King Air 200 – The Training Workbook
Copyright © 2011
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 BE200 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: AIRCRAFT - GENERAL ............................................................ 11
INTRODUCTION TO THE KING AIR 200 AND B200 ....................................................11
OBJECTIVES .......................................................................................................................11
GENERAL ...........................................................................................................................12
NOSE SECTION ..................................................................................................................13
COCKPIT .............................................................................................................................13
LIGHTING SYSTEMS ........................................................................................................15
CABIN CONFIGURATION ................................................................................................16
CABIN WINDOWS .............................................................................................................22
EMERGENCY EXIT ...........................................................................................................24
INTERIOR DIVIDERS ........................................................................................................24
AFT FUSELAGE .................................................................................................................24
EMPENNAGE .....................................................................................................................25
WINGS .................................................................................................................................25
POWER PLANT ..................................................................................................................27
ELECTRICAL SYSTEM .....................................................................................................27
PROPELLER SYSTEM .......................................................................................................27
FUEL SYSTEM ...................................................................................................................27
ANTI-ICE/DE-ICE SYSTEMS ............................................................................................28
ENVIRONMENTAL SYSTEM ...........................................................................................28
LIMITATIONS ........................................................................................................................... 29
AIRSPEED LIMITATIONS ................................................................................................31
WEIGHT LIMITS ................................................................................................................32
CENTER OF GRAVITY LIMITS .......................................................................................32
EMERGENCY PROCEDURES................................................................................................ 33
EXPANDED GENERAL PROCEDURES ............................................................................... 35
QUESTIONS ............................................................................................................................... 36
CHAPTER 2: ELECTRICAL SYSTEM ............................................................. 38 OBJECTIVES .......................................................................................................................38
ELECTRICAL POWER - DESCRIPTION AND OPERATION ........................................39
BATTERY SYSTEM ...........................................................................................................41
DC GENERATION - DESCRIPTION AND OPERATION ...............................................42
STARTER-GENERATORS .................................................................................................43
GENERATOR CONTROL UNIT .......................................................................................43
STARTER-GENERATOR PARALLELING ......................................................................44
OVER VOLTAGE PROTECTION......................................................................................44
REVERSE CURRENT PROTECTION ...............................................................................45
OVER EXCITATION PROTECTION ................................................................................45
COMPONENT LOCATION ................................................................................................45
AC GENERATION ..............................................................................................................46
EXTERNAL POWER ..........................................................................................................46
AVIONIC MASTER SWITCH ............................................................................................47
CIRCUIT BREAKERS ........................................................................................................48
LIMITATIONS ........................................................................................................................... 48
EXTERNAL POWER LIMITS ............................................................................................48
GENERATOR LIMITS ........................................................................................................48
STARTER LIMITS ..............................................................................................................49
EMERGENCY ELECTRICAL PROCEDURES .................................................................... 49
ABNORMAL ELECTRICAL PROCEDURES ....................................................................... 51
EXPANDED ELECTRICAL PROCEDURES ........................................................................ 53
QUESTIONS ............................................................................................................................... 55
CHAPTER 3: ANNUNCIATOR SYSTEM ......................................................... 58 OBJECTIVES .......................................................................................................................58
ANNUNCIATOR SYSTEM ................................................................................................58
WARNING PANEL .............................................................................................................58
CAUTION/ADVISORY PANEL ........................................................................................59
ANNUNCIATOR LIMITATIONS............................................................................................ 60
ANNUNCIATOR EMERGENCY PROCEDURES ................................................................ 60
ANNUNCIATOR ABNORMAL PROCEDURES ................................................................... 60
QUESTIONS ............................................................................................................................... 61
CHAPTER 4: FUEL SYSTEM ............................................................................. 62 OBJECTIVES .......................................................................................................................62
FUEL SYSTEM - DESCRIPTION AND OPERATION .....................................................62
FUEL GAUGES ...................................................................................................................64
FUEL DRAIN VALVES ......................................................................................................64
FUEL VENTS ......................................................................................................................65
FUEL PUMPS ......................................................................................................................65
AUXILIARY FUEL TRANSFER SYSTEM .......................................................................67
FUEL FILTERS ...................................................................................................................68
FUEL HEATER ...................................................................................................................69
CROSSFEED .......................................................................................................................69
FUEL PURGE SYSTEM .....................................................................................................70
FUEL SYSTEM LIMITATIONS .............................................................................................. 70
FUEL LIMITATIONS .........................................................................................................70
APPROVED ENGINE FUELS ............................................................................................70
EMERGENCY ENGINE FUELS ........................................................................................70
LIMITATIONS ON THE USE OF AVIATION GASOLINE .............................................71
APPROVED FUEL ADDITIVES ANTI-ICING ADDITIVES ..........................................71
FUEL BIOCIDE ADDITIVE ...............................................................................................72
EMERGENCY FUEL SYSTEM PROCEDURES .................................................................. 73
ABNORMAL FUEL PROCEDURES....................................................................................... 74
EXPANDED FUEL PROCEDURES ........................................................................................ 75
QUESTIONS ............................................................................................................................... 76
CHAPTER 5: ENGINE SYSTEM ....................................................................... 79 OBJECTIVES .......................................................................................................................79
GENERAL ENGINE DESCRIPTION .................................................................................79
PROPULSION SYSTEM CONTROLS ...............................................................................80
TURBOPROP ENGINE SYMBOLS AND THEIR MEANINGS ......................................82
AIR INTAKE SECTION .....................................................................................................83
COMPRESSOR SECTION ..................................................................................................83
COMPRESSOR BLEED VALVES .....................................................................................84
COMBUSTION SECTION ..................................................................................................85
TURBINE SECTION ...........................................................................................................85
EXHAUST SECTION ..........................................................................................................86
REDUCTION GEAR SECTION .........................................................................................86
THE ACCESSORY SECTION ............................................................................................86
ENGINE LUBRICATION SYSTEM ..................................................................................86
OIL TANK ...........................................................................................................................87
PUMPS .................................................................................................................................88
OIL FILTER .........................................................................................................................88
OIL COOLER ......................................................................................................................88
OIL TEMPERATURE .........................................................................................................89
OIL PRESSURE ...................................................................................................................89
CHIP DETECTION ..............................................................................................................89
FUEL HEATER ...................................................................................................................89
ENGINE FUEL SYSTEM ...................................................................................................90
FUEL CONTROL UNIT ......................................................................................................90
STARTING AND IGNITION SYSTEM ............................................................................91
AUTO IGNITION ...............................................................................................................92
FIRE DETECTION SYSTEM (BB-2 through BB-1438) ....................................................92
FIRE DETECTION SYSTEM (BB-1439 AND AFTER) ....................................................93
FIRE EXTINGUISHING SYSTEM ....................................................................................95
ENGINE SYSTEM LIMITATIONS ......................................................................................... 97
EMERGENCY ENGINE SYSTEM PROCEDURES ........................................................... 100
ABNORMAL ENGINE SYSTEM PROCEDURES .............................................................. 104
EXPANDED ENGINE SYSTEM PROCEDURES................................................................ 108
ENGINE STARTING (EXTERNAL POWER) .................................................................108
QUESTIONS ............................................................................................................................. 111
CHAPTER 6: PROPELLERS ............................................................................114 OBJECTIVES .....................................................................................................................114
GENERAL .........................................................................................................................114
BASIC PRINCIPLES .........................................................................................................115
PROPELLER GOVERNOR ..............................................................................................116
PRIMARY GOVERNOR ...................................................................................................116
OVERSPEED GOVERNOR ..............................................................................................118
FUEL TOPPING GOVERNOR .........................................................................................118
PROPELLER FEATHERING ............................................................................................119
AUTOFEATHER ...............................................................................................................119
PROPELLER BETA AND REVERSING .........................................................................120
PROPELLER SYNCHROPHASER ..................................................................................121
PROPELLER LIMITATIONS................................................................................................ 123
PROPELLER EMERGENCY PROCEDURES .................................................................... 123
PROPELLER ABNORMAL PROCEDURES ....................................................................... 123
PROPELLER EXPANDED PROCEDURES ........................................................................ 123
QUESTIONS ............................................................................................................................. 125
CHAPTER 7: PRESSURIZATION AND ENVIRONMENTAL SYSTEMS .............................................................................................................127
OBJECTIVES .....................................................................................................................127
INTRODUCTION ..............................................................................................................127
HEATING, COOLING AND PRESSURIZATION - DESCRIPTION AND OPERATION .....................................................................................................................128
HEATING TEMPERATURE CONTROL - DESCRIPTION AND OPERATION .....................................................................................................................130
AUTOMATIC OPERATION ............................................................................................130
MANUAL HEAT OPERATION .......................................................................................130
