearly 30,000 feet before maximum differential is reached. Although the King Air's pressure vessel is designed to withstand a normal maximum differential of 5.1 psid, the minimum allowable differential is 0. This means the aircraft structure cannot withstand a negative differential. If atmospheric pressure exceeds cabin pressure, a "negative pressure" relief diaphragm in the outflow valve opens to allow atmospheric pressure to relieve cabin negative pressure. "Pressure vessel" is that part of the aircraft cabin designed to withstand the pressure differential. In the King Air, the pressure vessel extends from the forward pressure bulkhead located between the cockpit and nose section to a rear pressure bulkhead located just aft of the cabin baggage compartment. The aircraft’s exterior skin makes up the outer seal. Windows are of round design for maximum strength. All cables, wire bundles, and plumbing passing through the pressure vessel boundaries are sealed to reduce leaks. "Environmental system " refers to the devices which control the pressure vessel's environment. Along with ensuring a circulation of air, this system controls temperature by utilizing heating and cooling devices as needed.




HEATING, COOLING AND PRESSURIZATION - DESCRIPTION AND OPERATION Cabin bleed air heating is accomplished by extracting bleed air from the compression stage (P3) of each engine and mixing it with ambient air in the flow control unit of each engine. A flow control unit mounted on the forward side of the firewall in each nacelle regulates the mixture of engine bleed air with ambient air from the cowling intake to produce a total airflow of 14 pounds per minute from both the right and left engine units. Bleed air comprises as much as ten pounds of the total airflow on cold days and as little as six pounds on hot days. The bleed air control valve is energized by a bleed air switch on the copilot’s subpanel. The ambient air control solenoid valve is energized closed on the ground by a landing gear safety switch on the left main landing gear to provide only warm bleed air to the cabin. When the airplane lifts off the ground, the landing gear safety switch de-energizes and immediately opens the left ambient air control valve. Approximately six seconds later the right ambient air control solenoid




valve opens. Air is ducted into the cabin through or around the air-to-air heat exchangers in the wing center section leading edges. Control of the air bypassing the air to air heat exchanger or being routed through the heat exchangers is accomplished by regulating the position of the bleed air bypass valves. These can be adjusted either manually or automatically by using the appropriate switch on the copilot’s subpanel. At the juncture of the bleed air l ines under the cabin floor on the right side of the fuselage, a check valve is installed to prevent the loss of pressure should either engine fail. The bleed air line is routed forward along the right side of the fuselage to a mixing plenum just forward of the copilot’s rudder pedals. Here the bleed air is mixed with recirculated cabin air. The bleed air lines from the engine compartment to the mixing plenum are wrapped with insulation and aluminum tape to reduce heat loss to a minimum. The air from the mixing plenum is routed through ducts behind the instrument panel to outlets on each side of the cockpit and to the defroster outlets for the windshield. A valve to each outlet and in the defroster duct controls the flow of heated air into the cockpit. These valves are regulated by push-pull controls on the subpanel. Low pressure ducting extends aft from the mixing plenum and distributes the conditioned air through the floor and overhead outlets on each side of the cabin. A butterfly valve located in the heated air duct is controlled by the CABIN AIR control knob on the copilot's sub-panel. When this knob is pulled out, only a minimum amount of warm air is permitted to pass through the valve to the cabin floor outlets, thereby increasing the amount of warm air available to the pilot and copilot heat outlets and to the defroster. At cruise power, the heating capacity of the system is sufficient to maintain a comfortable cabin temperature at ambient temperatures of -45°F.

Heating Temperature Control – Description and Operation The temperature control system consists of a cabin temperature mode selector switch, a manual temperature switch, a temperature control box, a cabin temperature sensor, and two heat exchanger bypass valves. The cabin temperature mode switch has four positions; MANUAL HEAT, MANUAL COOL, OFF and AUTO. The evaporator has a two-speed fan for air distribution, which is controlled by a three position VENT BLOWER switch on the subpanel. Positions on the VENT BLOWER switch are: AUTO, LOW and HIGH. The fan will operate in low speed when the mode switch is positioned to AUTO, MANUAL HEAT or MANUAL COOL.

Automatic Operation When the AUTO mode is selected, the heating and air-conditioning system is automatically controlled through the temperature control box. A signal from the temperature control box is transmitted to the bleed air bypass valves in the wing center section. Here the engine bleed air is regulated by the bypass valves to




control the amount of bleed air bypassing the air-to-air heat exchangers. When a signal from the temperature control box drives both bleed air bypass valves to

the maximum cool position, the refrigerant compressor clutch and condenser blower will energize. The clutch and fan will remain energized until the left valve rotates back past the 30° position. At this position, the micro switch on the valve operates to de-energize the clutch fan. A thermal switch is wired into the AUTO mode circuit to prevent the clutch and condenser blower from being energized until the ambient temperature is above 50°F, even though a cool signal is sent from the temperature control box.

MANUAL HEAT OPERATION

When the cabin temperature mode switch is in the MANUAL HEAT position, the temperature is controlled by selecting the position of the bypass valves with the momentary increase/decrease (MANUAL TEMP) control switch. When the MANUAL TEMP selector is switched to INCR, the left bypass valve is driven open to allow the engine bleed air/ambient air mixture to be routed around the heat exchanger for increased cabin heating. The switch must be held in the INCR position to actuate the bypass valves because the valves will stop moving when the MANUAL TEMP switch is released. If sufficient heating is not obtained by full actuation of the left bypass valve, an integral limit switch in the valve will close and the right bypass valve will begin to move. Allow approximately 30 seconds for each valve to drive to the full open or full closed position. When the airplane is on the ground, the ambient air shutoff valves are closed by actuation of the landing gear safety switch. This exclusion of ambient air permits all of the heat from the engine bleed air to be used for cabin heating. When the airplane lifts off the ground, the safety switch opens the circuit to the left ambient air valve. In order to prevent a pressure surge in the cabin, the right valve will open a few seconds after the left valve through a time delay circuit.

ELECTRIC HEAT

During extremely low temperature or low power settings, additional heating is available from an electrical heater containing eight heating elements rated at 1,000 watts each. In the ENVIRONMENTAL group on the copilots subpanel is the ELEC HEAT switch with three positions: GRD MAX - NORM - OFF. This switch is solenoid-held in GRD MAX position when on the ground and will drop down to the NORM position at lift-off when the landing gear safety




switch is opened. The maximum output of the electrical heater is 27,300 BTU during ground operation with all heater elements operating. Only four elements are available during flight for a total output of 13,650 BTU. The OFF position turns off all electric heat and leaves cabin heating to be provided by bleed air. The airplane electrical system is protected against an overload by a lockout circuit that prevents use of the electrical heater during operation of the propeller heat, engine air inlet heat, or windshield heat. A differential pressure switch mounted adjacent to the vent blower senses blower operation to prevent use of the electrical heater unless the blower is also operating.

FRESH AIR VENTILATION

Fresh-air ventilation is provided from two sources. One source, which is available during both the pressurized and the unpressurized mode, is the bleed air heating system. This air mixes with recirculated cabin air and enters the cabin through the floor registers. The volume of air from the floor registers is regulated by using the CABIN AIR control knob located on the copilot's subpanel. The second source of fresh air, which is available during the unpressurized mode only, is outside air obtained from a ram air scoop on the nose (left side). The ram air enters the evaporator plenum through a flapper door. The flapper door is open during the unpressurized mode. (In the pressurized mode the flapper door is held closed by a solenoid lock.) Cabin air forced into the evaporator plenum by a blower mixes with ram air from outside and is ducted around the electric heater and mixing plenum and into the ceiling outlet duct. Air ducted to each individual ceiling eyeball outlet can be directionally controlled by moving the eyeball in the socket. Volume is regulated by twisting the outlet to open or closed.

COOLING - DESCRIPTION AND OPERATION

The King Air C 90 air-conditioning system is similar to a home or automotive system. The unit is electrically driven, has a rated capacity of 16,000 BTU, and uses a refrigerant gas. The air-conditioner system consists of five major components. They are the evaporator, condenser, expansion valve, compressor, and the receiver-dryer. During operation, the compressor, driven by a 31/3 HP electric motor, compresses the refrigerant to a high pressure, high temperature gas. This gas then goes to the condenser where cooling air is drawn in




through a louvered intake in the right side of the nose and exhausted out through louvers in the left side by a fan on the pulley end of the compressor motor. This removes heat from the gas and condenses it to a liquid. The liquid is then stored in the receiver-dryer until it is needed. The refrigerant flows to the expansion valve as a liquid. Here it is metered to the evaporator at a rate that will allow all of the liquid to evaporate and return to the compressor at a reduced pressure. The heat required for this evaporation is absorbed from the air passing over the evaporator cooling fins.

AIR CONDITIONING TEMPERATURE CONTROL - DESCRIPTION AND OPERATION

The temperature control system consists of a cabin temperature mode switch, a manual temperature selector switch, a temperature control box, a cabin temperature sensor, a duct temperature sensor, two heat exchanger bypass valves and electrical relays. The cabin temperature mode switch has four positions; MANUAL HEAT, MANUAL COOL, OFF and AUTO. The forward evaporator has a two-speed blower for air distribution, which is controlled by a three position VENT BLOWER switch on the subpanel. Positions on the VENT BLOWER switch are: AUTO, LOW and HIGH. The low speed will come on when the mode switch is turned on to AUTO, MANUAL HEAT or MANUAL COOL.

PILOT TIP To keep the air conditioner in working order, it should be operated at least 10

minutes every month. This prevents the compressor seals from drying out.

Automatic Operation When the cabin temperature mode switch is in the AUTO position, the output signal from the temperature control box drives both bleed air bypass valves. As the left bypass valve passes through the 30° position, its externally mounted micro switch actuates and energizes the air conditioner compressor and condenser blower. The compressor and fan will operate until the left valve rotates back past the 30° position towards closed. When the AUTO mode is selected, the heating and air-conditioning system is automatically controlled through the temperature control box. A signal from the temperature control box is transmitted to the bleed air bypass valves in the wing center section. Here the engine bleed air is regulated by the bypass valves to control the amount of bleed air bypassing the air-to-air heat exchangers. When a signal from the temperature control box drives both bleed air bypass valves to the maximum cool position, the air conditioning compressor and condenser blower will energize. A thermal switch is wired into the AUTO mode circuit to prevent the clutch and condenser blower from being energized until the ambient temperature is above 50°F, even though a cool signal is sent from the temperature control box. Protection from refrigerant overpressure or underpressure is provided by a circuit which incorporates high and low pressure switches.




Manual Cool Operation With the cabin temperature mode switch in the MANUAL COOL position, the air conditioning compressor and condenser fan are energized through a time delay circuit. The time delay circuit prevents the compressor clutch from being energized until 10 seconds after being de-energized to allow the refrigerant pressure in the compressor to equalize so the compressor will not be turned on under high loads. Cabin temperature is controlled by actuation of the heat exchanger bypass valves through the MANUAL TEMP switch. The rotation of the valves will stop at the position at which the MANUAL TEMP switch is released. The bypass valves must be fully closed for maximum cooling.

PILOT TIP The air conditioner will not operate in manual unless the temperature switch is

held in the decrease position for 1 minute. PRESSURIZATION - DESCRIPTION AND OPERATION

The air used for cabin pressurization is obtained by bleeding air from the compressor stage P3 of each engine . A flow control units is mounted on the forward side of each nacelle firewall. These units mix ambient air with bleed air in order to control total air flow used for pressurization. Bleed air also supplies pressure to operate the air driven instruments, the door seal, rudder boost and the surface deice system. The bleed air and ambient air from the cowling intake are mixed together by the flow control units to produce a maximum total flow of 14 pounds per minute. B le e d air c o m p r is e s a s m u c h a s 1 0 p o u n d s o f air f lo w o n cold d a y s a n d a s l it t le a s 6 p o u n d s on hot days. The bleed air lines from the engine compartment to this mixing plenum are wrapped with insulation and aluminum tape to reduce the loss to a minimum.

Flow Control Unit Each flow control unit consists of an ejector and an integral bleed air modulating valve, firewall shutoff valve, and a check valve that prevents the bleed air from escaping through the ambient air intake. The flow of bleed air through the flow control unit is controlled as a function of atmospheric pressure and temperature. Ambient air flow is controlled as a function of




temperature only. When the bleed air valve switches on the co-pilot's left subpanel are turned on, a bleed air shutoff electric solenoid valve on each flow control unit opens to allow the bleed air into the unit. As the bleed air enters the flow control unit, it passes through a filter before going to the reference pressure regulator. The regulator will reduce the pressure to a constant value of 18 to 20 psi. This reference pressure is then directed to the various components within the flow control unit that regulate the output to the cabin. One reference pressure line is routed to the firewall shutoff valve located downstream of the ejector. A restrictor is placed in the line immediately before the shutoff valve to provide a controlled opening rate. At the same time, the reference pressure is directed to the ambient air modulating valve located upstream of the ejector and to the ejector flow control actuator. A pneumatic thermostat with a variable orifice is connected to the modulating valve. This pneumostat is located on the lower aft side of the fireseal forward of the firewall. The bi metallic sensing discs of the thermostat are inserted into the cowling intake. These discs sense ambient temperature and regulate the size of the thermostat orifices. W arm air will open the orifice and cold will restrict it until, at -30ºF, the orifice will be completely closed. When the variable orifice is closed, the pressure buildup will cause the modulating valve to close off the ambient air source. An electric solenoid valve located in the line to the pneumatic thermostat is wired to the LH landing gear safety switch. When the airplane is on the ground, the solenoid valve is closed, thereby directing the pressure to the modulating valve, causing it to shut off the ambient air source. The exclusion of ambient air allows faster cabin warm-up during cold- weather operation. An electric circuit containing a time delay relay is wired to the solenoid valves to allow the LH valve to operate 2 to 3 seconds before the RH valve. This precludes the simultaneous opening of the shutoff valves and a sudden pressure surge into the cabin.