RADIANT HEAT PANELS ..............................................................................................131
ELECTRIC HEAT .............................................................................................................132
FRESH AIR VENTILATION ............................................................................................132
COOLING - DESCRIPTION AND OPERATION ...........................................................133
AIR CONDITIONING TEMPERATURE CONTROL DESCRIPTION AND OPERATION .....................................................................................................................134
AUTOMATIC OPERATION ............................................................................................134
MANUAL COOL OPERATION .......................................................................................135
FORWARD EVAPORATOR FREEZE PROTECTION ...................................................136
PRESSURIZATION - DESCRIPTION AND OPERATION ............................................136
FLOW CONTROL UNIT ..................................................................................................136
OXYGEN SYSTEM ..........................................................................................................139
PRESSURIZATION AND ENVIRONMENTAL SYSTEMS LIMITATIONS .................. 141
EMERGENCY PRESSURIZATION AND ENVIRONMENTAL SYSTEMS PROCEDURES ......................................................................................................................... 141
ABNORMAL PRESSURIZATION AND ENVIRONMENTAL SYSTEMS PROCEDURES ......................................................................................................................... 146
PRESSURIZATION AND ENVIRONMENTAL SYSTEMS EXPANDED PROCEDURES ......................................................................................................................... 147
PRESSURIZATION TEST ................................................................................................147
OXYGEN SYSTEM PREFLIGHT INSPECTION ............................................................147
QUESTIONS ............................................................................................................................. 149
CHAPTER 8: LANDING GEAR, TIRES AND BRAKE SYSTEM ..............151 OBJECTIVES .....................................................................................................................151
GENERAL .........................................................................................................................151
GROUND HANDLING TOWING ....................................................................................152
PARKING ..........................................................................................................................153
NOSE LANDING GEAR ...................................................................................................153
DESCRIPTION AND OPERATION - MECHANICAL LANDING GEAR ....................154
WARNING SYSTEM MECHANICAL LANDING GEAR SYSTEM ............................156
DESCRIPTION AND OPERATION- HYDRAULIC LANDING GEAR ........................157
WARNING SYSTEM HYDRAULIC LANDING GEAR SYSTEM ................................161
TIRES .................................................................................................................................161
HYDRAULIC BRAKE SYSTEM .....................................................................................162
LANDING GEAR, TIRES AND BRAKE SYSTEM LIMITATIONS ................................ 164
LANDING GEAR CYCLE LIMITS ..................................................................................164
LANDING GEAR, TIRES AND BRAKE SYSTEM ABNORMAL PROCEDURES ......................................................................................................................... 164
LANDING GEAR, TIRES AND BRAKE SYSTEM EMERGENCY PROCEDURES ......................................................................................................................... 165
LANDING GEAR, TIRES AND BRAKE SYSTEM EXPANDED PROCEDURES ......................................................................................................................... 167
QUESTIONS ............................................................................................................................. 168
CHAPTER 9: PNEUMATIC AND VACUUM SYSTEM ...............................170 OBJECTIVES .....................................................................................................................170
DESCRIPTION ..................................................................................................................170
PNEUMATIC - DESCRIPTION AND OPERATION ......................................................170
VACUUM SYSTEM - DESCRIPTION AND OPERATION ...........................................171
ENGINE BLEED-AIR-WARNING SYSTEM - DESCRIPTION AND OPERATION .....................................................................................................................172
PNEUMATIC AND VACUUM SYSTEM LIMITATIONS ................................................. 174
PNEUMATIC AND VACUUM SYSTEM EMERGENCY PROCEDURES...................... 174
PNEUMATIC AND VACUUM SYSTEM ABNORMAL PROCEDURES ........................ 175
PNEUMATIC AND VACUUM SYSTEM EXPANDED PROCEDURES .......................... 175
QUESTIONS ............................................................................................................................. 176
CHAPTER 10: ANTI-ICE SYSTEM .................................................................177 OBJECTIVES .....................................................................................................................177
DESCRIPTION ..................................................................................................................177
ICE AND RAIN PROTECTION - DESCRIPTION AND OPERATION .........................177
AIRFOIL ............................................................................................................................177
DEICE BOOT - PROTECTIVE COATING ......................................................................179
AIR INTAKES ...................................................................................................................180
DUAL-MOTOR INERTIAL ICE SEPARATION SYSTEM ............................................181
AIR INTAKE ANTI-ICE LIP ............................................................................................182
BRAKE DEICE SYSTEM .................................................................................................183
WINDOWS AND WINDSHIELDS ..................................................................................184
PROPELLER DEICING ....................................................................................................185
PITOT HEAT .....................................................................................................................187
STALL WARNING VANE HEAT ....................................................................................187
ANTI-ICING SYSTEMS LIMITATIONS ............................................................................. 188
ANTI-ICE SYSTEM EMERGENCY PROCEDURES ......................................................... 189
ANTI-ICE SYSTEM ABNORMAL PROCEDURES ........................................................... 189
ELECTROTHERMAL PROPELLER DEICE (Manual System) ......................................190
ENGINE ICE VANE-FAILURE (L or R ICE VANE Annunciator) .................................191
ANTI-ICE SYSTEM EXPANDED PROCEDURES ............................................................. 191
QUESTIONS ............................................................................................................................. 193
CHAPTER 11: FLIGHT CONTROLS ..............................................................195 OBJECTIVES .....................................................................................................................195
FLIGHT CONTROLS ........................................................................................................195
ELEVATOR TRIM ............................................................................................................196
CONTROL LOCKS ...........................................................................................................198
GROUND MOORING/TOWING ......................................................................................198
WING FLAPS ....................................................................................................................199
YAW DAMPER .................................................................................................................201
STALL WARNING SYSTEM ...........................................................................................201
STALL WARNING ACTIVATES ....................................................................................201
RUDDER BOOST ..............................................................................................................202
FLIGHT CONTROL LIMITATIONS ................................................................................... 202
FLIGHT CONTROL EMERGENCY PROCEDURES ........................................................ 203
FLIGHT CONTROL ABNORMAL PROCEDURES .......................................................... 204
FLIGHT CONTROLS EXPANDED PROCEDURES .......................................................... 206
QUESTIONS ............................................................................................................................. 208
CHAPTER 12: PITOT STATIC SYSTEM .......................................................210 OBJECTIVES .....................................................................................................................210
PITOT AND STATIC PRESSURE SYSTEM ...................................................................210
OUTSIDE AIR TEMPERATURE .....................................................................................211
PITOT STATIC SYSTEM LIMITATIONS .......................................................................... 211
PITOT STATIC SYSTEM EMERGENCY PROCEDURES ............................................... 211
PITOT STATIC SYSTEM ABNORMAL PROCEDURES ................................................. 211
QUESTIONS ............................................................................................................................. 213
CHAPTER 13: OXYGEN SYSTEM ..................................................................214 OBJECTIVES .....................................................................................................................214
OXYGEN SYSTEM - DESCRIPTION AND OPERATION ............................................214
AUTO DEPLOYMENT PASSENGER OXYGEN SYSTEM ..........................................216
OXYGEN CYLINDERS ....................................................................................................216
OXYGEN PRESSURE-SENSE SWITCH .........................................................................217
OXYGEN SYSTEM LIMITATIONS ..................................................................................... 219
OXYGEN SYSTEM EMERGENCY PROCEDURES.......................................................... 219
OXYGEN SYSTEM ABNORMAL PROCEDURES ............................................................ 220
QUESTIONS ............................................................................................................................. 221
CHAPTER 14: POWER SETTINGS AND PROFILES ....................................................... 222
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CHAPTER 1
AIRCRAFT - GENERAL
INTRODUCTION TO THE KING AIR 200 AND B200
The King Air 200 workbook describes the airframe, engines and systems of the King Air 200
and B200. It is a compilation of operating information, tips and techniques that I have gathered
over the past 20 years as a King Air pilot and instructor. 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 completing this chapter, you will be able to:
Locate and describe:
Entry Door/Emergency Exit Avionics Area Fuselage Baggage Area Cabin Section Wing Section Fuselage Lights
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GENERAL
The King Air 200 is a high performance, all metal, low wing aircraft that has been in continuous
production since 1974. Originally introduced as the Super King Air, the word “super” was
dropped in 1996 as a marketing decision. An updated and improved version of the airplane
entered production in 1981 and became known as the B200. Approximately 3500 King Air 200’s
are in service today with numerous variants, including cargo and military versions. The airplane
is approved for day and night IFR and VFR flight operations and, if properly equipped, is
capable of flight into known icing. It has fully cantilevered wings and a T-tail. By locating the
horizontal stabilizer as high as possible, it stays out of the air disturbance created by the
propellers. The advantages of this design are less airframe vibration, wider C.G. range, and fewer
trim adjustments are necessary during airspeed or configuration changes. The fuselage is
pressurized to the skin between fore and aft pressure bulkheads. The control cables, torque
shafts, plumbing and wiring connections that pass through pressure walls are installed with fitted
seals or plug connectors that minimize air leakage. Like most modern turboprops, the King Air
200 fuselage is of semimonocoque construction and is fabricated utilizing aluminum frames,
bulkheads and keels that are reinforced by longerons and stringers. It is powered by two 850
SHP Pratt & Whitney turboprop engines. The 200 is equipped with two PT6A-41 engines while
the B200 utilizes the PT6A-42. The -42 engine is also rated at 850 shp but has internal
improvements that result in greater engine performance over a wider range of temperatures and
altitudes. 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 15 people,
although the normal corporate configuration is 7-8 passengers.
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NOSE SECTION
The nose section of the airplane houses the radar antenna dish and the avionics bay. It also
contains the hydraulic brake fluid reservoir, the vacuum system inlet and some components of
the air conditioner. Except for the compressor, the nose section is un-pressurized and is accessed
via removable panels on each side of the compartment. The radome is constructed of fiberglass
allowing radar waves to pass through it easily.
COCKPIT
Seats
The pilot's seats are adjustable both fore and aft, as well as vertically. Additionally, three tilt
adjustments are possible. 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. A preflight flow pattern is essential to the safe operation of the King Air by
a single pilot. Flow patterns do not replace checklists but are used to methodically set up the
aircraft prior to each phase of flight.
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Flow Patterns
Because of the wide variation in switch location, each pilot should develop a flow pattern that
incorporates their particular airplane. A good flow pattern starts at the end of the console and
follows the diagram arrows. Each switch is checked and positioned for the pertinent phase of
flight. This is a generic flow pattern that after completion should be followed by the appropriate
checklist.
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.
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Oxygen Masks
The quick donning oxygen masks for the crew are stored on the bulkhead behind the pilots.
Newer aircraft are equipped with masks stowed directly above each pilot station.
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 pilots, 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 interior light 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
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accompanying chimes. This three-position switch is placarded CABIN LIGHTS OFF. - NO
SMOKE & FSB.
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. 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 sub-panel. 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 seven passengers. All seats are equipped with seat belts and headrests. Some
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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. When 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
across from the airstair door.
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An optional forward facing unit may be installed in the aft baggage compartment. Either
installation is equipped with a hinged cushion cover turning the toilet into an additional
passenger seat. The seat belt and shoulder harness for the toilet seat is attached to the bulkhead.
Relief tubes are located on the left cabin side wall forward of the toilet and in the cockpit under
the pilot's seat.
PILOT TIP
If a Monogram electrically flushing toilet is installed, the sliding knife valve should be open at all times, except when actually servicing the unit.
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 410 pounds of baggage storage. Optional folding jumpseats can
be installed in the compartment. All baggage and cargo must be properly secured with the
webbing provided.
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PILOT TIP
Do not carry children in the baggage compartment unless secured by a seatbelt in a seat.
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 two drawers installed
beneath the couch and in the armrest cabinet located adjacent to the aft end of
the couch. An optional cabinet can be installed forward of the main cabin aft
partition.
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 on the B200. The forward assist cable
is easily detachable to provide more room for loading large baggage or cargo into the airplane.
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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 ensure structural integrity. Proper latching of the door can be verified by
both observing an annunciator light in the cockpit and by visually confirming the alignment
position marks on the bayonet 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
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safety release button and rotate the handle counterclockwise. 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 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.
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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 push- button 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 operate 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
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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 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 which adjusts the 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.
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WARNING
DO NOT LOOK DIRECTLY AT THE SUN, EVEN THROUGH POLARIZED WINDOWS. EYE DAMAGE COULD RESULT.
EMERGENCY EXIT
The emergency exit door (19” X 27”) is
located on the right cabin side wall just
aft of the copilot's seat. Inside the
airplane, the exit door is released by a
pull-down handle. The exit can be
opened from outside the aircraft by
pulling on a flush mounted handle. The
door is a non-hinged, plug-type which
removes completely from the frame
when the latches are released. The door can be locked from the inside with a key to prevent
access from the outside. The inside handle will override the locking mechanism. The exit should
be unlocked prior to flight to allow access to the cabin from the outside in the event of an
emergency. The key remains in the lock when the door is locked and can be removed only when
the door is unlocked. The key slot is in the vertical position when the door is unlocked. Removal
of the key from the lock before flight assures the pilot that the door can be removed from the
outside if necessary.
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.
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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 T-Tail empennage configuration. The aircraft is equipped with
a rudder boost system which will automatically apply pressure to the appropriate rudder if an
engine fails. 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. Static wicks minimize the effects of static build up on the aircraft
structures. The pneumatic de-ice boots are attached to the leading edges of the horizontal
stabilizers.
PILOT TIP
One static wick can be missing from each side of the horizontal stabilizer and one can be missing from the vertical stabilizer.
WINGS
The airplane utilizes a NACA 23000 series wing shape. This airfoil 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
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54'6" and incorporates a 6 degree wing dihedral. The total wing area is 303 sq. feet. The Beech
King Air 200 and B200 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. The
center section and outboard wing assemblies are semi-monocoque box construction. Both center
section spars are I-beam sections built up from extruded aluminum. The wing structure
incorporates continuous dual spar structures (front and rear) from tip to tip.
The forward wing spar structure, the most critical element of the wing from a structural integrity
standpoint, incorporates fail-safe type construction. The lower element of the forward spar cap is
made up of 3 elements bonded together. If a flaw should develop in the cross section of any
element, the flaw would stop progressing at the bond line of the adjoining element rather than
progressing completely through the section. A sealed integral (wet wing) fuel tank is installed in
the outboard end of each wing assembly. The tank interior is coated for corrosion protection.
Inboard of the integral tanks, bladder fuel tanks are installed. 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 provide
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for redundancy in the cockpit. Static wicks minimize the effects of static build up on the aircraft
structures.
PILOT TIP
One static wick can be missing or broken on each wing.
POWER PLANT
The aircraft is powered by two 850 shp Pratt and Whitney PT6A-41 or PT6A-42 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
two-stage 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 “dual fed” 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 inverters.
More information on the electrical system is supplied in Chapter 2 of this workbook.
PROPELLER SYSTEM
The aircraft 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 544 usable gallon system with each wing divided into a main and an
auxiliary system. The main system is comprised of five outboard wing tanks which include four
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bladder types and one wet-wing type and the nacelle bladder tank. These are all interconnected
by gravity feed lines and flow into the nacelle tank. 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.
An optional brake deice system is also available. 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.
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LIMITATIONS All airspeeds quoted in this section are indicated airspeeds (IAS) and assume zero instrument error.
AIRSPEEDS FOR SAFE OPERATION (12,500 LBS)
Maximum Demonstrated Crosswind Component……………….…………………… 25 Knots
Takeoff (Flaps Up)
Rotation………………………………………………………………………… 95 Knots 50-ft Speed………...……………………………………………………………121 Knots
Takeoff (Flaps Approach)
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Rotation………………………………………………………………………… 94 Knots 50-ft Speed………...……………………………………………………………106 Knots
Two-Engine Best Angle-of-Climb (Vx)………………………………………………..100 Knots Two-Engine Best Rate-of-Climb (Vy)……..…………………………………………..125 Knots
Cruise Climb: Sea Level to 10,000 feet…………………………………………………….. 160 Knots 10,000 to 20,000 feet………………………………………………………….140 Knots 20,000 to 25,000 feet………………………………………………………….130 Knots 25,000 to 35,000 feet………………………………………………………….120 Knots
Maximum Airspeed for Effective Windshield Anti-icing……………………………226 Knots Maneuvering Speed (VA)……………………………………………………………...181 Knots Turbulent Air Penetration……………………………………………………………..170 Knots
For turbulent air penetration, use an airspeed of 170 knots. Avoid over-action on power levers.
Turn off autopilot altitude hold. Keep wings level, maintain attitude and avoid use of trim. Do
not chase airspeed and altitude. Penetration should be at an altitude which provides adequate
maneuvering margins when severe turbulence is encountered.
Landing Approach: Flaps Down…………………………………………………………………..103 Knots Balked Landing Climb……………………………………………………….100 Knots
Intentional One-Engine-lnoperative Speed (VSSE)……………………...……………104 Knots Air Minimum Control Speed (VMCA)…………………………………….…………….86 Knots
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AIRSPEED LIMITATIONS
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WEIGHT LIMITS
Maximum Ramp Weight 12,590 pounds
Maximum Take-off Weight 12,500 pounds
Maximum Landing Weight 12,500 pounds
Maximum Zero Fuel Weight 11,000 pounds
Maximum Weight in Baggage Compartment:
BB-1091 and after:
When Equipped with Fold-up Seats 510 pounds
When Not Equipped with Fold-up Seats 550 pounds
Prior to BB-1091:
When Equipped with Fold-up Seats 370 pounds
When Not Equipped with Fold-up Seats 410 pounds
CENTER OF GRAVITY LIMITS
Aft Limit
196.4 inches aft of datum at all weights
Forward Limits
185.0 inches aft of datum at 12,500 pounds, with straight line variation to 181.0 inches aft of
datum at 11,279 pounds. 181.0 inches aft of datum at 11,279 pounds or less.
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EMERGENCY PROCEDURES The pilot in command of an aircraft is directly responsible for and is the final authority as to the
operation of that aircraft. In an emergency requiring immediate action, the pilot in command may
deviate from any rule in 14 CFR Part 91, Subpart A, General, and Subpart B, Flight Rules, to the
extent required to meet that emergency. The following section deals with situations that require
immediate and accurate action by the crew. Memory items are printed in bold type and should be
completed in a timely manner. However, acting too rapidly may compound the emergency and
place the aircraft in an unrecoverable situation. To prevent this, memory items must be
accomplished methodically and must include coordination between the pilots.
The following steps should be committed to memory and considered mandatory in any emergency:
1. Fly the airplane.
2. Identify the emergency.
3. Complete the appropriate checklist.
BOLD TYPE INDICATES MEMORY ITEMS!
CABIN OR CARGO UNLOCKED (CABIN DOOR Annunciator)
WARNING
DO NOT ATTEMPT TO CHECK THE SECURITY OF THE AIRSTAIR OR CARGO DOOR IN FLIGHT. REMAIN AS FAR FROM THE DOOR AS POSSIBLE WITH
SEATBELTS SECURELY FASTENED.
If the CABIN DOOR Annunciator illuminates, or If an Unlatched Airstair/Cargo Door is Suspected:
1. All Occupants - SEATED WITH SEAT BELTS SECURELY FASTENED
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2. Cabin Sign - NO SMOKE & FSB
3. Cabin Differential Pressure - REDUCE TO LOWEST VALUE PRACTICAL (zero preferred) by descending and/or selecting higher cabin altitude setting.
4. Oxygen - AS REQUIRED
5. Land at nearest suitable airport.
EMERGENCY EXIT
Emergency Exit Handle - PULL (This is a plug-type door and opens into the cabin)
CAUTION
The outside handle may be locked from the inside with the EXIT LOCK lever. The inside EXIT- PULL handle will unlatch the door regardless of the position of the EXIT LOCK
lever. Before flight, make certain the lock lever is in the unlocked position. On some models, the outside handle may be locked from the inside with a key. The inside handle will
unlatch the door, regardless of the position of the key lock, by overriding the locking mechanism. Before flight, make certain the door is unlocked.
SPINS
If a Spin is entered inadvertently:
1. Control Column - FULL FORWARD
2. Full Rudder - OPPOSITE DIRECTION OF SPIN
3. Power Levers – IDLE
4. Controls - NEUTRALIZE WHEN ROTATION STOPS
5. Execute a smooth pull out.
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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.
EXPANDED GENERAL PROCEDURES
CABIN DOOR ANNUNCIATOR CIRCUITRY CHECK
The following test shall be performed prior to the first flight of the day.