Outflow and Safety Valves Since air is delivered to the pressure vessel at a relatively constant rate of flow, the Pressurization Control System controls only the outflow of air from the pressure vessel to achieve control of the pressure differential. The outflow of pressurized cabin air is controlled by the outflow valve and safety valve utilizing a cabin pressure controller. The outflow and safety valves sense atmospheric pressure through vents that protrude through the aft pressure bulkhead. The outflow and safety valves are installed in a recessed area on the aft pressure bulkhead. Excess cabin pressure is vented into the access area immediately aft of the valves. The outflow valve is used for three purposes. First, it meters the outflow of cabin air in response to vacuum control forces from the controller. Second, it contains a preadjusted relief valve set to ensure that the cabin does not exceed 5 .1 psid. Third, it incorporates a negative pressure differential relief diaphragm which prevents the pressure differential from being negative. The safety valve also performs three functions. First, it is the "Dump Valve" which opens completely to relieve all pressure differential whenever the Pressure Control Switch is positioned in "Dump," or when the switch is in "Press" and the left landing gear safety switch is closed due to the weight of the aircraft compressing the gear strut. Second, it contains a preadjusted relief valve set to ensure that differential pressure does not exceed 5.1 psid. This provides protection against over -pressurization, should the outflow valve stick or be misadjusted. Last, like the outflow valve, it contains a negative pressure differential relief diaphragm. The pressurization controller, mounted in the cockpit pedestal, adjusts the opening of the outflow valve in order to regulate the outflow of air through the valve. It does this b y varying the amount of vacuum applied to the outflow valve. The face of the Controller contains two knobs. The left one is the rate knob and the right one is the altitude knob. With the rate knob, the pilot can select a desired cabin rate of climb and descent, from a minimum of approximately 50 fpm to a maximum of 2,000 fpm. With the altitude knob, the pilot can select a desired cabin pressure altitude, from 1,000 feet below sea level to 10,000 feet MSL. On the ground, the left landing gear safety switch closes to apply power to a normally open solenoid, which in turn closes to block off the source of vacuum to the controller. With no vacuum applied, the outflow valve moves to its spring-loaded, closed position. At liftoff the cabin will immediately begin to pressurize at the rate preset on the controller. Vacuum pressure for the pressure controller is controlled by the vacuum regulator that also regulates instrument vacuum. When the airplane is on the ground with the squat switch compressed, the cabin pressure control switch can be set to the TEST position to de-energize




the preset and safety solenoids and allow the pressure control system to function as though the airplane were in flight. The cabin pressure control switch mounted on the cockpit pedestal, contains three positions. The aft position is labeled "Test," the center position is "Press" (for "pressure"), and forward is "Dump." Normally, it is left in the center position. The switch must be lifted over a detent to go to the Dump position. When released from the Test position, it will return back to the center, due to spring force. Outside air can enter the cabin anytime the cabin pressure differential is zero and the cabin pressure control switch to set to DUMP. Ambient air is then allowed to flow into the fresh air inlet, and into the forward evaporator plenum. Cabin pressure altitude and the cabin-to-atmosphere pressure differential are indicated on the differential pressure indicator. The pressure differential is expressed in psig and the pressure altitude is expressed in thousands of feet. The climb rate indicator allows monitoring of the rate of change of cabin pressurization. If cabin pressure altitude exceeds 10,000 ft, the cabin altitude warning pressure switch closes and the warning annunciator light labeled ALT WARN will illuminate.

Oxygen System FAR’s require that anytime an aircraft flies above 25,000 feet, oxygen must be immediately available to the crew and passengers. Oxygen for flight at high altitudes is supplied by a cylinder mounted behind the aft cabin bulkhead. The cylinder is filled by a valve accessible through an access door on the right side of the fuselage. The system has two pressure gages, one located on the right side panel in the cockpit for in-flight use and one adjacent to the filler valve for checking the pressure of the system during filling. The oxygen system utilizes a 22-, 49-, or 66-cubic foot volume cylinder. Oxygen flows from the cylinder through a pressure line of copper tubing routed along the right side of the fuselage to the system regulator and shutoff valve. The shutoff valve is actuated by a push-pull type control located overhead between the pilot’s and copilot’s seats. The regulator is a constant-flow type which supplies low pressure oxygen through aluminum plumbing to the outlets and proved an adequate oxygen flow up to a cabin altitude of 30,000 feet. Each mask plug is equipped with its own regulating orifice. Normal storage of the pilots and copilots oxygen masks is in a container located on




the forward cockpit sidewalls. Diluter demand quick-donning oxygen masks are connected to the oxygen supply at all times. When the diluter demand masks are not in use they are stored in containers attached to the pilot’s and copilot’s headliner. Normal storage of the passengers' masks is in the seat-back pockets. The cabin oxygen outlets are located in the ceiling at the forward and aft ends of the cabin headliner. All masks are easily connected in by pushing the orifice in firmly and turning clockwise approximately one quarter turn. Disconnecting is easily accomplished by reversing the motion.

PILOT TIP The oxygen bottle is fully charged when it reads 1800psi when the 22 cu ft cylinder is used or 1850psi when 49 or 66 cu ft cylinder is used. Fill the oxygen system slowly by adjusting the recharging rate with the pressure regulating valve on the servicing cart because high pressure oxygen will cause excessive heating of the filler valve.

PRESSURIZATION LIMITATIONS

Cabin Differential Pressure Gage Green Arc (Approved Operating Range) 0 to 5.0 psi Red Arc (Unapproved Operating Range) 5.0 psi to end of scale

EMERGENCY ENVIRONMENTAL SYSTEM PROCEDURES

(BOLD TYPE INDICATES MEMORY ITEMS!)

Use of Oxygen WARNING! The following table sets forth the average time of useful consciousness (TUC) (time from onset of hypoxia until loss of effective performance) at various altitudes.

Cabin Pressure Altitude TUC 35,000 feet 1/2 - 1 minute 30,000 feet 1 - 2 minutes 25,000 feet 3 to 5 minutes 22,000 feet 5 to 10 minutes 12 - 18,000 feet 30 minutes or more




Loss of Pressurization In the event of pressurization loss at high altitude, don oxygen masks and descend as necessary.

HIGH DIFFERENTIAL PRESSURE Anytime the differential pressure goes into the Red Arc: 1. Cabin altitude Controller… SELECT HIGHER CABIN ALTITUDE SETTING If condition persists: 2. Bleed Air Valves… ………………………………...CLOSED 3. Cabin Pressure (after cabin is depressurized)… DUMP 4. Bleed Air Valves……………………………………OPEN

PILOT TIP The oxygen pressure provided to the passengers is not adequate for sustained

flight at cabin altitudes above 25,000 feet.

Smoke and Fume Elimination Attempt to identify the source of smoke or fumes. Smoke associated with electrical failures is usually gray or tan in color, and irritating to the nose and eyes. Smoke produced by environmental system failures is gen erally white in color, and much less irritating to the nose and eyes. If smoke is prevalent in the cabin, cabin oxygen masks should not be used.

Electrical Smoke or Fire 1. Oxygen

A. Oxygen Control Handle - PULL ON (Verify) B. Crew - DON MASKS (100% position)

2. Cabin Temp Mode - OFF 3. Vent Blower - AUTO 4. Aft Blower (if installed) - OFF 5. Avionics Master - OFF 6. Nonessential Electrical Equipment – OFF

If Fire or Smoke Ceases: 7. Individually restore avionics and equipment previously turned off. 8. Isolate defective equipment.

WARNING! Dissipation of smoke is not sufficient evidence that a fire has been extinguished. If it cannot be visually confirmed that no fire exists, land at the nearest suitable airport.

If Smoke Persists or if Extinguishing of Fire is Not Confirmed: 9. Cabin Pressure - DUMP 10. Land at the nearest suitable airport.




NOTE Opening a storm window (after depressurizing) will facilitate smoke and fume removal.

Environmental System Smoke or Fumes 1. Oxygen

a. Oxygen Control Handle - PULL ON (Verify) b. Crew - DON MASKS (100% position)

2. Cabin Temp Mode - OFF 3. Vent Blower - HI 4. Left Bleed Air Valve – CLOSED

If Smoke Decreases: 5. Continue operation with left bleed air off.

If Smoke Does Not Decrease: 6. Left Bleed Air Valve - OPEN 7. Right Bleed Air Valve - CLOSED 8. If smoke decreases, continue operation with right bleed air off.

NOTE Each bleed air valve must remain closed long enough to allow time for smoke purging to positively identify the smoke source.

Emergency Descent 1. Power Levers – IDLE 2. Propeller Levers - FULL FORWARD 3. Flaps - APPROACH 4. Landing Gear - DN 5. Airspeed - 182 KNOTS MAXIMUM ABNORMAL ENVIRONMENTAL PROCEDURES

Illumination of Cabin Door Warning Annunciator WARNING! Do not attempt to check the security of the cabin door. Remain as far from the door as possible with seat belts securely fastened until the airplane has landed. 1. If the CABIN DOOR warning annunciator illuminates, depressurize cabin (consider altitude first) by activating cabin pressurization dump switch on pedestal.

2. Do not attempt to check cabin door for security until cabin is depressurized and the airplane is on the ground.

Check security of cabin door (on the ground) by lifting cabin door step and checking position of arm and plunger. If unlocked position of arm is indicated, turn door handle toward locked position until arm and plunger are in position.




Cracked Windshield 1. If it is positively determined that the crack is on the outer panel, no action is required.

CAUTION Windshield wipers may be damaged if used on cracked outer panel. Heating elements may be inoperative in area of crack.

If it is determined that the crack is on the inner panel, descend or reset the pressurization controller to achieve 3 psi or less differential pressure within ten minutes. Visibility through the windshield may be significantly impaired.

ENVIRONMENTAL SYSTEM EXPANDED PROCEDURES

Pressurization Test 1. Bleed Air valves – Open 2. Condition Levers – High Idle 3. Cabin Altitude Selector Knob - 1000 feet below field pressure altitude 4. Rate Control selector Knob - Set index at 12-o'clock position 5. Cabin Pressurization Switch -Test position 6. Cabin VSI - CHECK FOR RATE OF DESCENT INDICATION 7. Cabin Pressurization Switch – Released 8. Cabin Altitude Selector Knob - Planned cruise altitude plus 1000 feet 9. Condition Levers – As required




OXYGEN DURATION

NOTE

A bottle pressure of 1850 psig at 15°C is fully charged (100% capacity). Read duration directly from table.

1. Read the oxygen pressure from the gage. 2. Read the IOAT (with battery ON). (Assume IOAT to be equal to

BOTTLE TEMPERATURE). 3. Determine the percent of usable capacity from the following graph (e.g.,

1100 psi at 0°C = 57%) 4. Compute the oxygen duration in minutes from the table by multiplying

the duration by the percent of usable capacity. e.g., a. Pilot and copilot plus 4 passengers = 8 people using oxygen

NOTE

A Pilot and copilot are each counted as 2 people with diluter demand masks set at 100% or NORMAL.

b. Cylinder Volume = 49 cu ft (1387 liters) c. Duration with full bottle = 41 minutes d. Duration with 57% capacity = .57 x 41 = 23 minutes

ENVIRONMENTAL SYSTEM QUESTIONS

1. When does the vent blower operate? .




2. When is the cabin temperature rheostat functional?

.

3. When is the manual temperature switch functional? .

4. Name the 3 functions of the outflow valve? , and ________________________________________.

5. What is the function of the by -pass valves located in the wing root?

________________________________________________________ _____________________________________________.

6. How long should the Manual Temp Switch be held in the Decrease

position to operate the air Conditioner? _____________________________________________________.

7. What is the normal allowable max differential pressure for the Model C

90A? .

8. Upon lift-off, the cabin fails to pressurize. List some of the possible reasons.

.

9. The airplane entry door must be in the ________________ position for flight.

10. What action should the pilot take if the outer pane of the windshield cracked?

.

11. The ALT WARNING annunciator light illuminates at: A. 10,000 ft B. 12,000 ft C. 12,500 ft D. 14,500 ft

12. List the memory items for Emergency Descent:

13.What is the UTC at 25,000 feet?




14.What will cause the electric heat to go from Ground Max to Normal

automatically? __________________________________________________________ __________________________________________________________ __________________________________________________________ .

15.T or F: With the cabin at 10,000 feet, the aircraft can climb to nearly

30,000 feet before maximum differential is reached.

16.What position should the condition levers be in for a pressurization test? High or Low

17.What position should the Vent Blower switch be in for Electrical smoke?

___________ Environmental Smoke? ____________.




Chapter 8

Landing Gear, Tires, and Brake System OBJECTIVES

With the use of this training manual the pilot will be able to:

1. Identify the major components which make up the landing gear system. 2. Identify those systems using hydraulic power. 3. Identify those systems using electrical power. 4. Identify the major components of the brake system. 5. Know the airspeed limitations of the landing gear system. 6. Identify various types of unsafe gear indications and utilize the

appropriate emergency checklist for each indication. GENERAL

The King Air C-90A utilizes an electrically operated hydraulic system. The system is controlled by a handle placarded LDG GEAR CONTROL – UP – DN on the right subpanel. The landing gear control handle must be pulled out of a detent before it can be moved from either the UP or the DN position. Visual indication of landing gear position is provided by individual green GEAR DOWN lights. The lights can be checked by depressing the lamp. A red light in the landing gear control handle indicates when the gear is in transit. Gear up is indicated when the red light goes out. This red light also comes on with the warning horn anytime all gears are not down and locked when the power levers are retarded to less than 79% N1. The bulb may be checked by a press-to-test switch mounted adjacent to the landing gear control handle. The landing gear in-transit light will indicate one or all of the following conditions: a. Landing gear handle is in the "up" position and the airplane is on the ground with weight on the landing gear. b. One or both power levers retarded below approximately 79% N1 and one or more landing gears not down and locked. Warning horn will sound. c. Any one or all landing gears not fully retracted or in the down and locked position. d. Warning horn has been silenced and will not operate.