1. Perform the following annunciator circuitry check:
a. Battery - ON
b. With door open and mechanism in locked position, ensure CABIN DOOR annunciator is ILLUMINATED. c. With door dosed and latched, but not locked, ensure the CABIN DOOR annunciator remains ILLUMINATED. d. With the door closed and locked, ensure that the CABIN DOOR annunciator is EXTINGUISHED. e. Battery - OFF
2. Ensure that the door is closed and locked using the following procedure:
a. Ensure that the door handle will not move out of the locked position without depressing the release button.
b. Lift the top door step and ensure that the red safety arm is around the plunger. Ensure that the green index mark on each of the 4 locking bolts aligns with the black pointer in the observation port.
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AIRPLANE – GENERAL
QUESTIONS
1) To open the emergency exit:
a) Turn the release handle clockwise and pull the door down and in.
b) Unlock the exit with the key and push the door out and away from the airplane.
c) Turn the release handle counterclockwise and push the door out.
d) Pull the door release handle downward and inward.
2) True or False: 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) True or False: 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 ________.
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9) List:
a) Va
b) Vne
c) Vlo
d) Vle
e) Vmc
10) The maximum zero fuel weight is ________.
11) True or False: The maximum ramp weight is 12,500lbs.
12) The maximum weight in the aft baggage compartment is ________.
13) What does the white triangle on the airspeed indicator represent?
14) What are the emergency procedures for an illuminated Door Light annunciator warning?
15) If the emergency exit has a key lock, can you remove the key if the door is locked?
16) If the emergency exit does not have a key lock, how do you ensure that it is locked?
17) Assuming the emergency exit is locked, can people enter the aircraft through it?
18) Assuming the emergency exit door is locked, can passengers exit the aircraft through the
emergency hatch?
19) True or False: The aircraft is approved for spins.
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CHAPTER 2
ELECTRICAL SYSTEM
OBJECTIVES
After completing this chapter, you will be able to:
1) Locate the switches for the:
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 for illumination and corrective action (if required) for the following:
a. Generator
b. Inverter (if required)
c. Battery charge
d. Ignition
4) Utilizing the aircraft electrical schematic locate:
a. Battery
b. Hot-wired bus
c. Generators
d. Current limiters
e. Generator busses
f. Dual fed busses
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g. Ground power plug
h. Inverters
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. Current limiter check
b. Normal engine start
7) State procedure for detecting:
a. A failed current limiter
b. A failed current limiter 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 Super King Air 200 electrical system is a 28-volt DC, "dual fed" 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 are located on the overhead panel and
indicate the load on each generator. The generator buses are interconnected by the isolation bus
through two 325-ampere current limiters. The current limiters will isolate the battery from a fault
on a generator bus. The current limiters should be checked prior to each flight. A reading of zero
on the left or right volt meter indicates that the current limiter is out on the side reading zero. The
entire bus system operates as a single bus, with power being supplied either by the battery or the
generators. There are four dual-fed sub-buses which receive power from either the left or right
generator bus after passing through a 60-amp limiter, a 70-amp diode, and a 50-amp circuit
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breaker. All aircraft electrical loads are divided among these buses. The equipment on the buses
is arranged so that all items with duplicate functions, such as right and left landing lights, do not
share a common bus. A dual inverter system is installed on the aircraft to provide AC power for
certain engine instruments and avionics equipment. The left generator bus powers the number 1
inverter and the right generator bus powers the number 2 inverter. The INVERTER selector
switch, located on the pilot's sub-panel, activates the selected inverter and provides 400-hertz,
115-volt, alternating current to the avionics equipment, and 400-hertz, 26 VAC to the
torquemeters. The battery is capable of starting the engines and can provide up to 30 minutes of
backup power in the event of a dual generator failure.
PILOT TIP
During the second engine start, turn off the operating engine’s generator. Attempting to start the second engine while the operating engine’s generator is energized will damage the 325A current
limiters. This procedure is not required on S/Ns BB 1444 and later.
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BATTERY SYSTEM
A fully charged 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, the battery should provide adequate power for approximately 30 minutes. The
battery's voltage can be checked by using the volt/load meters located on the overhead panel.
Pressing the knobs on both load meters checks the battery voltage and the condition of the
current limiters. No voltage indicates that a current limiter is out. 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 this 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
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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 for the duration of the flight.
PILOT TIP
The battery may be damaged if exposed to voltages higher than 30V for extended periods of time.
DC GENERATION - DESCRIPTION AND OPERATION
The major components of the DC generation and control system include the two starter-
generators and the battery. 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. In order to turn the generator ON, the generator switch must be held upward in the
reset position for one full 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.
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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 using the volt/load meter on the overhead
panel.
GENERATOR CONTROL UNIT
Aircraft BB-89 and subsequent
The generator control units (GCU) are self-contained components mounted below the center
aisle floor forward of the main spar. Each starter-generator has its own GCU to provide voltage
regulation, generator paralleling, reverse current sensing, and over-voltage and over-excitation
protection. During normal operation, each generator control unit 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. 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
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current relay is open. When the reverse current relay is closed, the annunciator will extinguish
and the volt/loadmeters should indicate starter-generator output.
Aircraft BB-2 through BB-89
On these aircraft a voltage regulator provides voltage regulation, generator paralleling, reverse
current sensing, and over-voltage and over-excitation protection. Each generator is equipped
with a voltage regulator that maintains a constant voltage output.
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 their 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 within 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 of regulator voltage are automatically
performed by the GCU's 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. Anytime one
starter-generator is on-line and the other is off-line at the same voltage, the paralleling circuit
will cause the regulators to decrease output voltage of the former and increase output voltage of
the latter, until both starter-generators are on-line.
PILOT TIP
During an engine start, ensure that the generator switch is in the OFF position. This prevents the generation of field current during engine start. The presence of field current during an engine
start will reduce the torque available from the starter and may lead to a hotter start.
OVER VOLTAGE PROTECTION
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The generator control units (GCU) 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 GCU will open the line
contactor relay and remove the starter-generator from the bus.
OVER EXCITATION PROTECTION
Over excitation protection is provided by the GCU. The GCU over excitation protection circuit
will activate in the event that starter-generator voltages begins to increase without control, but
does not go into over-voltage. If the generator field reaches its design limit, the generator will
drop off-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
GCU 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.
COMPONENT LOCATION
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The voltage regulators, 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.
AC GENERATION
AC power is supplied by one of two inverters installed in the wing center section
outboard of each engine nacelle. An inverter select switch, placarded INVERTER
NO 1, OFF, INVERTER NO 2 is located on the pilot's subpanel. When either
inverter is selected, DC power is supplied to that inverter and connects 26 VAC
and 115 VAC outputs to various instruments and systems requiring AC power. Typical avionics
that use AC power include the autopilot/flight director, RMI, attitude gyro and the ADF.
On aircraft BB-1095 and prior, the torquemeters are also AC powered. The inverter warning
annunciator light is energized anytime the inverter fails or power is removed. The warning light
on the King Air 200 reads INST INV while the warning light on the B200 reads INVERTER.
The AC meter is located on the overhead panel adjacent to the DC volt/load meters. The meter
normally monitors frequency, unless the button in the lower left hand corner of the meter is
pressed, at which time it will display voltage. For normal operation, the 115v inverter output
must be 107-120VAC at 390-410 Hz.
EXTERNAL POWER
The external power receptacle is located on the right
wing just outboard of the engine nacelle. The receptacle
is designed for use with an auxiliary ground unit having
a standard AN plug. A switch in the external power
plug receptacle illuminates a yellow caution light, EXT
PWR, on the caution/advisory annunciator panel. This
annunciator light receives power from the hot battery bus. A voltage of 24 to 28 VDC is required
to close the external power relay. The airplane electrical system is protected against damage
from reverse polarity by a relay and diode in the external power circuit. When an external power
source is used, the Ground Power Unit (GPU) must be capable of producing 1000 amperes for 5
seconds, 500 amperes for two minutes and 300 amperes continuously. Use of an inadequate
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ground power unit can cause damage to the airplane's electrical system. External power can be
used to operate all of the airplane’s electrical equipment, including the avionics.
PILOT TIP
The output setting must not be set to exceed 1000 amperes on ground power units. Any current set in excess of 1000 amperes may over-torque and damage the starter.
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. Only use a ground power source equipped with an AN-type plug.
b) Before connecting an external power unit, turn off all radio equipment and generator
switches, but turn the battery on to protect transistorized equipment against transient
voltage spikes.
c) If battery voltage indicates less than 20 volts, the battery must be recharged or replaced
with a battery indicating 20 volts or greater, before using auxiliary power. The battery
switch must be ON when starting engine with auxiliary power, and generators should be
OFF until auxiliary power has been disconnected.
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.
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PILOT TIP
Voltage is required to energize the avionics power relays in order to remove power from the avionics equipment.
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 panel is
located below the fuel gauges and to the left of the pilot. The larger 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. However, 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 and may result
in a fire. 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 "Emergency" section of your POH.
LIMITATIONS EXTERNAL POWER LIMITS
External power carts must 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
Maximum sustained generator load is limited as follows:
In Flight:
Sea Level to 31,000 feet altitude -100%
Above 31,000 feet altitude - 88%
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Ground - 85%
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
The pilot in command of an aircraft is directly responsible for and is the final authority as to the
operation of that aircraft. In an emergency requiring immediate action, the pilot in command may
deviate from any rule in 14 CFR Part 91, Subpart A, General, and Subpart B, Flight Rules, to the
extent required to meet that emergency. The following section deals with situations that require
immediate and accurate action by the crew. Memory items are printed in bold type and should be
completed in a timely manner. However, acting too rapidly may compound the emergency and
place the aircraft in an unrecoverable situation. To prevent this, memory items must be
accomplished methodically and must include coordination between the pilots.
The following steps should be considered mandatory in any emergency:
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1) Fly the airplane.
2) Identify the emergency.
3) Complete the appropriate checklist.
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 Ready - PULL ON (Verify) 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) 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.
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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.
INVERTER FAILURE
1) Select other inverter.