The function of the landing gear in-transit light is to indicate that the landing gear is in transit or the position of the landing gear does not match that of the handle. It also indicates that the landing gear warning horn has been silenced and not rearmed. The light will remain on when the horn is silenced.




The up indicator, down indicator and warning horn systems are completely independent systems. A malfunction in any one system will leave the other two systems unaffected.

GROUND HANDLING TOWING

Always ensure that the control locks are removed before towing the airplane. Serious damage to the steering linkage can result if the airplane is towed while the control locks are installed. Do not tow the airplane with a flat shock strut. Even brief towing or taxiing with a deflated strut can cause severe damage. The nose gear strut has turn limit warning marks to warn the tug driver when turning limits of the gear will be exceeded. Damage will occur to the nose gear and linkage if the turn limit is exceeded. A nose gear steering stop block is installed to warn the pilot if tow limits have been exceeded. The maximum nose wheel turning angle is 48° left and right.

PILOT TIP Do not push or pull the airplane using the

propellers or control surfaces. PARKING

The parking brake may be set by pulling outward on the parking brake control, located on the extreme left side, below the pilot's subpanel, and depressing the toe portion of the pilot's rudder pedals. The parking control closes dual valves in the brake lines that trap the hydraulic pressure applied to the brakes and prevents pressure loss through the master cylinders. To release the parking brake, depress the pilot's brake pedals to equalize the pressure on both sides of the parking brake valves and push the parking brake control fully in. The tow bar connects to the upper torque knee fitting of the nose strut. The airplane is steered with the tow bar when moving the airplane by hand, or an optional tow bar is available for towing the airplane with a tug. Although the tug will control the steering of the airplane, someone should be positioned in the pilot's seat to operate the brakes in case of an emergency.

NOSE LANDING GEAR

Using differential power and brakes, the nose gear can be pivoted to its maximum angle of 48 degrees to the right or left of center, allowing the airplane to be turned within a 35'6" wing tip radius. Upon retraction, the nose landing gear




assembly is fully enclosed in the wheel well. The gear door mechanism is a mechanical design that does not require sequencing valves. Three high intensity lights are mounted on the nose gear assembly. The dual landing lights on the nose gear provide coverage of light for landing at night. The single taxi light is aimed down to illuminate the ramp area ahead of airplane during ground operations. These lights will remain illuminated with the gear up until the switch is placed in the off position. An air-oil type shock strut on the nose wheel is filled with compressed air and hydraulic fluid to absorb landing shocks and decrease any bouncing tendencies. A shimmy damper is mounted on the right side of the nose gear strut. This hydraulic cylinder dampens any nose wheel shimmy during take off and landing. A linkage connected to the rudder pedals permits nose wheel steering when the nose gear is down. Since motion of the pedals is transmitted via cables and linkage to the rudder, rudder deflection occurs when force is applied to any of the rudder pedals. With the nose landing gear retracted, some of the force applied to any of the rudder pedals is absorbed by a spring-loaded link in the steering system so that there is no movement at the nose wheel, but rudder deflection still occurs. The nose wheel is self centering upon retraction.

PILOT TIP The landing and taxi lights remain on after the gear has been retracted.

DESCRIPTION AND OPERATION - LANDING GEAR

The landing gear are retracted and extended by an electrically operated hydraulic system. The electrically operated hydraulic power pack is located in the center of the wing center section forward of the main spar. The hydraulic power pack consists of a 28-vdc motor, pump, two-section reservoir, filter, four-way selector valve, up and down selector solenoid, gear-up pressure switch and low liquid level sensor. To prevent cavitation of the pump, engine bleed air regulated to 18 to 20 psi is plumbed into the power pack reservoir and the system fill reservoir. Associated plumbing for a normal extend mode, emergency extend mode, and normal retract mode is routed from the power pack to each main gear actuator and the nose gear actuator. The plumbing for the normal extend mode and the emergency extend mode is fitted as separate plumbing to the shuttle valve in each actuator.

P I L O T T I P

If any of the following conditions exist, is likely that an unsafe gear indication is due to an unsafe gear and is not a false indication. 1. The inoperative gear down annunciator illuminates when tested. 2. The red light in the handle is illuminated.




3. The gear warning horn sounds when one or both power levers are retarded below a preset N1.

CAUTION NEVER RELY ON THE SAFETY SW ITCH TO KEEP THE GEAR DOW N. THE LANDING GEAR CONTROL SWITCH MUST BE IN THE DOWN POSITION.

LANDING GEAR CONTROLS A landing gear control switch handle, placarded UP/DN is located on the pilot’s inboard subpanel. A solenoid operated downlock latch prevents the landing gear control handle from being raised while the airplane is on the ground. The landing gear safety (squat) switch releases the latch when the airplane leaves the ground. If necessary, the latch can be manually overridden by pressing down on the red button placarded KN LOCK REL. To prevent accidental gear retraction on the ground, a safety (squat) switch on the main struts breaks the control circuit whenever the struts are compressed. CAUTION NEVER RELY ON THE SAFETY (SQUAT) SWITCH TO KEEP THE GEAR DOWN WHILE TAXIING, ON LANDING OR TAKE-OFF ROLL. ALWAYS CHECK THE POSITION OF THE LANDING GEAR SWITCH. The landing gear control circuit is protected by a 2-ampere circuit breaker located on the pilot’s inboard subpanel. Power for the pump motor is supplied through the landing gear motor relay and a 60-ampere relay circuit breaker, both of which are located in the power pack bay.

LANDING GEAR RETRACT CYCLE When the landing gear control handle is moved to the UP position, the solenoid mounted on the valve body end of the pump is energized to actuate the gear selector valve to allow system fluid under pressure from the pump to flow to the retract side of the system. The nose gear actuator will unlock when 200 to 400 psi of hydraulic pressure is applied to the retract port of the nose gear actuator. The landing gear will begin to retract after the nose gear actuator is unlocked. As the actuator pistons move to retract the landing gear, the fluid in the actuators exits through the normal extend port of the actuators and is carried back to the power pack through the normal extend plumbing. When the hydraulic fluid enters the power pack, the gear selector valve directs the return fluid to the primary reservoir. The landing gear is held in the retracted position by positive hydraulic pressure. When the system pressure reaches the high pressure limit, the gear-up pressure switch, mounted on the power pack




assembly, will interrupt current to the pump motor. This same pressure switch will actuate the pump motor should the system pressure drop to the low pressure limit. An accumulator precharged to 800 +/- 5% psi, located in the left wing inboard of the nacelle, is designed to aid in maintaining the system pressure in the gear-up mode. LANDING GEAR EXTEND CYCLE When the landing gear control handle is moved to the DN position, the solenoid is positioned to allow fluid under pump pressure to flow to the extending side of the system. As the actuator pistons move to extend the landing gear, the fluid in the actuators exits through the normal retract port of the actuators and is carried back to the power pack through the normal retract plumbing. Fluid from the pump opens a pressure check valve in the power pack to allow the return fluid to flow into the primary reservoir. When the actuators fully extend the landing gear, an internal mechanical lock in the nose gear actuator and a mechanical lock on each main gear drag brace will hold the landing gear in the down position. In this position, the internal locking mechanism in the nose gear actuator and the mechanical lock on the main gear drag braces will actuate downlock switches to interrupt current to the pump motor. The motor will continue to run until all three landing gears are down and locked. A low pressure vent valve in the power pack (open below 80 psig and closed above 80 psig) relieves any thermal expansion in the retract side of the system when the landing gear is down and locked. HYDRAULIC FLUID LOW WARNING SYSTEM A yellow HYD FLUID LOW annunciator located in the CAUTION/ADVISORY panel will illuminate in the event the hydraulic fluid level in the landing gear power pack becomes critically low. When low fluid level is indicated, the landing gear should not be extended or retracted by using the hydraulic power pack; however, the landing gear can be extended with the emergency hand pump. A sensing unit mounted on the motor end of the power pack provides the necessary switching circuitry to illuminate the low fluid light. The optically operated sensing unit has an integrated self-test circuit. The integral self-test circuit is energized by a switch on the instrument panel and functionally tests the sensing unit’s internal circuitry. MANUAL EMERGENCY OPERATION OF LANDING GEAR Manual landing gear extension is provided through a manually powered hydraulic system. An emergency hand pump, placarded LANDING GEAR ALTERNATE EXTENSION, is located on the floor between the pilot’s seat and the pedestal. The pump is used when emergency extension of the gear is required. To extend the gear with this system pull the landing gear control




circuit breaker on the pilot’s inboard subpanel and place the landing gear control handle in the DN position. Remove the emergency pump handle from the securing clip and pump the handle up and down to extend the gear. As the handle is pumped, hydraulic fluid is drawn from the hand pump suction port of the power pack into the pump and exited under pressure. Fluid under pressure from the pump is routed to the power pack hand pump pressure port and to the shuttle valve in each actuator. Fluid pressure at the shuttle valves will position the valves to allow fluid to flow into the actuator cylinders. As the actuator pistons move to extend the landing gear, the fluid in the actuators exits through the normal retract port of the actuators and is carried back to the power pack through the normal retract plumbing. The fluid routed to the power pack hand pump pressure port from the hand pump unseats the internal hand pump dump valve to allow the return fluid to flow into the primary reservoir. Continue to pump the handle up and down until the green GEAR DOWN indicator lights on the pilot’s inboard subpanel illuminate. Ensure that the pump handle is in the full down position prior to placing the pump handle in the securing clip. When the emergency pump handle is stowed, an internal relief valve is actuated to relieve the hydraulic pressure in the pump.

WARNING! After an EMERGENCY landing gear extension has been made, do not move any landing gear controls or reset any switches or circuit breakers until the cause of the malfunction has been determined and corrected. A service valve located forward of the power pack may be used, in conjunction with the emergency hand pump, to raise and lower the gear from maintenance purposes. FILL RESERVOIR A fill reservoir, located just inboard of the left nacelle and forward of the front spar, contains a cap and dipstick assembly, marked HOT/FILL, COLD/FILL, for convenience of maintaining system fluid level.

LANDING GEAR WARNING SYSTEM

Visual indication of the landing gear position is provided by two red in-transit lights, located in the landing gear control switch handle, and a green GEAR DOWN indicator light assembly labeled NOSE – L – R, located adjacent to the landing gear control switch. Illumination of the red in-transit lights indicates that the landing gear is extending or retracting; the lights go out when the gear is up. Illumination of a green GEAR DOWN light indicates that the landing gear is down and locked. The red lights in the control handle can be checked by pressing the HDL LT TEST push-button switch located to the right of the landing gear control switch. The check the GEAR DOWN lights, press the light assembly case.




The landing gear warning system is provided to warn the pilot that the landing gear is not down and locked during specific flight regimes. Various warning modes result, depending upon the position of the flaps. With the flaps in the UP or APPROACH position and either or both power levers retarded below approximately 78 - 80% N1, the warning horn will sound intermittently and the landing gear control handle lights will illuminate. The horn can be silenced by pressing the WARN HORN silence button adjacent to the landing gear control handle or on the left power lever. The lights in the landing gear control handle cannot be cancelled. The landing gear warning system will be rearmed if the power levers are advanced sufficiently. With the flaps beyond the APPROACH position, the warning horn and landing gear control handle lights will be activated regardless of the power settings, and cannot be cancelled.

TIRES

The airplane utilizes a pair of 8.5 x 10, 8 ply tubeless, rim-inflation tires on each main gear assembly. For increased service life, 10-ply-rated tires of the same size may be installed. A 6.5 x 10, 6-plyrated tire is installed on the nose gear.

PILOT TIP Tires that have picked up a film of fuel, hydraulic fluid, or oil should be washed down as soon as possible, in order to prevent deterioration of the rubber.

Maintaining proper tire inflation pressures will help prolong tire service life. Check tires frequently to maintain pressures within recommended limits, and maintain equal pressures on both tires of each dual-wheel installation. Proper inflation pressures will help avoid damage from landing shocks, contact with sharp stones and ruts, and will minimize tread wear. When inflating the tires, inspect them for cuts, cracks, breaks, and tread wear. Inflate the main wheel tires between 52-58 psi. The nose wheel tires should be inflated to between 50 and 55psi.

PILOT TIP The aircraft manufacturer does not recommend using recapped

tires on the airplane.




HYDRAULIC BRAKE SYSTEM

The hydraulic brake is designed for use with MIL-H-5606 hydraulic Fluid and to withstand 550 psi operating pressure with zero psi back pressure. The depression of either set of pedals moves the piston rod and the piston in the master cylinder attached to each pedal.. The hydraulic pressure resulting from the movement of the pistons in the master cylinders is transmitted through flexible hoses and fixed aluminum tubing to the multiple disc brake assemblies on the main landing gear wheels. This pressure forces the brake pistons on the wheel to press against the multiple linings and discs of the brake assembly. Dual parking valves are installed adjacent to the rudder pedals between the master cylinders of the pilot’s rudder pedal and the wheel brakes. After the pilot’s brake pedals have been depressed to build up pressure in the brake lines, both valves can be closed simultaneously by pulling out the parking brake control on the left subpanel. This closes the valves to retain the pressure that was previously pumped into the brake lines. The parking brake is released when the pedals are depressed briefly (to equalize pressure on both sides of the valves) and the control is pushed in, causing the valves to open. Brake system servicing is limited to maintaining the hydraulic fluid level in the reservoir mounted on the bulkhead in the upper left corner of the nose avionics compartment. A dipstick is provided for measuring the fluid level. When the reservoir is low on fluid, add sufficient quantity of approved hydraulic fluid (MIL-H-5606) to fill the reservoir to the full mark on dipstick.