ABNORMAL ELECTRICAL PROCEDURES
GENERATOR INOPERATIVE (L or R DC GEN Annunciator)
1) Loadmeter - VERIFY GENERATOR IS OFF (0% LOAD) 2) Generator - RESET, THEN ON
If generator will not reset:
1) Generator – OFF 2) Loadmeter - DO NOT EXCEED 100% (88% Above 31,000 feet)
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BATTERY CHARGE RATE (BATTERY CHARGE Annunciator)
Ground Operations:
The BATTERY CHARGE annunciator will illuminate after an engine start. Do not take off with
the annunciator illuminated unless a decreasing battery charge current is confirmed. See Nickel-
Cadmium Battery Check in POH.
In Flight:
In-flight illumination of the BATTERY CHARGE annunciator indicates a possible battery
malfunction.
1) Battery Switch – OFF
2) BATTERY CHARGE Annunciator Extinguished - CONTINUE TO DESTINATION
BATTERY CHARGE Annunciator Still Illuminated - LAND AT NEAREST SUITABLE
AIRPORT.
EXCESSIVE LOADMETER INDICATION (over 100%)
1) Battery - OFF (monitor loadmeter)
If Loadmeter Still Indicates Above 100%:
1) Nonessential Electrical Equipment – OFF
If Loadmeter Indicates 100% or Below.
1) Battery – ON
CIRCUIT BREAKER TRIPPED
1) Nonessential Circuit - DO NOT RESET IN FLIGHT
2) Essential Circuit:
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a) Circuit Breaker - PUSH TO RESET
b) If Circuit Breaker Trips Again - DO NOT RESET
BUS FEEDER CIRCUIT BREAKER TRIPPED
(Fuel Panel Bus Feeders and Right Circuit Breaker Panel Bus Feeders)
- A short is indicated, do not reset in flight.
AVIONICS MASTER POWER SWITCH FAILURE
If the Avionics Master Power Switch Fails to Operate in the ON Position:
1) Avionics Master Circuit Breaker – PULL
PILOT TIP
Turning on the Avionics Master Power switch removes power that holds the avionics relay open. If the switch fails to the OFF position, pulling the Avionics Master circuit breaker will remove
power to the relay and should restore power to the avionics buses.
EXPANDED ELECTRICAL PROCEDURES
HOT BATTERY BUS CHECK WITH THE BATTERY SWITCH OFF.
1) Fuel Firewall Valves CLOSED
2) Standby Boost Pumps ON - Listen for operation.
3) Battery Switch ON -FUEL PRESS lights illuminate immediately.
4) Fuel Firewall Valves OPEN -FUEL PRESS lights extinguish.
5) Standby Boost Pumps OFF - FUEL PRESS lights illuminate.
CURRENT LIMITER CHECK
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1) One Generator TURN OFF EITHER LEFT OR RIGHT
2) Left and Right Volt/Loadmeters PRESS BOTH
3) 28 volts on both loadmeters NORMAL
4) Less than 28 volts on any loadmeter FAILED CURRENT LIMITER
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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-200?
a) The NiCad or lead-acid battery.
b) Ground power.
c) The two 250 amp starter-generators.
d) Both a & b above.
3) Why is the King Air 200 electrical system called "Dual Fed"?
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 200 has two __ volt and __ AMP D.C. starter -generators that are regulated to
__ volts ± .25 volts.
6) True or False: Certain engine instrument gauges use AC power.
7) What is the minimum the battery voltage for a battery start?
8) True or False: The starter-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 two 325 amp current limiters?
11) What are the 4 primary functions of the Generator Control unit?
12) What does the reverse current relay do?
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13) How many amps can the lead-acid battery provide for 1 hour?
a) 34
b) 42
c) 24
d) 12
14) True or False: 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-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 plug receptacle 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 starting engine's 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?
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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 seconds ON, 30
minutes OFF
b) 10 seconds ON, 30 seconds OFF, 40 seconds ON, 60 seconds OFF, 60 seconds ON, 90
seconds OFF
c) 20 seconds ON, 60 seconds OFF, 20 seconds ON, 60 seconds OFF, 20 seconds ON, 90
minutes OFF
d) 15 seconds ON, 50 seconds OFF, 15 seconds ON, 60 seconds OFF, 10 seconds ON, 5
minutes OFF
23) Explain how to check the current limiters.
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CHAPTER 3
ANNUNCIATOR SYSTEM
OBJECTIVES
After completing this chapter, you will be able to:
1) Identify the components of the annunciator system.
2) Describe the light 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 a red warning annunciator panel located in the center of the
glareshield, and a yellow caution and green advisory annunciator panel located on the center sub-
panel. Two red MASTER WARNING flashers are located in the glareshield in front of each
pilot. The two yellow MASTER CAUTION flashers are located just inboard of the MASTER
WARNING flashers and the PRESS TO TEST button is located immediately to the right of the
warning annunciator panel.
L ENG FIRE INVERTER CABIN DOOR ALT WARN R ENG FIRE L FUEL PRESS R FUEL PRESS L OIL PRESS L GEN OVHT A/P TRIM FAIL R GEN OVHT R OIL PRESS
L CHIP DETECT L BL AIR FAIL A/P DISC R BL AIR FAIL R CHIP DETECT
WARNING PANEL
L DC GEN HYD FLUID LOW PROP SYNC ON RVS NOT READY R DC GEN DUCT OVERTEMP
L ICE VANE BATTERY CHARGE EXT PWR R ICE VANE L AUTOFEATHER ELEC TRIM OFF AIR COND N1 LOW R AUTOFEATHER L ICE VANE EXT BRAKE DEICE ON LDG/TAXI LIGHT PASS OXY ON R ICE VANE EXT L IGNITION ON L BL AIR OFF FUEL CROSSFEED R BL AIR OFF R IGNITION ON
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CAUTION/ADVISORY PANEL
The annunciator lights are the word-readout type. Whenever a fault condition covered by the
annunciator system occurs, a signal is generated and the appropriate annunciator is illuminated.
If the fault requires the immediate attention and reaction of the pilot, the appropriate red warning
annunciator in the glareshield panel illuminates and both MASTER WARNING flashers begin
flashing. Any annunciator light illuminated on the warning panel will remain on until the fault is
corrected. However, the MASTER WARNING flashers can be extinguished by pushing the face
of either MASTER WARNING flasher, even if the fault is not corrected. This allows the
MASTER WARNING flashers to reset and be ready to displaying additional warnings. After the
fault that caused the warning to illuminate is corrected, the affected warning annunciator will
extinguish, but the MASTER WARNING flashers will continue flashing until one of them is
depressed. Whenever an annunciator-covered fault occurs that requires the pilot's attention but
not his immediate reaction, the appropriate yellow caution annunciator in the caution/ advisory
panel illuminates, and both MASTER CAUTION flashers begin flashing. The flashing
MASTER CAUTION lights can be extinguished by pressing the face of either of the flashing
lights to reset the circuit. This action resets the Master Caution panel and if another fault occurs
causing a caution annunciator light to illuminate, the MASTER CAUTION flashers will be
activated again. An illuminated caution annunciator on the caution/advisory annunciator panel
will remain on until the fault condition is corrected, at which time it will extinguish. However,
the MASTER CAUTION flashers will continue flashing until one of them is depressed. The
caution/advisory annunciator panel also contains the green advisory annunciators. There are no
master flashers associated with these annunciators, since they are only advisory in nature. They
indicate a functional situation that does not demand the immediate attention or reaction of the
pilot. An advisory annunciator can be extinguished only by correcting the condition indicated on
the illuminated lens.
All warning, caution, and advisory annunciator lights and the yellow MASTER CAUTION
flashers feature a "bright" and a "dim" mode of illumination intensity. The "dim" mode will be
selected automatically whenever all of the following conditions are met:
1) A generator is on-line.
2) The overhead flood lights are off.
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3) The pilot flight lights are on.
4) The ambient light level in the cockpit is below a preset value.
Unless all of these conditions are met, the "bright" mode will be selected automatically. On later
airplanes, and earlier airplanes with modified annunciator circuitry, The MASTER WARNING
flasher also features both a "bright" and "dim" mode of illumination. 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 right of the warning
annunciator panel in the glareshield, illuminates all the annunciator lights, MASTER
WARNING flashers, and MASTER CAUTION flashers. Any lamp that fails to illuminate when
tested should be replaced.
PILOT TIP
The annunciator light bulbs can be changed by pressing in the center of the indicator and removing it from the panel. Pull the bulb from the rear of the panel and replace it with a new
#327 bulb.
ANNUNCIATOR LIMITATIONS NONE
ANNUNCIATOR EMERGENCY PROCEDURES NONE
ANNUNCIATOR ABNORMAL PROCEDURES NONE
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ANNUNCIATOR SYSTEM
QUESTIONS 1) Name the three annunciator panels and the color of the lights associated with these panels.
2) The annunciator system features master warning and master caution flashers. Where are these located?
3) What would make them illuminate?
4) The annunciator panels will automatically dim when: (Circle correct answer)
a) The master light switch is: (On, Off)
b) The pilot's flight light switch is: (On, Off)
c) The overhead flood light switch is: (On, Off)
d) The cockpit light level is: (Low, High)
e) At least one generator is: (Off, On)
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CHAPTER 4
FUEL SYSTEM
OBJECTIVES
After completing this chapter, you will be able to:
1) Identify fuel system controls, components, functions and gauges.
2) Explain fuel annunciator lights, probable cause for illumination 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, and identify selected components.
9) Describe operation of fuel transfer system.
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 fuel system consists of a series of rubber-bladder cells and an integral wet wing tank in each
wing connected by a crossfeed line. The fuel system in each wing is further divided into a main
and auxiliary fuel system with a total usable fuel capacity of 544 gallons. The main fuel system
in each wing consists of a nacelle tank, two wing leading edge tanks, two box section bladder
tanks, and an integral wet wing tank. All the tanks are interconnected and fuel flows into the
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nacelle tank by gravity. The total usable fuel capacity of the main fuel system is 386 gallons.