Brake Wear Allowance The only other requirement related to servicing involves the wheel brakes themselves. Brake lining adjustment is automatic, eliminating the need for periodic adjustment of the brake clearance. Check brake wear periodically to assure that dimension "A", in the Brake Wear illustration, does not reach zero. When it reaches zero, you should contact your mechanic.




PILOT TIP The parking brake should be left off and wheel chocks installed if the airplane is to be left unattended. Changes in the ambient temperature can cause the brakes to release or to exert excessive pressures.

SHOCK STRUTS

With the airplane empty except for fuel and oil, the nose strut should be extended 3 to 3-1/2 inches and the main strut should be extended 3 inches.

PILOT TIP Do not fill shock struts with oxygen.

LANDING GEAR LIMITATIONS

KCAS KIAS

Maximum Landing Gear Operating Speed

182E 164R

182E 163R

Do not extend or retract landing gear above the speeds given.

Maximum Landing Gear Extended Speed

182 182 Do not exceed this speed with landing gear extended.

EMERGENCY LANDING GEAR SYSTEM PROCEDURES

ABNORMAL LANDING GEAR PROCEDURES

Hydraulic Fluid Low [HYD FLUID LO] 1. Landing Gear – ATTEMPT TO EXTEND NORMALLY UPON ARRIVING AT

DESTINATION 2. If Landing Gear Fails To Extend – SEE LANDING GEAR MANUAL

EXTENSION Landing Gear Manual Extension

If the landing gear fails to extend after placing the Landing Gear Control down, perform the following: 1. Landing Gear Relay Circuit Breaker (Pilot’s right subpanel) - PULL 2. Landing Gear Control – CONFIRM DN 3. Alternate Extension Handle – UNSTOW AND PUMP

a. Pump handle up and down until the three green gear-down annunciators are illuminated.

b. While pumping, do not lower handle to the level of the securing clip as this will result in loss of pressure.




If all three green gear-down annunciators are illuminated: 4. Alternate Extension Handle - STOW 5. Landing Gear Controls – DO NOT ACTIVATE a. The Landing Gear Control and the Landing Gear Relay Circuit Breaker MUST NOT BE ACTIVATED b. The landing gear should be considered UNSAFE until the airplane is jacks and the system has been cycled and checked.

If one or more green gear-down annunciators do not illuminate for any reason and a decision is made to land in this condition: 6. Alternate Extension Handle – CONTINUE PUMPING a. Continue to pump until maximum resistance is felt. b. When pumping is complete, leave handle at the top of the stroke, DO NOT LOWER AND STOW

Prior to Landing: 7. Alternate Extension Handle – PUMP AGAIN a. Pump the handle again until maximum resistance is felt. b. When pumping is complete, leave handle at the top of the stroke, DO NOT LOWER AND STOW

After Landing: 8. Alternate Extension Handle – PUMP WHEN CONDITIONS PERMIT a. Pump the handle again, when conditions permit, to maintain

hydraulic pressure until the gear can be mechanically secured. b. DO NOT LOWER STOW HANDLE c. DO NOT ACTIVATE THE LANDING GEAR CONTROL OR THE

LANDING GEAR RELAY CIRCUIT BREAKER d. The landing gear should be considered UNLOCKED until the

airplane is on jacks and the system has been cycled and checked. LANDING GEAR EXPANDED PROCEDURES

Practice Landing Gear Manual Extension

1. Airspeed – Below 182 KNOTS 2. Landing Gear Relay Circuit Breaker (pilot’s subpanel) – PULL 3. Landing Gear Control – DN 4. Alternate Extension Handle – PUMP UP AND DOWN UNTIL [L],

[R] AND [NOSE] ILLUMINATE AND FURTHER RESISTANCE IS FELT 5. Alternate Extension Handle - STOW

Landing Gear Retraction After Practice Manual Extension After a practice manual extension of the landing gear, the gear may be retracted hydraulically as follows:

1. Alternate Extension Handle – CONFIRM STOWED 2. Landing Gear Relay Circuit Breaker (pilot’s subpanel) – PUSH IN 3. Landing Gear – UP




Landing Gear Will Not Retract 1. Landing Gear Relay CB – CHECK IN 2. Pump Handle - STOW 3. Landing Gear Handle – UP

If gear will not retract: 4. Landing Gear Handle – DOWN 5. Landing Gear Relay CB – PULL 6. Maximum Airspeed – 182

LANDING GEAR SYSTEM QUESTIONS

1. The maximum speed for alternate gear extension with the manual system is: A. 120 K B. 130 K C. 140 K D. 182 K

2. What is the tire pressure for the mains:

Nose gear tire: .

3. T or F: Brake wear can be checked during preflight.

4. Where is the brake fluid reservoir located?

5. When is the landing gear horn silence button disabled?

6. If manually extending the landing gear, when would you stop pumping? Why?

7. Where is the landing gear relay control circuit breaker located?

8. The red light in the gear handle will illuminate when: A. The gear is not down and locked. B. The landing gear is not up and locked. C. The landing gear is in transit. D. All of the above.

9. The gear warning horn will sound when the gear is not down and:

A. Either power lever is reduced to a certain setting. B. The wing flaps are extended beyond the approach setting. C. The hydraulic system pressure falls below 1,500 psi. D. Both a and b.




10.The emergency landing gear extension system utilizes: A. A hand crank located behind the pilot's seat. B. A hand pump located in the cockpit. C. A nitrogen blow-down bottle. D. A mechanical drop-down release.

11. T or F: Once the gear has extended manually, it can be retracted normally.

12. Airspeeds for the landing gear:

A. Maximum gear extended speed KCAS. B. Maximum gear extension speed KCAS. C. Maximum gear retraction speed KCAS.

13. Is the parking brakes hydraulic or mechanical? (Circle one)




Chapter 9

Pneumatic and Vacuum System OBJECTIVES

After completing this unit, the student should be able to:

1. State the air source for pneumatic operation. 2. State the vacuum source. 3. State acceptable pneumatic and vacuum gauge readings 4. Describe pilot action to activate the surface deice system.

DESCRIPTION

The PNEUMATIC and VACUUM SYSTEMS training section of the workbook present a description and discussion of pneumatic and vacuum systems. The sources for pneumatic air, and vacuum along with acceptable gauge readings are discussed.

PNEUMATIC - DESCRIPTION AND OPERATION

The pneumatic system uses engine bleed air from the third state (P3) of the compressor. One engine can supply sufficient bleed air to operate all the systems requiring pressurized air. During operation with one engine inoperative, a check valve in the bleed air line from each engine prevents flow back through the line on the side of the inoperative engine. A pressure gage calibrated in pounds per square inch indicates air pressure available to operate the various systems. Air temperature of approximately 650°F (depending on the power setting and ambient air temperature) is bled from each engine compressor at a flow rate sufficient to produce the 18 psi of pressure required to operate the door seal, deice boots, the bleed air warning system (if installed), the hour meter and on some models, the emergency exit door seal . The bleed air for these systems comes off the compressor bleed air line at each engine. This bleed air is routed aft from the engine to a firewall shutoff valve, through a check valve and on to a pressure regulator valve. The pressure regulator valve is located adjacent to the check valves under the RH seat deck immediately forward of the rear spar. The loss of heat in the pneumatic plumbing will reduce the temperature of the bleed air from a maximum temperature of 650°F to approximately 70°F above ambient air temperature by the time it reaches the pressure regulator valve. The regulator valve is set at approximately 18 psi of pressure and incorporates a safety valve that will limit pressure to approximately 21 psi by venting




excess pressure overboard. From the pressure regulator valve, lines are routed to the various aircraft systems that utilize pneumatic pressure. A pressure gauge calibrated in pounds per square inch indicates air pressure available to operate the various systems. On airplanes LJ 1 through LJ 1062 the pressure gauge is located on the right side panel next to the vacuum suction gauge. On airplane LJ 1063 and higher the pressure gauges are located on the copilot subpanel.

VACUUM SYSTEM - DESCRIPTION AND OPERATION

Bleed air from the engines is routed through the venture of an ejector to produce the vacuum necessary for operation of the instruments and deflation of the surface deicer boots. The vacuum is regulated by a vacuum regulator valve designed to admit into the system the amount of air required to maintain sufficient vacuum for proper operation of the instruments. The Vacuum Regulator Valve has an air filter attached to protect the vacuum system and the individual instruments on all Kings Airs are equipped with air filters to protect them from contamination. The vacuum system furnishes vacuum to operate the surface deice system, the copilot's gyro instruments, the air-operated turn and slip indicator, the vacuum (gyro suction) gage, and the cabin pressurization control system . The vacuum is produced by an ejector that is operated by the pneumatic system




using bleed air from the engines. To produce the vacuum, pneumatic air is passed through the ejector venturi which draws air from the vacuum system regulator valve, the instrument air filter, the cabin pressure controller and the cabin safety outflow valve. Each of these components has filtered inlets that must be cleaned or replaced at a scheduled time. The vacuum is regulated by a vacuum regulator valve that admits into the system the amount of air required to maintain sufficient vacuum (5.9 in. Hg.) for proper operation of the vacuum-operated systems and components. The surface deicer system uses vacuum to deflate the deicer boots after being inflated by pneumatic pressure.

The cabin pressurization control system uses vacuum to operate the controller and outflow valves. The vacuum ports of the flight instruments are plumbed to a vacuum manifold which is located to the right of the air plane centerline and aft of the pressure bulkhead. The instrument air inlet ports are plumbed to the air intake manifold that is connected to the instrument air filter. The port on the end of each manifold is plumbed to the vacuum (gyro suction) gage. The second port of each manifold is plumbed to the turn and slip indicator. When an electric turn and bank indicator is installed, these ports are capped. The third port of each manifold is plumbed to the directional gyro indicator. The fourth port of each manifold is plumbe d to the gyro horizon indicator.

PILOT TIP The instrument filter is located at the top of the avionics compartment and

should be replaced every 500 hours.

ENGINE BLEED AIR CONTROL Bleed air valve switches are located on the copilot’s left subpanel. When the switches are placed in the OPEN position, environmental bleed air is available to pressurize the airplane. Amber annunciators placarded L BL AIR OFF and R BL AIR OFF will illuminate to indicate that the respective bleed air valve switch is in the CLOSED position. The annunciators indicate only switch position and not the position of the respective bleed air valve.

DOOR SEAL SYSTEM

The cabin entrance door and emergency exit on the C-90A utilize air from the pneumatic system to inflate the door seals after takeoff. Bleed air regulated at 4 psi is tapped from the left engine manifold and directed to the door seal. The bleed air is used to pressurize the landing gear hydraulic system reservoir and the escape hatch seal is pressurized from the pneumatic system manifold




FLIGHT HOUR RECORDER

The Hobbs meter is located on the copilot's rights subpanel. In order for it to operate, pneumatic air must be supplied along with DC power through the flap control circuit breaker. In addition, the right landing gear squat switch must be in the extended position.

PNEUMATIC LIMITATIONS

Pneumatic Gage Green Arc (Normal Operating Range) 12 to 20 psi Red Line (Maximum Operating Limit) 20 psi

Gyro Suction Gage Narrow Green Arc (Normal from 35,000 to 15,000 feet) 2.8 to 4.3 in. Hg Wide Green Arc (Normal from 15,000 feet to Sea Level) 4.3 to 5.9 in. Hg 35K marked on face of gage at 3.0 in. Hg 15K marked on face of gage at 4.3 in. Hg

PNEUMATIC SYSTEM EMERGENCY PROCEDURES

None.

PNEUMATIC SYSTEM ABNORMAL PROCEDURES

None.

PNEUMATIC SYSTEM EXPANDED PROCEDURES

Vacuum/Pneumatic Pressure Check (2000 RPM) 1. Left Bleed-Air Switch OFF 2. Pneumatic/Vacuum Gage PNEU 12-20/VAC 4.3-5.9 psi 3. Right Bleed-Air Switch OFF 4. Pneumatic/Vacuum Gage ZERO 5. Bleed-Air Warning Lights ILLUMINATED 6. Left Bleed-Air Switch OPEN 7. Pneumatic/Vacuum Gage PNEU 12-20/ VAC 4.3-5.9 psi 8. Bleed-Air Warning Lights EXTINGUISH 9. Right Bleed-Air Switch OPEN




PNEUMATIC AND VACUUM SYSTEM QUESTIONS

1. How is the vacuum source created?

2. The cabin pressurization control system uses to operate the controller and outflow valves.

3. Normal gyro suction is psi.

4. Normal pneumatic pressure is psi.

5. The following engine instruments may confirm a bleed air leak:

A. Increase Torque and increase N1 B. Increase in RPM and ITT C. Increase in ITT and decrease in RPM D. Decrease in torque and increase in ITT




Chapter 10

Anti-Icing Systems OBJECTIVES

After completing this unit, the student should be able to:

1. Describe anti-icing systems. 2. Understand conditions requiring the use of anti-icing systems. 3. Explain operation of all anti-icing systems. 4. Describe means of verifying correct operation. 5. Describe use of alternate anti-icing systems.

DESCRIPTION

The ANTI-ICING SYSTEMS section of this workbook presents a description and discussion of the airplane anti -icing systems. All of the anti- ice and deice systems in this airplane are described in detail, showing location, controls, and how they are used. The purpose of this training unit is to acquaint the pilot with all the systems available for flight in icing or heavy rain conditions, and their controls. Procedures in case of malfunction in any system are included. This also includes information concerning preflight deicing and defrosting. Flight in known icing conditions requires knowledge of conditions conducive to icing and of all systems available to prevent excessive ice from forming on the airplane.

ICE AND RAIN PROTECTION - DESCRIPTION AND OPERATION

The airplane is equipped with a variety of ice and rain protection systems that can be utilized during inclement weather conditions.

Airfoil The pneumatic deice boots on the wings and tail remove ice formed during flight. Regulated bleed air pressure and vacuum are cycled to the pneumatic boots for the inflation -deflation cycle. The selector switch that controls the system permits automatic single-cycle operation or manual operation. The deice system is operated with bleed air pressure obtained from the engine compressors. This air is routed through a regulator valve that is set to maintain the pressure required to inflate the deice boots on the leading edge of each wing and the horizontal and vertical stabilizer. To assure operation of the




system should one engine fail, a check valve is incorporated in the bleed line from each engine to prevent the escape of air pressure into the chamber of the inoperative compressor. The bleed air from the engine is also routed through ejectors that employ the venturi effect to produce vacuum for deflation of the deice boots and operation of certain flight instruments. The inflation and deflation phases of operation are controlled by means of distributor valves. The deice system is actuated by a three-way toggle switch on the LH subpanel. This

switch is spring-loaded to return to the OFF position

Wing Boots

from either the MANUAL or SINGLE position. When the switch is pushed to the SINGLE position, one complete cycle of deicer operation automatically follows as the valves open to inflate the deice boots. After an inflation period of approximately seven seconds, a timer switches the valve to the VACUUM, position and deflates the boots. When the switch is pushed to the MANUAL position, the boots will inflate and will stay inflated positions as long as the switch is held in the manual position. Upon release of the switch, the distributor valves return to the VACUUM position and the deice boots remain deflated until the switch is actuated again.

For most effective deicing operation, allow at least 1/2 inch of ice to form before attempting ice removal. Very thin ice may crack and cling to the boots instead of shedding. Maintain a minimum speed of 140 KNOTS during sustained icing conditions to prevent ice accumulation on

unprotected services of the wing. The boots should never be inflated for takeoff or landing.

Tail Boots




NOTE The National Transportation Safety Board states that:

• As little as ¼ inch of leading-edge ice can increase the stall speed 25 to 40 knots.

• Sudden departure from controlled flight is possible with only ¼ inch of leading-edge ice accumulation at normal approach speeds.

• In theory, ice bridging could occur if the expanding boot pushes the ice into a frozen shape around the expanded boot, thus rendering the boot ineffective at removing the ice.

• The Safety Board has no known cases where ice bridging has caused an incident or accident, and has investigated numerous incidents and accidents involving a delayed activation of deice boots.

• Ice bridging is extremely rare, if it exists at all. • Early activation of the deice boots limits the effects of leading-edge

ice and improves the operating safety margin. • Using the autopilot can hide changes in the handling qualities of the

airplane that may be a precursor to premature stall or loss of control.

• Leading-edge deice boots should be activated as soon as icing is encountered, unless the aircraft flight manual or the pilot’s operating handbook specifically directs not to activate them.

• If the aircraft flight manual or the pilot’s operating handbook specifies to wait for an accumulation of ice before activating the deice boots, maintain extremely careful vigilance of airspeed and any unusual handling qualities.

• Be aware that some aircraft manufacturers maintain that waiting for the accumulation of ice is still the most effective means of shedding ice.

Ice inspection lights are mounted on the outside of each engine nacelle and illuminate the leading edge of the wing. They are controlled by a single switch labeled ICE located on the pilot’s right sub-panel.

PILOT TIP The ice lights operate at a very high temperature. Do not operate for

extended periods of time while on the ground.




Deice Boot – Protective Coating Age Master No. 1 and Icex coating are both products of the B.F. Goodrich Company. Age Master No. 1 is a liquid coating that protects rubber products from weathering and ozone and extends the life of the boots. Icex coating is a silicone-based coating specifically compounded to lower the strength of ice adhesion on the surface of the deicer boots. Icex will not damage the rubber boots and offers additional protection from the harmful elements of the atmosphere.

Age Master No.1 Application Age Master No. 1 is a protective coating which chemically bonds with the rubber in the deicer boot and helps resist the deteriorating effects of ozone, sunlight, weather, oxidation and pollution. The coating should be applied as instructed on the label of the container. For continued protection of the boot surface, the coating should be applied every 150 hours. Two treatments per year should be adequate.

Icex Application Icex coating is a silicone-based material that lowers the strength of ice adhesion on the surface of the deicer boots. When properly applied, Icex provides a smooth, polished film that evens out microscopic irregularities on the rubber surface. Ice formations have less chance to cling and are removed faster and cleaner when the boots are operated. Icex should be applied as instructed on the label of the container.

INERTIAL ICE SEPARATION SYSTEM

An inertial ice separation system is installed in each engine air inlet to prevent moisture particles from entering the engine inlet during icing conditions. When icing conditions are encountered, a movable inertial ice vane is lowered into the inlet airstream to induce an abrupt turn in the airflow before entering the engine inlet screen. The heavy ice-laden air is then discharged overboard through an opening in the lower cowling at the aft end of the air duct. The inertial ice vanes are extended and retracted by switches located on the pilot’s left subpanel. The switches are placarded ENGIN ANTI-ICE – LEFT – RIGHT – ON – OFF – ACTUATORS – STANDBY – MAIN. Vane position during operation is indicated by a slight decrease in torque with switches ON. In addition, the actuators have dual motors to provide a redundant system.




The ACTUATORS switch allows the selection of either MAIN or STANDBY actuator motor. The vanes have only two positions, there are no intermediate positions. The system is monitored by L and R ENG ANTI-ICE (green) and L and R ENG ICE FAIL (amber) annunciators. Illumination of the L or R ENG ANTI-ICE (green) annunciator indicates that the system is actuated. Illumination of the L or R ENG ICE FAIL (amber) annunciator indicates that the system did not operate to the desired position. Immediate illumination of the L or R ENG ICE FAIL (amber) annunciator indicates loss of electrical power, whereas delayed illumination indicates an inoperative actuator.

PILOT TIP

Icing conditions occur even though you are not getting surface ice. When in visible moisture at temperatures of +5ºC or colder, extend the ice

vanes. The engine ice vanes should be extended for all ground operations to help prevent FOD. Always maintain oil temperature within limits.

AIR INTAKE ANTI-ICE LIP

The lip around each air intake leading edge is heated by engine exhaust to prevent the formation of ice during inclement weather. This system is completely automatic and requires no pilot action.

ENGINE FUEL CONTROL HEAT

The compressor bleed air line to each engine fuel control unit is protected against icing by electrically heated jackets. Cams on the CONDITION levers activate switches that control the electric power to the air line heaters. Fuel control heat is “ON” for all flight operations when the condition levers are moved out of the fuel cutoff range.

WINDOWS AND WINDSHIELDS

Electrical heating elements embedded in the wind shield provide adequate protection against the formation of ice while air from the cabin heating systems prevents fogging to ensure visibility during operation under icing conditions. Normally a temperature of 90°F - 110°F is maintained.




Switches in the ICE PROTECTION group on the pilot’s right subpanel, placarded, WSHLD ANTI-ICE – NORMAL – OFF – HI – PILOT – COPILOT, are used to control windshield heat. When the switches, PILOT’S and COPILOT’S, are in the NORMAL (up) position, the secondary areas of the windshields are heated. When the switches are in the HI (down) position, the primary areas are heated. The primary areas are smaller areas and are heated to higher temperatures. Each switch must be lifted over a detent before it can be moved to the HI position. This lever-lock feature prevents inadvertent selection of the HI position when moving the switches from the NORMAL to the OFF (center) position. The electrically heated laminated glass and plastic windshield is subject to gradual process of delamination due to the effect of chemical action and differentials of temperature and pressure incurred during pressurized flights at varying altitudes and under varying weather conditions. This delamination is not detrimental to the structural integrity of the windshield, but it may significantly decrease visibility or the deicing capability of the windshield. Beyond certain limits, either of these effects will require the replacement of the windshield.

PILOT TIP

Erratic operation of the magnetic compass may occur while windshield heat is being used. Objects viewed through the windshield will be distorted when the

windshield heat is on.

Cleaning Plastic Windows The plastic windows should be kept clean and waxed at all times. Only approved plexiglass cleaners such as Maguiar’s Mirror Glaze, Permatex Plastic Cleaner or Parko Antistatic Plastic Polish should be utilized. To prevent scratches and crazing, wash the windows carefully with plenty of soap and water, using the palm of the hand to feel and dislodge dirt and mud. A soft cloth, chamois, or sponge may be used, but only to carry water to the window surface. Rinse the window thoroughly, then dry it with a clean, moist chamois. Rubbing the surface of the plastic window with a dry cloth will serve only to build up an electrostatic charge that attracts dust.

TIP

It is equally essential that the windshield wipers be thoroughly cleaned, for grit trapped by the wipers is a common source of scratches in the windshield

when the wipers are operated.

PROPELLER DEICING

The propellers are protected against icing by electrothermal boots that automatically cycle to prevent the formation of ice on each blade. The propeller electric deice system includes: an electrically heated boot for each propeller blade, a timer, an on-off switch and an ammeter. When the switch is turned on the ammeter registers 14 to 18 amperes of current to the prop boots. The




current flows from the timer through the brush assemblies to the slip rings, where it is distributed to the individual propeller deicer boots.

Heat produced by the heating elements in the deicer boots reduces the adhesion of the ice. The ice is then removed by the centrifugal effect of the propeller and the blast of the airstream. Power to the deice boot heating elements is cycled in a continuous programmed sequence.

Power to these deice boots is cycled in 90 -second phases. The first 90-second phase heats all the deicer boots on the RH propeller. The second phase heats all the deicer boots on the LH propeller. The deicer timer completes one full cycle every three minutes. As the deicer timer moves from one phase to the next, a momentary deflection of the propeller ammeter needle may be noted.

NOTE

The heating sequences for the deicer boots noted in the previous section are for normal operation. However, since the timer does not return to any given point when the power is turned off, it may restart at any sequence point.

With the propeller heat switch on, the prop amp gauge located on the pilot’s left subpanel, should indicate current flow. Normal current flow is indicated by green arc showing between 14 to 18 amps. If the current rises beyond 20 amps, the system should be turned off. Loss of one heating element does not mean entire system must be turned off, although ice may build up on one propeller.

PILOT TIP

Operating the propeller heat with the engines off will damage the heating elements .




PITOT HEAT

A heating element in the pitot mast prevents the pitot opening from becoming clogged with ice. The heating element is controlled by a switch placarded PITOT, LEFT and RIGHT located on the left inboard subpanel. It is not recommended to operate the pitot heat while on the ground except to test the system or to remove ice and snow from the mast.

STALL WARNING VANE HEAT

The lift transducer is equipped with anti-icing capability on both the mounting plate and the vane. The heat is controlled by a switch in the ice group located on the pilot's right sub-panel identified: STALL WARN. The level of heat is minimal for ground operation, but is automatically increased for flight operation through the left landing gear safety switch.

PILOT TIP

Prolonged use of the stall warning and pitot heat on the ground will damage the heating elements.

WARNING! The heating elements protect the lift transducer vane and face plate from ice. However, a buildup of ice on the wing may change or disrupt the airflow and prevent the system from accurately indicating an imminent stall. Remember that the stall speed increases whenever ice accumulates on any airplane.

FUEL VENTS

The fuel system is vented through a recessed vent coupled to a static vent on the underside of the wing adjacent to the nacelle. One vent (NACA) is recessed to prevent icing. The second vent is heated to prevent icing and serves as a backup should the NACA vent become plugged.

FUEL HEAT

An oil-to-fuel heat exchanger, located on the engine accessory case, operates continuously and automatically to heat the fuel sufficiently to prevent ice from collecting in the fuel control unit. Fuel heat is automatic and requires no action by the pilot.




ICING LIMITATIONS

Minimum Ambient Temperature for Operation of Deicing Boots -40°C

Minimum Airspeed for Sustained Icing Flight -140 Knots

Maximum Airspeed with Windshield Icing – 226 Knots

Sustained flight in icing conditions with flaps extended is prohibited except for approach and landings.

ICE VANES, LEFT and RIGHT, shall be extended for operations in ambient temperatures of +5°C or below when flight free of visible moisture cannot be assured.

ICE VANES, LEFT and RIGHT, shall be retracted for all takeoff and flight operations in ambient temperatures of above +15°C.

EMERGENCY ICING SYSTEM PROCEDURES

None. ABNORMAL ICING SYSTEM PROCEDURES

Electrothermal Propeller Deice Abnormal Readings on Deice Ammeter. (Normal Operation: 14 to 18 amps) 1. Zero Amps:

A. a Prop Deice - CHECK ON B. If OFF, reposition to ON after 30 seconds. C. If in ON position with zero amps reading, system is inoperative: position

the switch to OFF.

2. Below 18 amps: A. Continue operation. B. If propeller imbalance occurs, increase rpm briefly to aid in ice removal.

3. Over 24-28 amps:

A. Continue operation. B. If propeller imbalance occurs, increase rpm briefly to aid in ice removal.

4. Above 28 Amps: A. Avoid icing conditions, since continued operation of the system cannot be assured. B. Do not operate the system, except in emergencies. C. Restrict time of operation to a minimum.




ICING EXPANDED PROCEDURES

1. Power Levers 1,800 RPM 2. Ice Vane SWITCH 3. Torque Drop CHECKED 4. Ice Vane SWITCH

WARNING Either the MAIN or STANDBY actuator must be operational on each engine before takeoff.

1. Engine Anti-ice Actuators – STANDBY 2. Engine Anti-ice – OFF [L ENG ANTI-ICE] & [R ENG ANTI-ICE] –

EXTINQUISHED 3. Engine Anti-ice Actuators – MAIN 4. Engine Anti-ice – ON [L ENG ANTI-ICE] & [R ENG ANTI-ICE] –

ILLUMINATED Windshield Anti-ice, Pilot’s and Copilot’s – CHECK INDIVIDUALLY

1. Windshield Anti-ice – HI (observe increase on left and right loadmeters) 2. Windshield Anti-ice – OFF, THEN NORMAL (observe increase on left and right

loadmeters) 3. Windshield Anti-ice – OFF

Electrothermal Propeller Deice – CHECK

CAUTION DO NOT OPERATE PROPELLER DEICE WHEN THE PROPELLERS ARE STATIC.