The filler cap for this system of tanks is located on the leading edge of the wing 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 auxiliary fuel
system consists of a fuel tank on each side of the center section with a usable capacity of 79
gallons each. The auxiliary fuel system consists of a center section tank with its own filler
opening, and an automatic fuel transfer system to transfer the fuel into the main fuel system. Do
not put fuel in the auxiliary tanks unless the main tanks are full. If the auxiliary tanks are full,
fuel will be automatically used from these tanks prior to the wing tanks. During automatic
transfer of auxiliary fuel the nacelle tanks are constantly refilled by a jet transfer pump. A check
valve in the gravity feed line from the outboard wing prevents reverse fuel flow from the nacelle
tank back into the wing tank. Anytime the auxiliary fuel tanks are empty, fuel in the main wing
tank will gravity flow into the nacelle tanks. The main and auxiliary fuel systems are equipped
with five fuel sump drains, a drain manifold and a firewall filter drain in each wing. All fuel is
filtered with a firewall-mounted 20-micron filter. These filters incorporate an internal bypass
which opens to permit uninterrupted fuel supply to the engine in the event of filter icing or
blockage.
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FUEL SYSTEM SCHEMATIC
FUEL GAUGES
The fuel quantity indicator system is a
capacitance type system with one fuel gauge per
wing. A spring loaded selector allows the pilot
to switch from the main tank readout to the
auxiliary tank readout. 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 any other types of fuels are used, the system will not indicate correctly.
The gauges are marked in pounds.
FUEL DRAIN VALVES
The drain valve for the firewall fuel filter is located to the right of the filter at the firewall near
the bottom of the nacelle. The nacelle tank has two drains located on the bottom center of the
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nacelle forward of the wheel well. The inboard drain is for the standby boost pump and the
outboard drain is for the nacelle fuel sump and strainer. The leading edge tank has a drain on the
underside of the wing just outboard of the nacelle. The integral wet wing fuel tank has a sump
drain located approximately midway on the underside of the wing. The drain for the auxiliary
tank is at the wing root midway between the main and aft spars. The drains should be checked
for fuel contamination during each preflight.
PILOT TIP
Allow a 3 hour settle period whenever possible after fueling before checking for contamination.
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 just outboard of the nacelle. A NACA vent is installed and recessed
to prevent icing. The second vent is electrically heated to prevent icing and serves as a backup
should the NACA vent become plugged.
FUEL PUMPS
The wing tanks gravity feed into the nacelle tank through a fuel line. A flapper-type check valve
in the end of the gravity feed line prevents any flow of fuel back into the wing tanks. Fuel is
pumped to the engine by the engine-driven low pressure boost pump mounted on the accessory
section of the engine. The low pressure pump operates any time the gas generator (N1) is turning
and provides fuel pressure to the high pressure engine driven fuel pump. The low pressure pumps
put out sufficient fuel pressure for all conditions except operation in the crossfeed mode or while
using aviation gasoline at altitudes above 20,000 feet. The purpose of this pump is to provide
pressurized fuel to the high pressure engine driven fuel pump. The low pressure pump provides
lubrication and prevents cavitation of the high pressure fuel pump. It is not an emergency backup
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, 200-mesh strainer. This pump provides sufficient fuel pressure to ensure a proper
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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,
the engine-driven low pressure boost pump is backed up by an electrically driven standby fuel
pump located in the bottom of each nacelle tank. In addition to serving as a backup unit in the
event of a malfunction in the engine-driven low pressure boost pump, the electrically driven
standby pump provides the pressure required for crossfeed operations. Failure of the engine
driven low pressure 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 illuminates the red FUEL PRESSURE warning light in the annunciator panel. If this
occurs, the standby boost pump should be turned on. The red FUEL PRESSURE light will
extinguish at approximately 11 psi as fuel pressure increases.
CAUTION
OPERATION WITH THE FUEL PRESSURE LIGHT ON IS LIMITED TO 10 HOURS BETWEEN OVERHAUL OR REPLACEMENT OF THE ENGINE-DRIVEN FUEL
PUMP.
The standby pumps are controlled by toggle switches on the fuel-control panel. The power
source for the standby boost pumps is supplied from the number 3 and number 4 dual fed buses.
This power is available only when the master switch is turned on. The alternative source of
power to the standby 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 standby pump switches
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must be in the off position to prevent discharge of the battery.
PILOT TIP
Remember to check that the fuel crossfeed switch and both standby boost pump switches are turned off after shutdown. These items are powered by the hot battery bus and will discharge the
battery if left on.
AUXILIARY FUEL TRANSFER SYSTEM
Fuel pressure from the engine-driven low pressure boost pump provides the motive flow to
operate the jet transfer pump. The jet pump transfers fuel from the auxiliary tanks to the nacelle
tanks. The transfer jet pumps are actuated by toggle switches on the fuel-control panel. This
switch selects either the automatic (AUTO) or manual (AUX TRANSFER OVERRIDE)
position. When the switch is placed in the AUTO position, the motive flow valve will open
approximately 30 to 50 seconds after the engine starts. This time delay prevents the loss of fuel
pressure during engine starting. During auxiliary fuel transfer, a pressure switch located in the
fuel line is set to actuate between 5 to 7 psi. If the fuel pressure in this line does not increase, the
NO TRANSFER light on the fuel-control panel will illuminate indicating that the motive flow
valve is still closed and fuel is not transferring from the auxiliary tank. If this occurs, select the
AUX TRANSFER OVERRIDE position using the auxiliary fuel transfer switch. This action will
bypass the automatic fuel transfer feature and apply power directly to the motive flow valve.
Once the motive flow valve has opened, the jet transfer pump will pump fuel from the auxiliary
fuel tank into the nacelle fuel tank as long as either the engine-driven boost pump or the
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electrical standby boost pump is operating and there is fuel in the auxiliary tank. An overflow
line returns excess fuel delivered by the jet transfer pump back to the auxiliary tank. When the
auxiliary fuel tank is empty, a low-level float switch closes the motive flow valve after a 30- to
60-second time delay. This delay prevents cycling of the motive flow valve which could be
caused by sloshing fuel. The automatic fuel-control module simultaneously removes the power
to close the motive flow valve to prevent continued operation of the jet transfer pump. The
auxiliary fuel system will not feed fuel into the main fuel system if there is a simultaneous failure
of the engine driven low pressure boost pump and the electrically driven standby pump on the
same side or if there is a failure of the motive flow valve. This condition will cause the
illumination of the NO TRANSFER light on the fuel-control panel. The firewall shutoff valve
for each engine fuel system is actuated by its respective FUEL FIRE- WALL VALVE switch on
the pilot's fuel-control panel.
When the FUEL FIREWALL VALVE switch is closed, its respective firewall shutoff valve
shuts off the flow of fuel to the engine. The firewall shut off valves receive power from the
number 3 and number 4 dual fed buses. This power is available only when the master switch is
turned on. The alternative source of power for the firewall shutoff valves is directly from the
battery through the hot battery bus. Only fuel is cut off to the engine with this switch.
FUEL FILTERS
From the firewall shutoff valve, fuel is routed to the engine-driven boost pump and then to 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. There is no indication in the
cockpit if the fuel filter is being bypassed. 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.
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PILOT TIP
The normal interval for inspecting all fuel filters is 150 hours.
FUEL HEATER
Dissolved water cannot be filtered from the fuel with micronic type filters, but can be released by
lowering the fuel temperature. Since this can occur during flight, a fuel heater is installed on each
engine. 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. Fuel heater
operation is automatic whenever the engine is running and requires no pilot action.
CROSSFEED
Crossfeed is only to be conducted during single engine operations. Each nacelle tank is
connected to the opposite engine by a crossfeed line. Crossfeed operation is controlled by a
manually operated crossfeed switch on the fuel-control panel. This switch energizes a solenoid
that opens the crossfeed valve. This action simultaneously energizes the standby pump on the
side from which fuel is desired and de-energizes the motive flow valve in the opposite fuel tank
system. When the crossfeed valve is open, the green FUEL CROSSFEED light on the
annunciator panel will illuminate. The crossfeed does not transfer fuel from tank to tank. Its
primary function is to supply fuel from one side to the opposite engine during an engine-out
condition. If the standby boost pumps on both sides are operating and the crossfeed valve is
open, fuel will be supplied to the engines in the normal manner because the pressure on each side
of the crossfeed valve will be equal.
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CAUTION
THE STANDBY BOOST PUMP MUST BE OPERATIONAL ON THE SIDE FROM WHICH THE FUEL IS BEING SUPPLIED.
FUEL PURGE SYSTEM
The fuel system on airplane serials BB-2 through BB-665 is equipped with a fuel drain collector
system. Airplane serials BB-666 and after are equipped with a fuel purge system. The fuel purge
system is designed to burn any residual fuel in the fuel manifolds during engine shutdown.
During engine operation, compressor discharge air (P3 air) is routed through a filter and check
valve, pressurizing a small air tank mounted on the engine. During engine shutdown, the
pressure differential between the air tank and fuel manifold causes air to be discharged into the
fuel manifold system. This air forces all residual fuel out through the nozzles and into the
combustion chamber where it is consumed. This action causes a momentary rise in engine speed.
FUEL SYSTEM LIMITATIONS
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
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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 20,000 feet pressure altitude (FL 200) or below if either standby pump
is inoperative.
3) Crossfeed capability is required for climbs above 20,000 feet pressure altitude (FL 200).
4) Operation above 31,000 feet (FL 310) is prohibited.
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.
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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
Water in jet fuel creates an environment favorable to the growth of microbiological sludge in the
settlement areas of the fuel cells. This sludge, plus other contaminants in the fuel, can cause
corrosion of metal parts in the fuel system as well as clogging of the fuel filters. 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 200 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 Usable Fuel Quantity 544 gallons (3645 pounds)
• Each Main Fuel Tank System 193 gallons (1293 pounds)
• Each Auxiliary Fuel Tank 79 gallons (529 pounds)
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FUEL IMBALANCE
Maximum allowable fuel imbalance between wing fuel systems is 1000 pounds.