1. Prop Deice – ON 2. Deice Ammeter – 18 to 24 AMPS (monitor for 90 seconds to ensure automatice

timer operation) 3. Prop Deice – OFF

Surface Deice System – CHECK

1. Pneumatic Pressure – GREEN ARC (12-20 PSI) 2. Surface Deice Switch – SINGLEL AND RELEASE

a. Pneumatic Pressure Gage – WILL DECREASE MOMENTARILY b. Boots – CHECK BOTH WING AND BOTH HORIZONTAL STABILIZER

BOOTS VISUALLY, IF POSSIBLE, FOR INFLATION AND VACUUM HOLD DOWN

c. Wing Boots will inflate in approximately 6 seconds, followed by horizontal stabilizer boots.

3. Surface Deice Switch – MANUAL AND HOLD a. Pneumatic Pressure Gage – WILL DECREASE MOMENTARILY b. Boots – CHECK BOTH WING AND BOTH HORIZONTAL STABILIZER

BOOS VISUALLY, IF POSSIBLE, FOR INFLATION




4. Surface Deice Switch – RELEASE a. Boots – CHECK BOTH WING AND BOTH HORIZONTAL STABILIZER

BOOTS VISUALLY, IF POSSIBLE, FOR INFLATION AND VACUUM HOLD DOWN

Pitot Heat – CHECK (observe slight increase in loadmeter)

ENCOUNTERING ICING CONDITIONS

NOTE For 60 years, pilots have been taught to wait for a prescribed accumulation of leading-edge ice before activating the deice boots because of the believed threat of ice bridging

WARNING Due to distortion of the wing airfoil, ice accumulations on the leading edges can cause a significant loss in rate of climb and in speed performance, as well as increases in stall speed. Even after cycling the deicing boots, the ice accumulation remaining on the boots and unprotected areas of the airplane can cause large performance losses. For the same reason, the aural stall warning system may not be accurate and should not be relied upon. Maintain a comfortable margin of airspeed above the normal stall airspeed. In order to minimize ice accumulation on unprotected surfaces of the wing, maintain a minimum of 140 knots during operations in sustained icing conditions. Prior to a landing approach, cycle the deicing boots to shed any accumulated ice.

1. Engine Ice Protection Before visible moisture is encountered at +5°C and below, or: At night when freedom from visible moisture is not assured at +5°C and below. (Operation of strobe lights will sometimes show ice crystals not normally visible.)

a. Engine Anti-ice – ON [L ENG ANTI-ICE] & [R ENG ANTI-ICE] – ILLUMINATED

b. Engine Instruments – DROP IN TORQUE AND INCREASE IN ITT INDICATES PROPER OPERATION

NOTE For Illumination of the L and/or R [ENG ICE FAIL] indicates a failure of the selected Engine Anti-ice System. Immediate illumination indicates loss of power to the actuator(s). Select the other actuator(s).

c. Power – RESET, IF DESIRED (observe engine limitations)




WARNING If in doubt, actuate the Engine Anti-ice System. Engine icing can occur even though no surface icing is present. If freedom from visible moisture cannot be assured, engine ice protection should be activated. Visible moisture is moisture in any form; clouds, ice crystals, snow, rain, sleet, hail or any combination of these.

2. Auto Ignition – ARM

NOTE Engine Auto Ignition must be armed for icing flight, precipitation, and operation during turbulence. To prevent prolonged operation of the ignitors with the system armed, do not reduce power levers below 425 ft-lbs torque.

3. Electrothermal Prop Deice – ON a. The system may be operated continuously in flight, and will function

automatically until the switch is turned off. b. Prop RPM – MODULATE BRIEFLY TO RELIEVE PROPELLER

IMBALANCE DUE TO ICE. REPEAT AS NECESSARY.

4. Surface Deice WARNING All components of the surface deice system must be monitored during icing flight to ensure the system is functioning normally. These components include: Pneumatic Pressure Gage. The gage should indicate 12-20 psi before boots are activated. The pressure will momentarily decrease when the boots are activated. Gyro Suction Gage. The gage should indicate in the area of the green arc corresponding to the airplane altitude. The vacuum will momentarily decrease when the boots are activated.

Pneumatic Boots. Visually monitor the boots, where possible, to ensure ice is being removed. CAUTION Operation of the surface deice system in ambient temperatures below -40°C can cause permanent damage to the deice boots.

When Ice Accumulates to ½ to 1 inch:

a. Surface Deice Switch – SINGLE and RELEASE b. Repeat as required.




If Single Position of the Surface Deice Switch Fails: c. Surface Deice Switch – MANUAL AND HOLD FOR A MINIMUM OF 6

SECONDS, THEN RELEASE d. Repeat as required.

5. Windshield Anti-Ice – NORMAL/HI

NOTE To ensure adequate windshield anti-icing protection, operation in icing conditions at or below ambient temperatures of -24°C is not recommended. In the event of windshield icing, reduce airspeed as required.

PILOT TIP

Turn off or limit the use of the auto pilot in order to better “feel” changes in the handling qualities of the airplane.

ANTI-ICING SYSTEM QUESTIONS

1. Windshield heat : A. Affects the compass B. Is used all the time C. Is prohibited when outside air temperature is 30ºF or colder D. Will shattered a cold soaked windshield.

2. Use the inertial separators whenever the temperature is and

is present.

3. T or F: Use of flaps in icing condition is prohibited.

4. Minimum speed for flight in icing conditions is K.

5. T or F: The wing and tail boots sequence at the same time in the CYCLE

6. The engine inlet lips are heated by: A. Bleed air from the P3 section of the engine B. Exhaust gases C. Electrothermal boots D. NACA design prevents icing of the inlets.

7. The deice boots should not be cycled if the outside air temperature is below:

A. -50ºC B. -40ºC C. -40ºF D. -30ºC




8. T or F: Continuous use of the pitot on the ground is recommended:

9. If the boots are manually inflated for more than 10 seconds: A. The boots may develop rips and tears B. The boots will automatically deflate C. Ice may form on the expanded boot and not be removable D. Add drag to the wing

10. Define icing conditions.

11. Should the inertial separators ever be used on the ground?

12. Describe the working principle of the inertial separators ("ice vanes").

13. How would you know if the inertial separators have actually lowered?

14. T or F: Damage will occur if windshield heat is used on the ground?

15. What caution should be considered regarding the use windshield heat?

16. Under what conditions could the stall warning-system be inaccurate?

17. How can you determine that the propeller deice timer is working correctly?




Chapter 11

Flight Controls OBJECTIVES

After completion of this section of the workbook, the student should be able to:

1. Explain the operation of the primary flight controls. 2. Describe the location and operation of the trim tabs and controls. 3. Explain the use of the control locks. 4. Explain the operation of the flaps. 5. Describe the stall warning system.

FLIGHT CONTROLS

Dual controls are provided for the pilot and copilot. The ailerons and elevators are operated by conventional push-pull control yokes interconnected by a T- column. The flight controls are cable-operated conventional surfaces which require no power assistance for normal control by the pilot or copilot. All primary flight control surfaces are manually controlled through cable and bellcrank systems. Each system incorporates surface travel stops and linkage adjustments. The rudder pedals are interconnected by a linkage below the cockpit floor. The rudder pedal bellcranks are adjustable to two positions. The ailerons, elevators and rudder may be secured with control locks in the cockpit.

PILOT TIP Do not push or pull the aircraft by the propellers or control surfaces




ELEVATOR TRIM

Manual control of the elevator trim is accomplished by utilizing a trim wheel located on the left side of the throttle pedestal. The electric elevator-trim system is controlled by an Elevator - On - Off switch located on the pedestal. It incorporates a dual-element thumb switch on each control wheel, a trim- disconnect switch on each control wheel, and a Pitch Trim circuit breaker on the right side panel. The Elevator Trim switch must be on for the system to operate. Both elements of either dual- element thumb switch must be simultaneously pushed forward to achieve nose-down trim and moved aft for nose-up trim. When the trim switch is released, it returns to the center (Off) position. Any activation of the trim system by the copilot's trim switch can be overridden by the pilot's trim switch. Pedestal Trim Switch A before take-off check of both dual element thumb switches should be made by

moving each of the four switch elements individually. One switch element should not activate the system. A two level, push-button, momentary-on, trim- disconnect switch is located inboard of the trim switch on the outboard grip of each control wheel. The electric elevator-trim system can be disconnected by depressing either of these switches.

If the autopilot is engaged, depressing either trim-disconnect switch to the first of the two levels disconnects the autopilot and the yaw damp system.




Depressing the switch to the second level disconnects the autopilot, the yaw damp system, and the electric elevator-trim system. A green annunciator on the caution/advisory annunciator panel placarded ELEC TRIM OFF, alerts the pilot whenever the system has been disabled with a trim-disconnect switch and the Elevator Trim switch is on. The system can be reset by recycling the Elevator Trim switch on the pedestal. The manual-trim control wheel can be used to change the trim anytime, whether or not the electric trim system is in the operative mode.

PILOT TIP

Do not allow the trim system to move pass the limits on the elevator trim indicator either manually, electrically or by the autopilot.

CONTROL LOCKS

The control locks are provided to prevent movement of the controls while the airplane is parked. The control lock consists of a U-shaped clamp and two pins connected by a chain. The pins lock the primary flight controls and the U- shaped clamp fits around the engine power control levers and serves to warn the pilot not to start the engine with the control locks insta lled. It is important that the locks be installed or removed together to preclude the possibility of an attempt to taxi or fly the airplane with the power levers released and the pins still installed in the flight controls.

GROUND MOORING/TOWING

Three tie-down eyes are provided, one on each wing and another on the tail. To secure the airplane, chock all the wheels fore and aft and tie the airplane down utilizing all three tie-down points.

CAUTION REMOVE THE CONTROL LOCKS BEFORE TOWING THE AIRPLANE. IF TOWED WITH A TUG WHILE RUDDER LOCK IS IN PLACE, SERIOUS DAMAGE TO THE STEERING LINKAGE MAY OCCUR.

With the tow bar is connected to the nose strut, the airplane can be steered with the nose wheel when moving it by hand or with a tug. When moving the airplane, do not push on the propeller or control surfaces.




CAUTION NEVER TOW OR TAXI THE AIRPLANE WITH A FLAT STRUT. EVEN BRIEF TOWING OR TAXING IN THIS CONDITION WILL RESULT IN SEVERE DAMAGE. NEVER EXCEED THE TURNING LIMITS MARKED ON THE NOSE GEAR STRUT DURING GROUND HANDLING. IF THE TURN LIMITATION IS EXCEEDED DURING GROUND HANDLING, DAMAGE TO THE STEERING LINKAGE AND NOSE STRUT WILL OCCUR.

WING FLAPS

The C-90A‘s operational speed limit for flaps provides for easy traffic pattern transition. Flaps are selectable to 3 positions: up, approach (35%), and down (100%). The airplane’s flap tracks are not exposed when flaps are retracted. This design eliminates exposed surfaces that could collect ice and potentially interfere with flap operation. The flaps, two panels on each wing, are driven by an electric motor through a gearbox mounted on the forward side of the rear spar. The motor incorporates a dynamic braking system which helps to prevent overtravel of the flaps. The gearbox drives four flexible drive shafts connected to a jackscrew actuator at each flap. The flaps are operated by a sliding switch lever located just below the condition levers. Flap travel, from 0% to 100% (fully down) is registered in percentage on an electric flap indicator at the top of the pedestal forward of the power levers. The indicator is operated by a potentiometer driven by the right inboard flap. Any of the three flap positions, UP, APPROACH or DOWN may be selected by moving the flap selector lever up or down to the selected switch position indicated on the pedestal. A side detent provides for quick selection of the APPROACH position (35% flaps). From the UP position to the APPROACH position, the flaps cannot be stopped at an intermediate point. Between the APPROACH position and DOWN, the flaps may be stopped as desired by moving the handle to the DOWN position until the flaps have moved to the desired position, then moving the flap handle back to APPROACH. The flaps may also be raised to any position between DOW N and APPROACH by raising the handle to UP until the desired setting is reached, then returning the handle to APPROACH. The APPROACH detent acts as a stop for any position greater than 35%. Moving the flap handle out of the UP position renders the landing gear warning horn silence function inoperative. With the flap handle out of the UP position, the landing gear warning horn can be silenced only by lowering the landing gear or advancing the power levers. A second approach position switch will cause the warning horn to sound continuously when the flaps are lowered beyond the approach position until the landing gear is extended, regardless of the power lever setting. On later models, three detents provide for quick selection of UP, APPROACH, and DOWN positions. The flaps cannot be stopped in an intermediate position. The flap motor power circuit is protected by a 20-ampere flap motor circuit breaker placarded FLAP MOTOR, located on the right circuit breaker panel. A




5-ampere circuit breaker (FLAP IND & CONTROL) for the control breaker is located on the right circuit breaker panel. Lowering the flaps will produce these results:

• Attitude – Nose Up • Airspeed – Reduced • Stall Speed – Lowered • Trim – Nose-down Adjustment Required to Maintain

YAW DAMPER

The Yaw Damp system is designed to provide the pilot with help in maintaining directional control and increase ride comfort. The system is normally incorporated in the autopilot, but can be operated separately. The yaw damper must be disengaged during takeoff and landing. Operating instruction can be found in Flight Manual Supplement. RUDDER BOOST The King Air C90A Series airplanes are equipped with a pneumatic type rudder boost system. This system, when engaged, aids the pilot in maintaining directional stability should engine failure or a large power variation between engines occur. The system senses the bleed air pressure of both engines at the differential pressure switch, and should a power variation occur, a shuttle in the switch will move towards the low side. This actuates the corresponding solenoid valve, which opens and allows bleed air to travel to its servo. The servo, in turn, moves the attaching cable that moves the rudder cable and the rudder in the direction needed to stabilize the airplane. The main components of this system are the differential pressure switch, inline filter, rudder boost pressure regulator, left and right solenoid valves, left and right pressure relief valves, left and right servos and the associated plumbing. Each servo is attached to the primary rudder cable by a cable and clamp. Slack in the rudder and rudder servo cables is eliminated by tension from a spring assembly on each servo cable. The system is energized by a two-position toggle switch on the pedestal placarded RUDDER BOOST OFF. The circuit is protected by the five-ampere RUDDER BOOST circuit breaker in the Circuit Breaker Panel.