FUEL CROSSFEED
Crossfeeding of fuel is permitted only when one engine is inoperative.
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.
AUXILIARY FUEL
Do not put any fuel into the auxiliary tanks unless the main tanks are full.
OPERATING WITH LOW FUEL PRESSURE
Operation of either engine with its corresponding fuel pressure annunciator (L FUEL PRESS or
R FUEL PRESS) illuminated is limited to 10 hours before overhaul or replacement of the
engine-driven fuel pump. Windmilling time need not be charged against this time limit.
WARNING
ALTHOUGH THE AIRPLANE IS APPROVED FOR TAKEOFF WITH ONE STANDBY BOOST PUMP INOPERATIVE, CROSSFEEDING OF FUEL WILL NOT BE
AVAILABLE FROM THE SIDE OF THE INOPERATIVE STANDBY BOOST PUMP.
EMERGENCY FUEL SYSTEM PROCEDURES The pilot in command of an aircraft is directly responsible for and is the final authority as to the
operation of that aircraft. In an emergency requiring immediate action, the pilot in command may
deviate from any rule in 14 CFR Part 91, Subpart A, General, and Subpart B, Flight Rules, to the
extent required to meet that emergency. The following section deals with situations that require
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immediate and accurate action by the crew. Memory items are printed in bold type and should be
completed in a timely manner. However, acting too rapidly may compound the emergency and
place the aircraft in an unrecoverable situation. To prevent this, memory items must be
accomplished methodically and must include coordination between the pilots.
The following steps should be committed to memory and considered mandatory in any
emergency:
1) Fly the airplane.
2) Identify the emergency.
3) Complete the appropriate checklist.
BOLD TYPE INDICATES MEMORY ITEMS!
FUEL PRESSURE LOW [L FUEL PRESS] OR [R FUEL PRESS]
1) Standby Pump (failed side) ON
2) [FUEL PRESS] EXTINGUISHED
3) Oil Temperature and Pressure Gages (failed side) MONITOR
ABNORMAL FUEL PROCEDURES CROSSFEED (One-Engine-Inoperative Operation)
1) Crossfeed LEFT OR RIGHT, AS REQUIRED [FUEL CROSSFEED] - ILLUMINATED
2) Standby Pumps OFF
3) Auxiliary Tank Transfer AUTO
4) Fuel Balance MONITOR
If Fuel is Required from the Inoperative Engine's Auxiliary Fuel Tank and the Reason for
Shutdown was Not an Engine Fire or Fuel Leak:
1) Firewall Shutoff Valve (inoperative engine) OPEN [FUEL PRESS] - EXTINGUISHED
2) No Transfer Light (inoperative engine) EXTINGUISHED IN 30 TO 50 SECONDS
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To Discontinue Crossfeed:
1) Crossfeed Flow Switch OFF (centered)
AUXILIARY FUEL TRANSFER FAILURE (NO TRANSFER Light)
1) Auxiliary Tank Transfer OVERRIDE
2) No Transfer Light EXTINGUISHED (If light does not extinguish, auxiliary fuel may not be
available.)
3) Auxiliary Fuel Quantity MONITOR
4) Auxiliary Tank Transfer AUTO (when auxiliary fuel tank is empty)
EXPANDED FUEL PROCEDURES
FUEL SYSTEM CHECK
Conduct the following checks with Battery ON:
1) Firewall Shutoff Valves – CLOSE
2) Standby Pumps – ON Listen For Operation, Verify both FUEL PRESS lights Illuminated
3) Firewall Shutoff Valves - OPEN Verify both FUEL PRESS lights extinguished
4) Standby Pumps - OFF Verify both FUEL PRESS lights Illuminated
5) Crossfeed – LEFT, then RIGHT while Verifying FUEL CROSSFEED light illuminates and
FUEL PRESSURE lights extinguish.
6) Crossfeed – OFF
7) Auxiliary Tank Transfer – AUTO
8) No Transfer Light - PRESS TO TEST
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FUEL SYSTEM
QUESTIONS
1) List the items on the fuel panel that receive power from the hot battery bus:
2) True or False: The engine will continue to operate at reduced power with boost pump pressure
after the failure of the high pressure fuel pump.
3) True or False: The jet pump is DC powered from the number 2 Dual Fed bus.
4) Maximum fuel imbalance 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 standby boost pump?
a) It functions as a backup pump in the event of primary boost pump failure.
b) It is used with aviation gas in climbs above 20,000 feet.
c) It is used in crossfeed operation.
d) All of the above.
7) Total fuel capacity: __ gallons __lbs.
Main Tanks: __ gallons __lbs.
Aux Tanks: __ gallons __lbs.
8) When is crossfeed use authorized?
a) For single-engine operation.
b) For climbs above 20,000 feet when aviation gas is used.
c) When one standby pump is inoperative.
d) When fuel pressure decreases below 10 ± psi.
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9) Maximum Zero Fuel weight is _____lbs.
10) Which of the following limitations applies to operation with aviation gas?
a) A maximum altitude of 20,000 feet with both standby boost pumps operative and 150 hours
between overhauls
b) A maximum altitude of 31,000 feet with standby boost pump inoperative and 150 hours
between overhauls
c) A maximum altitude of 20,000 feet with one standby pump inoperative and 150 hours
between overhauls
d) A maximum of 150 hours between overhauls only
11) Is a fuel anti-icing additive required for this aircraft?
12) Illumination of the fuel pressure warning light indicates:
13) True or False: The engine will continue to operate at reduced power with boost pump
pressure after the failure of the high pressure fuel pump.
14) True or False: The “NO TRANSFER” light will come on for 30-50 seconds after the
auxiliary fuel is completely transferred to the main system.
15) You fuel the airplane with jet fuel and mix in 100 gallons of AVGAS. Each engine must be
charged ____ hour(s) against its 150 hour AVGAS limitation.
16) When selecting crossfeed, left to right, the automatic fuel transfer module will do what to the
following items?
a) Right electric boost pump
b) Left electric boost pump
c) Right motive flow valve
17) What are the memory items for illumination of a Fuel Pressure Low annunciator light?
18) How long should you let the fuel settle before checking for contaminates?
a) 1 hour
b) 2 hours
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c) 3 hours
d) 4 hours
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CHAPTER 5
ENGINE SYSTEM
OBJECTIVES
After completing this chapter, you will 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 200 was introduced with Pratt & Whitney PT6A-41 engines. The -41 is flat rated
to 850 SHP at 2000 RPMs. The B200 is equipped with the -42 engine. This engine is identical to
the -41 but incorporates improvements in the first stage axial flow compressor and internal
changes to the exhaust duct. This allows a 10% increase in altitude cruise performance. The Pratt
& Whitney PT6A engine is a light weight, reverse flow, free turbine engine driving a propeller
via 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
two power turbines and the power turbine shaft. The two rotors 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
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gearbox starts the engine. The compressor draws air into the engine via an annular air inlet case,
increases its pressure across the 3 axial stages and one centrifugal impeller and delivers it around
the combustion chamber. Air enters the combustion chamber via small holes and, at the correct
compressor speed, fuel is introduced into the combustion chamber. Two spark igniters located in
the combustion chamber ignite the mixture. The hot gases are then 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 vane
ring and hit the turbine blades and create a rotational movement of the compressor turbine to
drive the compressor. The expanding gases travel across the power turbines and provide
rotational energy to drive the propeller shaft. The reduction gearbox reduces the power turbines
speed (approximately 30,000 RPM) to one suitable for propeller operation (1600 to 2000 RPM).
This is done through a 15 to 1 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 hydromechanical 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.
PROPULSION SYSTEM CONTROLS
The propulsion system is operated by three sets of controls:
1) The power levers
2) The propeller levers
3) The condition levers
The power levers control engine power from idle through take-off power by operation of the gas
generator (N1) governor in the fuel control unit. Increasing N1 rpm results in increased engine
power. The condition levers have three positions; FUEL CUT-OFF, LOW IDLE and HIGH
IDLE. Each lever controls the fuel cutoff function of the fuel control unit and limits idle speed at
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56-62% N1 for low idle, and 70% N1 for high idle. The propeller levers are operated
conventionally and control the constant speed propellers through the primary governor.
PILOT TIP
If excessive ITT's occur during any one of the following conditions, adjust the condition levers to a higher N1 speed.
• When high generator loads are required. • During operations at high ambient air temperatures.
• During operations at high field elevations. • When maximum reverse is required.
:
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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
Ng (or N2) Gas generator speed (RPM or %) Nf (or N2) Power turbine speed (RPM or %) Np Propeller speed (rp 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 long with oil temperature, oil pressure and fuel flow.
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The engines used on the King Air 200 have seven major sections: 1) Air intake section, 2)
Compressor section, 3) Combustion section, 4) Turbine section, 5) Exhaust section, 6)
Reduction gear section, 7) Accessory drive section.
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 via 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
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stage and prior to the air entering the combustion can. This air is referred to as P3 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. Compressor bleed valves compensate 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
valves, one on each side of the compressor located at the 9 o'clock and 3 o'clock position of the
engine, are pneumatic pistons which reference the pressure differential between the axial and
centrifugal stages. The function of these valves is to prevent compressor stalls and surges in the
low N1 operating range. At low N1 RPM, both valves are in the open position. At takeoff and
cruise N1 RPM both bleed valves will be closed. If both compressor bleed valves were to stay
closed, a compressor stall would result from the attempt to accelerate the engine to takeoff
power. If one or both valves were to stick in the open position, the ITT would increase, the
torque decrease, while N1 RPM would remain the same.
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PILOT TIP
• Throttle back if a continuous compressor surge is encountered.
• Accelerate slowly if an engine is prone to surging.
• A surge may damage the compressor and hot section. Have the engine bleed valve checked if surging is encountered.
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 turbines and propellers 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 three reaction turbines. The two-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 section of the engine 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.