PILOT TIP

A buildup of ice on the wing may change or disrupt the airflow and prevent the system from accurately indicating an imminent stall.




STALL WARNING SYSTEM

The stall warning system senses angle of attack through a transducer vane mounted on the leading edge of the left wing. When the lift transducer vane determines that a stall is imminent, the switch completes a circuit to a warning horn and illuminates a red warning light. The horn sounds and the light illuminates at 7 to 9 knots about the stall. The system has a heater that can be selected by the pilot prior to entering icing conditions.

FLIGHT CONTROL LIMITATIONS

Maneuver Limits The BEECHCRAFT King Air C-90A’s are Normal Category Airplanes. Acrobatic maneuvers, including spins, are prohibited.

FLIGHT LOAD FACTOR LIMITS

FLAPS UP FLAPS DOWN

3.32 positive g's 2.00 positive g's 1.33 negative g's 3.29 positive g’s at 10,100 lbs.

FLIGHT CONTROL EMERGENCY PROCEDURES


Flight Controls Unscheduled Electric Pitch Trim Activation

1. Airplane Attitude – MAINTAIN USING ELEVATOR CONTROL 2. AP/Trim Discount – DEPRESS FULLY & HOLD [TRIM] - ILLUMINATES

NOTE Autopilot will disengage when the disconnect switch is depressed.

3. Manually retrim airplane. 4. AP/Trim Disconnect – DEPRESS FULLY & HOLD [TRIM] – ILLUMINATED




If Trim Continues to Run: 5. AP/Trim Disconnect – DEPRESS FULLY & HOLD [TRIM] – ILLUMINATED 6. Pitch Trim Circuit Breaker – PULL 7. AP/Trim Disconnect – RELEASE 8. Manually retrim airplane. 9. Autopilot – DO NOT ENGAGE

Unscheduled Rudder Boost Activation Rudder boost operation without a large variation of power between the engines indicates a failure of the system.

1. Directional Control – MAINTAIN USING RUDDER PEDALS 2. Rudder Boost – OFF

If Condition Persists:

3. Rudder Boost Circuit Breaker – PULL 4. Either Bleed Air Valve – PNEU & ENVIR OFF 5. Rudder Trim – AS REQUIRED 6. Perform normal landing.

CAUTION DO NOT reactivate electric trim system until cause of malfunction has been determined.

Spins If the spin is entered inadvertently:

Immediately move the control column full forward, apply full rudder opposite to the direction of the spin, and reduce power on both engines to idle. These three actions should be done as nearly simultaneously as possible; then continue to hold this control position until rotation stops and then neutralize all controls and execute a smooth pullout. Ailerons should be neutral during recovery.

FLIGHT CONTROL ABNORMAL PROCEDURES

Flaps Up Landing Refer to the POH PERFORMANCE Section, for Flaps Up Landing Distance and Approach Speed. 1. Approach Speed - CONFIRM 2. Autofeather (if installed) - ARM 3. Pressurization - CHECK 4. Cabin Sign - NO SMOKE & FSB 5. Flaps – UP




CAUTION Do not silence the landing gear warning horn, since the flap actuated portion of

the landing gear warning system will not be actuated during a flap up landing.

6. Landing Gear – DN 7. Lights – REQUIRED NOTE 8. Under low visibility conditions, landing and taxi lights should be left off due to

light reflections. 9. Radar - AS REQUIRED 10. Surface Deice - CYCLE (as required)

NOTE

If crosswind landing is anticipated, determine Crosswind Component from the PERFORMANCE section of the POH. Immediately prior to touchdown, lower

upwind wing and align the fuselage with the runway. During rollout, hold aileron control into the wind and maintain directional control with rudder and

brakes. Use propeller reverse as desired.

When Landing Assured: 11. Approach Speed - ESTABLISHED 12. Yaw Damp - OFF 13. Propeller Levers - FULL FORWARD 14. Power Levers – IDLE

After Touchdown:

15. Power Levers - LIFT AND SELECT REVERSE 16. Brakes - AS REQUIRED

FLIGHT CONTROL EXPANDED PROCEDURES

Electric Elevator Trim 1. Verify that the ELEV TRIM switch is on. 2. Check operation of the dual-element thumb switches.

WARNING! Operation of the electric trim system should occur only by movement of pairs of switches. Any movement of the elevator trim wheel while actuating only one switch denotes a system malfunction. If a malfunction of the electric trim system is indicated, electric trim must be disengaged and trim changes made with manual trim only.




FLIGHT CONTROLS QUESTIONS

1. List the maximum flap air speeds: Approach Flaps Full Flaps

2. Explain how to select 60% flaps.

3. In what speed range could you not select intermediate flaps?

4. Where is the circuit breaker located for the flap motor?

5. Refer to the emergency procedures. List the procedures for the no flap landing.

6. Is any one of the four flap segments different than the others?

7. Where is the aileron trim tab located?

8. Where is the electric trim switch located?

9. T or F: The flaps have no asymmetrical protection.

10. T or F: The yaw damper must be operational for flight.

11. The wing flaps are: A. Electric B. Hydraulic C. Electrically actuated/hydraulically operated.

12. T or F: The King Air C-90A is a Normal Category airplane. 13. T or F: Rudder Boost is hydraulic powered and electrically activated. 14. T or F: The pilot’s trim switch over-rides the copilot’s trim switch.




Chapter 12

Pitot Static System OBJECTIVES

After completing this section of the work book, the student will be able to:

1. Identify the major components of the pitot static system. 2. Describe how the pilot and copilot instruments receive pitot and static

pressure. 3. Be able to drain the pitot static system. 4. Describe the alternate static source.

PITOT AND STATIC PRESSURE SYSTEM

The pitot and static pressure system provides a source of impact pressure and static air for operation of selected flight instruments. The pitot portion of the system is comprised of the pitot mast mounted on each lower side of the nose, the wiring connecting the heating element of the mast into the electrical system and the tubing between the mast and the airspeed indicators. The impact pressure entering the masts is transmitted to the dual airspeed indicators mounted on the instrument panel through separate tubing routed along each upper side of the nose compartment. Since the pitot mast is the lowest point in each line from the airspeed indicators, the resultant natural drainage eliminates the need for drain valves. Two circuit breaker switches on the left inboard subpanel control the heating elements that prevent the pitot openings in the mast from becoming clogged with ice. The static portion of the system includes two static ports on each side of the fuselage aft of the aft pressure bulkhead. Lines connect the static ports to the instruments in the crew compartment and an alternate line supplies static air for the pilot's instruments should the fuselage static ports become obstructed. The static lines are routed from the static ports to the top center of the fuselage and immediately over to the right side of the fuselage. They are then routed forward along the fuselage beneath the windows to the rate-of-climb indicator, altimeter and airspeed indicator at the instrument panel. The static line drain valves are located behind the access door located in the lower right flight compartment wall adjacent to the instrument panel. The static lines should be drained any time the aircraft has been exposed to rain, either on the ground or during flight. Should abnormal or erratic instrument readings indicate that the normal static source is restricted; the alternate air source may be utilized. This alternate system supplies static air from the interior of the aft fuselage. The alternate static air line is routed through the aft pressure bulkhead forward along the right




side of the fuselage to the static air selector valve. This selector valve is located below the copilot's circuit breaker panel adjacent to the instrument panel. The static air selector valve is held in the normal position by a clip. The alternate air source is selected by raising the clip and moving the toggle from NORMAL to ALTERNATE. The pilot's instruments then function on the alternate air source.

OUTSIDE AIR TEMPERATURE

The outside air temperature indicator is installed in the pilot's overhead panel or the pilot's left sidewall panel. The indicator dial is on the inside of the compartment with the stem of the instrument protruding through the skin of the airplane. The instrument is hermetically sealed against dust and moisture. The instrument consists of a bimetal element which is attached to the staff and pointer. A hollow stainless steel stem encloses the element. A sunshield is installed over the stem for protection.

PITOT STATIC LIMITATIONS None.




PITOT STATIC SYSTEM EMERGENCY PROCEDURES None.

PITOT STATIC SYSTEM ABNORMAL PROCEDURES

Pilot’s Alternate Static Air Source THE PILOT'S ALTERNATE STATIC AIR SOURCE SHOULD BE USED FOR CONDITIONS WHERE THE NORMAL STATIC SOURCE HAS BEEN OBSTRUCTED. When the airplane has been exposed to moisture and/or icing conditions (especially on the ground), the possibility of obstructed static ports should be considered. Partial obstructions will result in the rate of climb indication being sluggish during a climb or descent. Verification of suspected obstruction is possible by switching to the alternate system and noting a sudden sustained change in rate of climb. This may be accompanied by abnormal indicated airspeed and altitude changes beyond normal calibrated differences.

Whenever any obstruction exists in the Normal Static Air System, or when the Alternate Static Air System is desired for use:

A. Pilot's Static Air Source (right side panel) - ALTERNATE B. For Airspeed Calibration and Altimeter Correction, refer to the

PERFORMANCE section of the POH.

NOTE Be certain the static air valve is in the NORMAL position when the alternate

system is not needed.

PITOT STATIC SYSTEM QUESTIONS

1. What are the restrictions against the use of pilot heat?

2. Describe how L & R pitot masts provide separate pitot pressure to pilot and co-pilot airspeed indicators.

3. Where is the location of the emergency (alternate) static source?

4. Does this source provide alternate static pressure to pilot and co-pilot or pilot only?




5. When should the static air line drain petcocks be drained? Why?

6. Why would you not drain them in normal flight after leaving a heavy rainstorm?




Chapter 13

Oxygen System OBJECTIVES

With the use of this training manual the pilot will be able to:

1. Identify the major components which make up the oxygen system. 2. Explain the emergency procedures regarding the use of oxygen. 3. Be familiar with the time of useful consciousness at varying altitudes.

OXYGEN SYSTEM - DESCRIPTION AND OPERATION

Oxygen for flight at high altitudes is supplied by a cylinder mounted behind the aft cabin bulkhead. The cylinder is filled by a valve accessible through an access door located on the right side of the fuselage. The system has two pressure gauges. One is located on the right side panel in the cockpit for in-flight use and the other is located by the filler valve. A push/pull handle (PULL ON), located aft of the overhead light. This handle operates a cable which opens and closes the shut-off valve located at the oxygen supply bottle in the aft, unpressurized area of the fuselage. When this handle is pushed in, no oxygen supply is available anywhere in the airplane. It should be pulled out prior to engine starting to ensure that oxygen will be immediately available anytime it is needed. When this handle is pulled out, the primary oxygen supply line is charged with oxygen, provided the oxygen supply bottle is not empty (Check the oxygen supply pressure gage on the right subpanel and verify that sufficient oxygen is available for the flight). The oxygen supply line delivers oxygen to the two crew oxygen outlets in the cockpit and to the passenger oxygen system. This system provides a constant oxygen flow and is adequate up to a cabin altitude of 30,000 feet. The pilot's oxygen masks are normally stowed behind the pilot’s and copilot’s head. The oxygen outlets are located on the forward cockpit sidewalls. The passenger’s oxygen masks are located in the seatback pockets. All masks are connected to the oxygen system by pushing the plug into the oxygen outlet firmly and turning clockwise approximately 1/4 turn. The passenger oxygen outlets are located in the ceiling at the forward and aft and of the cabin.




Oxygen Cylinders The oxygen system uses steel oxygen cylinders that are available in four sizes. The standard system utilizes the 22-cubic-foot cylinder and some optional systems use the 49 or 66 cubic-foot cylinder. The oxygen cylinder should be filled to a pressure of 1800 psi at a temperature of 70°F. To prevent overheating the oxygen system should be filled slowly. Oxygen cylinders used in the airplane are of two types. Light weight cylinders, stamped "3HT" on the plate on the side, must be hydrostatically tested every three years and the test date stamped on the cylinder. This bottle has a service life of 4,380 pressurizations or 24 years, whichever occurs first, and then must be discarded. Regular weight cylinders, stamped "3A" or "3AA", must be hydrostatically tested every five years and stamped with the retest date. Service life on these cylinders is not limited.

PILOT TIP Offensive odors may be removed from the oxygen system by purging. This should be accomplished anytime the system pressure drops below 50psi.

OXYGEN LIMITATIONS

Filling the Oxygen System When filling the oxygen system, only use Aviator's Breathing Oxygen, MIL-0- 27210.

WARNING! DO NOT USE MEDICAL or INDUSTRIAL OXYGEN. It contains moisture which

can cause the oxygen valve to freeze.

OXYGEN EMERGENCY PROCEDURES





Use of Oxygen

WARNING! The following table sets forth the average time of useful consciousness (TUC)

(time from onset of hypoxia until loss of effective performance) at various altitudes. Cabin Pressure Altitude TUC

35,000 feet 1/2 - 1 minute 30,000 feet 1 - 2 minutes

25,000 feet 3 to 5 minutes 22,000 feet 5 to 10 minutes

12 - 18,000 feet 30 minutes or more

1. Oxygen Control Handle - PULL ON (verify) 2. Crew - DON MASKS 3. Passengers - PLUG IN- DON MASKS 4. Oxygen Duration - CONFIRM

OXYGEN ABNORMAL PROCEDURES None.

OXYGEN SYSTEM QUESTIONS

1 . Why is it unnecessary to remove the oxygen filler valve access plate to check oxygen system pressure?

2 . What is the normal system pressure for a full bottle?

3 . List some precautions to observe during oxygen filling.

4 . Assuming a well-maintained oxygen system, what must the crew do to obtain oxygen? What must passengers do to obtain oxygen?

5 . What is the average TUC at 25,000 feet?




6 . It is acceptable to use medical oxygen if aviator’s breathing oxygen is not available. True False (Choose one)

7 . At 10,000 feet cabin altitude the passenger masks drop automatically. True False (Choose one)

8 . Where is the oxygen refill valve? ____________________________ .

9 . Where is the oxygen bottle? ______________________________.




King Air C-90A/B Profiles and Power Settings


Circle Approach and Landing

NOTE: This is a category B Aircraft, but airspeeds of 121 through 140 KIAS require using category C minimums.

1 ARRIVAL

1. Plan Circling Maneuver 2. Follow Normal Approach

Procedures to MDA

2

MDA

MAP

THRESHOLD 1. Gear—Re-check Down 2. Airspeed— VREF

3. Power—Idle

6

FINAL

MDA (minimum descent altitude)

1. Level off at MDA at least 1 mile prior to MAP if possible.

2. Torque — 600 - 1,000 lbs 3. 120-130 KIAS 4. Maneuver within visibility

criteria. 5. Maintain MDA

3

1. 120-130 KIAS When Landing assured:

2. Flaps—Down 3. Transition to VREF

4. Yaw Damper—Off

5

MAP (and during circling maneuver)

1. Determine that visual contact with the runway environment can be maintained and a normal landing can be made from a circling approach, or initiate a missed approach.

2. Maintain MDA during circling

4 BASE

maneuver. CAUTION

1. Commence descent from a point where a normal landing can be made.

To ensure constant reversing characteristics, the propeller control must be in full increase RPM position. If possible, propellers should be moved out of reverse at approximately 40 knots to minimize blade erosion. Care must be exercised when reversing on runways with loose sand, dust or snow on the surface. Flying gravel will damage propeller blades and dust or snow may impair the pilot’s visibility.

Pilot Tip Reverse is most effective at higher speeds and braking is most effect at lower speeds.

1

Nonprecision Approach — Procedure Turn

4 5

2

9

1 Initial Approach 1. Obtain ATIS 2. Preview Approach

& Missed Approach

Procedure Turn Outbound 1. Start Timing 2. Flaps — Approach

3. Navaids - Tune / Ident /Load GPS

3. 120 - 130 KIAS Procedure Turn Inbound 1. FD—As Desired 2. Reset Altitude Alerter

Arrival

1. Torque—Approx. 500lbs 2. 140 KIAS

Station Passage

1. Start Time 2. Set Altitude Alerter 3

3 FD As Desired 4. Start Before Landing Checklist

FAF

Intercept Final Approach 1. Course — Inbound

MAP - Missed Approach 1. Power Maximum 2. Pitch— 7° Nose-Up (FD-GA) 3. Flaps—Up 4. Gear— Up 5. Complete Missed Approach Procedure

10 MAP

MDA

10

FAF

6 7 Approach Inbound

1. Reset Altitude Alerter

8 Final Approach Fix 1. Start Timing 2. Gear Down 3. Torque - approx. 200lbs 4. Complete Before Landing

Checklist MAP - Landing Assured

5. 120 - 130 KIAS

THRESHOLD

1. Flaps—Down 2. Transition to VREF

3. Yaw Damper—Off

11

MDA (minimum descent altitude)

1. Level off at MDA at least 1 mile prior to MAP if possible.

2. Torque — 600 - 1,000 lbs 3. 120-130 KIAS

Landing 12 1. Props Full Forward 2. BETA or Reverse 3. Brakes—As Necessary

1. Gear—Re-check Down 2. Airspeed— VREF

3. Power—Idle

CAUTION To ensure constant reversing characteristics, the propeller control must be in full increase RPM position. If possible, propellers should be moved out of reverse at approximately 40 knots to minimize blade erosion. Care must be exercised when reversing on runways with loose sand, dust or snow on the surface. Flying gravel will damage propeller blades and dust or snow may impair the pilot’s visibility.

Pilot Tip 3 Reverse is most effective at higher speeds and braking is most effect at lower speeds.

One Engine Inoperative — Visual Approach & Landing

4

1 Initial Approach 1. Obtain ATIS 2. Descent Checklist Complete

Go Around 1. Power — Max 2. Gear — Up 3. Flaps — Up 4. Airspeed — Increase to VYSE

(Blue Line)

Arrival 2 1. Torque—Approx. 1000lbs 2. 140 KIAS 3. Start One-engine Inoperative

Approach & Landing Checklist

7 Threshold

1. Gear — Recheck Down 2. Airspeed — VREF

3. Power — Idle 8

Downwind 3 1. Flaps — Approach 2. 130-140 KIAS

9 1. BETA or Reverse—As Necessary 2. Brakes—As Necessary

1. Gear — Down 2. Prop — Full Forward

4

Base

6 Final

5 1. 120-130 KIAS

1. 120KIAS WHEN LANDING ASSURED: 2. Flaps — Down 3. Transition to VREF

4. Yaw Damper — Off 5. One - Engine - Inoperative

Approach and Landing Checklist —Complete



5

Landing From an ILS

1 5 Glide Slope Intercept

Initial Approach 1. Obtain ATIS 2. Preview Approach

1. Torque—Approx. 500 lbs 2. 120 KIAS (VYSE MIN)

OM & Missed Approach 3. Navaids - Tune / Ident 4. Descent Checklist Complete

6 DH - Missed Approach MM 1. Power Maximum 2. Pitch— 7° Nose-Up (FD-GA) 3. Flaps—Up 4. Gear—Up 5. Complete Missed Approach Procedure

2 Arrival

1. Torque—Approx. 500lbs 2. 140 KIAS (Typical) 3. FD — As Desired 4. Start Before Landing Checklist

DH 3 Approach Inbound

1. Flaps — Approach 2. 120-130 KIAS

4 Approaching Glide Slope 1. Gear — Down 2. Complete Before Landing Checklist

8 Landing

6

7 Threshold

DH—Visual & Landing Assured

1. Flaps — Down 2. Transition to VREF

3. Yaw Damper — Off

1. Props — Full Forward 2. BETA or Reverse 2. Brakes—As Necessary

1. Gear — Recheck Down 2. Airspeed — VREF

3. Power — Idle



6

Visual Approach and Landing

1 Initial

Rejected Landing 1. Power Maximum 2. Pitch— 10° Nose-Up 3. Airspeed—100 KIAS 3. Establish Normal Climb When

Clear of Obstacles 5. Flaps—Up

1. Obtain ATIS 2. Descent Checklist Complete

6. Gear—Up

8 Threshold 1. Gear—Re-check Down 2. Airspeed— VREF

3. Power—Idle

2 Arrival

1. Torque—Approx. 800lbs 2. 140 KIAS (Typical)

7 Final 1. 120 KIAS (VYSE MIN)

When Landing assured: 2. Flaps—Down 3. Transition to V

3. Start Before Landing Checklist 4. Yaw REF

Damper—Off

6

Downwind 1. Flaps—Approach 2. 120 KIAS

3 Abeam Touchdown Point 1. Gear — Down

9 Landing 1. Props—Full Forward 2. BETA or Reverse 3. Brakes— as Necessary

4 2. Before Landing Checklist Complete

Base 5 1. 120 KIAS

7

Approach to Stall — Landing Configuration

Horn or Buffet V2

1 Beginning of Maneuver Initial Condition 1. Torque—200lbs 2. Propellers— 1,900 RPM 3. Maintain Initial Heading 4. Maintain Initial Pitch 5. Flaps—Approach

(White Triangle) 6. Gear—Down (Below VLE ) 7. Flaps— Down 100%

(Below Top of White Arc) 8. Pitch Attitude Prior to Horn

or Buffet May Reach 10° - 15°, Depending on Technique

9. Horn Will Sound Approximately 10kts above Buffet

3

2 Stall and Recovery At Horn or Buffet—

Recover 1. Simultaneously Advance the

Power Levers Toward MAX Torque, Propeller Levers Full Forward, Reduce the Pitch Attitude as Necessary to Stop the Stall Warning, and Roll the Wings Level.

2. Establish Positive Rate of Climb

3. Flaps— Up, at or Above 100 KIAS

4. Gear— Up

Completion of Maneuver Completion 1. Level Off at Initial Altitude

and Heading 2. Reset Power as Required

8

Approach to Stall — Takeoff Configuration

Horn or Buffet V2

1 Beginning of Maneuver Initial Condition 1. Torque—200lbs 2. Propellers— 2,200 RPM 3. Maintain Initial Heading 4. Maintain Initial Altitude 5. Flaps—Approach

(Below White Triangle) 6. At 110 KIAS or Less,

Simultaneously Set the Torque to 700 lbs (Simulated 100% Torque), Establish a Bank Angle of 20° (No More Than 30°), and Raise the Nose and Climb)

7. Student May be Required to Perform this Maneuver While Maintaining 15° - 30° Angle of Bank or While Maintaining a Heading

8. Clear Area in Direction of Turn

9. Decrease Speed Approximately 1 Knot per Second

10. Pitch Attitude Prior to Horn or Buffet May Reach 15°-25°, Depending on Technique

3

2 Stall and Recovery At Horn or Buffet—

Recover 1. Reduce the Pitch Attitude as

Necessary to Stop the Stall Warning, and Roll the Wings Level.


3. Flaps— Up, at or Above VYSE

(Blue Line)

Completion of Maneuver Completion 1. Level Off at Initial Altitude


9

Approach to Stall — Clean Configuration

Horn or Buffet V2

1 Beginning of Maneuver Initial Condition 1. Torque—200lbs 2. Propellers— 1,900 RPM 3. Maintain Initial Heading 4. Maintain Initial Altitude 5. Pitch Attitude Prior to Horn

or Buffet May Reach 10° - 15°, Depending on Technique

6. Horn Will Sound Approximately 10kts above Buffet

3

2 Stall and Recovery 1. Simultaneously Advance the

Power Levers Toward MAX Torque, Reduce the Pitch Attitude as Necessary to Stop the Stall Warning, and Roll the Wings Level.


Completion 1. Level Off at Initial Altitude


10

Rejected Takeoff

Em

1 Before Takeoff 1. Follow Normal Takeoff

Procedures Until Initiating Abort at or Below V1

3 Clear of Runway 1. Complete After Landing

Checklist

2 ergency or Malfunction

At or Below V1 1. Recognize Reason for

Rejected Takeoff 2. Power Levers— Idle 3. Braking— As Necessary 4. Reverse— As Necessary 5. Maintain Runway Heading

Pilot Tip If rejected takeoff is due to reasons other than one engine power loss,

Reverse is most effective at higher speeds and braking is most effect at lower speeds.

11

Normal Takeoff and Departure

Cruise 8 1. Accelerate to Cruise Speed 2. Set Cruise Power 3. Complete Cruise Checklist

VYSE or Above 1. Flaps— Up 2. Yaw Damper— On 3. Climb Power— Set

6 Climb-Out 1. Accelerate to 150 KIAS 2. Landing / Taxi Lights— Out 3. Complete Climb Checklist

5

Area Departure / Climb Profile 7 1. 150 KIAS to 10,000 Ft 2. 130 KIAS to 10,000 - 20,000 Ft 3. 120 KIAS 20,000 - 25,000 Ft 4. 110 KIAS 25,000 - 35,000 Ft

Takeoff 4 1. Rotate at V1 to Approx 7°

Nose Up 2. Establish Positive Rate of

Climb 3. Landing Gear— Up

3 Takeoff Roll 1. Recheck Torque / ITT 2. Annunciators— Check

1 Before Takeoff 1. Checklist— Complete 2. Recheck V1 and V2

2 In Position 1. Hold Brakes 2. Props— 2,000 RPM

(On Governors) 3. Release Brakes 4. Set Torque

12

Beechcraft King Air 90 Series

Approach Power Setting Recommendations

Flight Condition Torque Flaps Gear IAS ROC (ftlbs) (kts) (fpm)

Two Engines - 1900 RPM Initial Maneuvering: Level 800 Up Up 160 0 Descending 300 Up Up 160 -1000 Approaching FAF 500 Apr Up 120 0 FAF Inbound: Precision Approach 500 Apr Dn 120 -600 Non-Precision Approach Descending 300 Apr Dn 120 -1000 Level at MDA (Category B) 800 Apr Dn 120 0 Visual Final 500 Dn Dn 100 ± -500 Missed Approach 1200*** Up Up 120 ± *

Single Engine – 2200 RPM

Initial Maneuvering: Level 900 Up Up 130 0 Descending 300 Up Up 130 -1000 Approaching FAF 900 Apr Up 120 0 FAF Inbound: Precision Approach 900 Apr Dn 120 -600 Non-Precision Approach Descending 300 Apr Up 120 -1000 Level at MDA (Category B) 900 Apr Up 120 0 Missed Approach 1315*** Up Up 120 ± *

* Torque may vary by approximately 100 tt-lbs due to differences in weight and in general airplane rigging/condition.

** For a precision approach, adjust torque as necessary to maintain 120

knots on the glide path. The torque figures presented work well for a 3° glide path in still air.

*** Unless restricted by climb ITT limits. For E90, use 1400 ft-lbs or 700°.

**** Or redline ITT, whichever comes first

13

King Air C90 Fuel System

NOTE Total Usable Fuel 384 Gallons (US)

NOTE

Right system is identical to left system except that the left contains the crossfeed valve. It should also be noted that the purge valve and fuel line are located on the inboard side of the nacelle and that there is a thermal relief valve and line from the crossfeed line in the right fuel system.

• Valve has holes for flow out at reduced rate. Only 28 of 44 gallons will not gravity feed to nacelle.

NOTE

A fuel capacitance gaging system utilizes a single fuel quantity gage for each wing fuel system. This gage can be switched to designate the amount of fuel in the nacelle tank or the total fuel in the system.

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