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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
6) Freon compressor on the right engine only
7) Low pressure fuel boost pump
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
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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.
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PILOT TIP
The dipstick indicates one quart below full when the oil level is normal. Minimum oil quantity for operations is four quarts low. Overfilling may cause a discharge of oil through the breather until
a satisfactory level is reached. 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.
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
gas generator speeds (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 60 – 200 Psig by a pressure relief valve in
the engine. Actual range on each model is dependent upon the aircraft serial number.
OIL FILTER
The engine oil filter is located under the square cover plate at the three-o'clock position of the
compressor inlet case and just behind the aft fire seal. The filter element should be replaced after
1000 hours of use and inspected for cleanliness and condition at 150-hour intervals. This filter
element is not cleanable and must be replaced if it has been subjected to heavy contamination
from the engine oil system.
OIL COOLER
The oil cooler radiator is located inside the lower engine nacelle. 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.
PILOT TIP
The engine ice vanes should be extended for all ground operations to minimize ingestion of ground debris. Turn engine anti-ice off, when required, to maintain oil temperature within limits.
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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.
PILOT TIP
Maximum oil consumption is 1 quart every 10 hours.
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.
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
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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 (90°F), the core valve is completely closed and oil bypasses the heater core.
ENGINE FUEL SYSTEM
The engine fuel system consists of the engine driven low pressure fuel pump, an oil to fuel
heater, the high pressure engine driven fuel pump, the fuel control unit (FCU), the flow divider
which sends fuel to the 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 modifying 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
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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.
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PILOT TIP
After engine start, the generator will not come on line if the starter switch has been left in the start 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 (BB-2 through BB-1438)
The 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
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
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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.
FIRE DETECTION SYSTEM (BB-1439 AND AFTER)
The fire detection system on these airplanes is designed to provide an immediate warning in the
event of a fire or overtemperature condition in either engine compartment. The main component
of the system is a temperature sensing element, which is routed through the three sections of
each engine nacelle and terminated in a responder unit. The responder unit is attached to the
engine mount in each engine accessory section at approximately the two o'clock position just
forward of the engine firewall. The responder unit contains two sets of contacts: a set of integrity
switch contacts for continuity test functions of the fire detection circuitry and a set of alarm
switch contacts which completes the circuit to actuate the fire warning system when the sensor
element detects an overtemperature condition in critical areas of the engine compartment. The
signals sent to the left or right annunciator-fault-detection printed circuit cards will illuminate the
respective red L or R ENG FIRE warning annunciator in the warning annunciator panel located
on the center glareshield. The left and right annunciator-fault-detection printed circuit cards will
also trigger the annunciator-control-circuit which will illuminate the pilot's and co-pilot's red
MASTER WARNING lights located in the glareshield. If the optional fire extinguishing system
is installed, the fire extinguisher control switches will illuminate. The MASTER WARNING
lights will continue to flash, even if the fire is extinguished. The MASTER WARNING lights
may be turned off by depressing the legend face of either light. At this time, the MASTER
WARNING lights will remain extinguished, even if a fire still burns inside the engine
compartment. The MASTER WARNING lights will automatically begin to flash again anytime
an additional warning annunciator is illuminated.
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The red L or R ENG FIRE warning annunciator is illuminated when the respective fire detection
element senses an overtemperature condition of sufficient magnitude to activate the alarm switch
contacts of the responder unit.
The red L or R ENG FIRE warning annunciator will automatically extinguish after the sensor
element in the engine compartment cools. The sensor element consists of a sealed outer tube
filled with an inert gas and an inner core filled with an active gas. The gases within the tubes
form a pressure barrier that keeps the contacts of the responder integrity switch closed for
continuity test functions of the fire alarm. As the temperature around the sensor element
increases, the gases within the tube begin to expand. If the pressure from the expanding gases
reaches a preset point, the contacts of the responder alarm switch close, illuminating the
respective red L ENG FIRE or R ENG FIRE warning annunciator and flashing the MASTER
WARNING lights.
The integrity (fault) pressure switch operates in the reverse manner of the alarm pressure switch.
The calibration gas (helium) sealed inside the sensor element normally holds the integrity
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pressure switch in a closed position, but allows the switch to open when the outer portion of the
sensor element is severed. Therefore, if the fire detection system is tested with the integrity
pressure switch open, the unit would fail to test, indicating a fault in continuity.
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).
The fire detection system is designed to actuate the alarm when any of the following conditions occur:
• When any one-foot section of the sensor element is heated to 900°F.
• When the average temperature of the entire sensor element reaches 450°F.
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
72° F. Four spray nozzles are positioned under the engine exhaust area, with another pair
mounted 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
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nozzles. The power to the switches is derived from the hot battery bus. These switches
incorporate three indicator lights. Airplanes BB-2 through BB-1485 are colored and marked as
follows: The red light, placarded L or R ENG FIRE-PUSH TO EXT, warns of the presence of
fire in the engine compartment. The amber light, placarded D, indicates that the system has been
discharged and the cartridge is empty. The green light, placarded OK, is provided only for the
preflight test function. Airplanes BB-1484, and after, are colored and marked as follows: A
yellow light, placarded EXTINGUISHER PUSH, warns of the presence of fire in the engine
compartment. A yellow light, placarded DISCH, indicates that the system has been discharged
and the cartridge is empty. A green light, placarded OK, is provided only for the preflight test
function. To actuate the system, raise the safety-wired clear plastic switch cover and press the
face of the lens. When the system is depleted, the amber or yellow D light will illuminate and
remain illuminated, regardless of the battery switch position, until the depleted extinguisher
cartridge has been replaced. The fire extinguisher circuits should be checked during the preflight
inspections by rotating the test switch through the L and R EXT positions on the switch. The
amber or yellow D and green OK lights on the extinguisher switches should illuminate. The
pressure gage mounted on each extinguisher supply cylinder should be checked during the
preflight inspection to assure that each cylinder is fully charged.
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ENGINE SYSTEM LIMITATIONS
NUMBER OF ENGINES Two ENGINE MANUFACTURER Pratt & Whitney Canada (Longueuil, Quebec, Canada) ENGINE MODEL NUMBER PT6A-42 POWER LEVERS Do not lift power levers in flight. STARTER LIMITS 40 seconds on, 60 seconds off; 40 seconds on, 60 seconds off; 40 seconds on, 30 minutes off. APPROVED ENGINE OILS The following oils are fully approved for use in Pratt &Whitney Canada PT6A-41 and -42
engines. Always refer to the latest revision of P&WC SB 3001 for a current list of approved
oils.
• Aeroshell Turbine Oil 500
• Aeroshell Turbine Oil 560
• Castrol 205
• Exxon Turbo Oil 2380
• Mobil Jet Oil 254
• Mobil Jet Oil II
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Do not mix different oil brands together.
PT6A-42
ENGINE OPERATING LIMITS
The following limitations shall be observed. Each column presents limitations. The limits presented do not necessarily occur simultaneously.
FOOTNOTES:
1) Torque limit applies within range of 1600 - 2000 propeller RPM (N2). Below 1600
propeller RPM, torque is limited to 1100 ft-lbs.
2) When gas generator speeds are above 27,000 RPM (72% N1) and oil temperatures are between 60°C and 71°C, normal oil pressures are:
100 to 135 psi below 21,000 feet; 85 to 135 psi at 21,000 feet and above.
During extremely cold starts, oil pressure may reach 200 psi. Oil pressure between 60
and 85 psi is undesirable; it should be tolerated only for the completion of the flight,
and then only at a reduced power setting not exceeding 1100 ft-lbs torque. Oil pressure
below 60 psi is un- safe; it requires that either the engine be shut down, or that a
landing be made at the nearest suitable airport, using the minimum power required to
sustain flight. Fluctuations of plus or minus 10 psi are acceptable.
3) A minimum oil temperature of 55°C is recommended for fuel heater operation at take-off power.
4) Oil temperature limits are -40°C and 99°C. However, temperatures of up to 104°C are
permitted for a maximum time of 10 minutes.
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5) These values are time limited to 5 seconds.
6) High ITT at ground idle may be corrected by reducing accessory load and/or
increasing N1 RPM.
7) At approximately 70% N1.
8) Cruise torque values vary with altitude and temperature.
9) This operation is time limited to 1 minute.
10) These values are time limited to 10 seconds.
11) Values above 99°C are time limited to 10 minutes.
PT6A-41
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EMERGENCY ENGINE SYSTEM PROCEDURES The pilot in command of an aircraft is directly responsible for and is the final authority as to the
operation of that aircraft. In an emergency requiring immediate action, the pilot in command may
deviate from any rule in 14 CFR Part 91, Subpart A, General, and Subpart B, Flight Rules, to the
extent required to meet that emergency. The following section deals with situations that require
immediate and accurate action by the crew. Memory items are printed in bold type and should be
completed in a timely manner. However, acting too rapidly may compound the emergency and
place the aircraft in an unrecoverable situation. To prevent this, memory items must be
accomplished methodically and must include coordination between the pilots.
The following steps should be committed to memory and considered mandatory in any
emergency:
1) Fly the airplane.
2) Identify the emergency.
3) Complete the appropriate checklist.
BOLD TYPE INDICATES MEMORY ITEMS!
All airspeeds quoted in this section are indicated airspeeds (IAS) and assume zero instrument error.
EMERGENCY AIRSPEEDS (12,500 LBS)
One-Engine inoperative Best Angle-of-Climb (VXSE) 115 kts.
One-Engine inoperative Best Rate-of-Climb (VySE) 121 kts.
Air Minimum Control Speed (VmcA) 86 kts.
Emergency Descent 181 kts
Maximum Range Glide 135 kts
ENGINE FAILURE
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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
Proceed with the Emergency Engine Shutdown for the following situations:
• ENGINE TORQUE INCREASE - UNSCHEDULED
• ENGINE FIRE IN FLIGHT
• ENGINE FAILURE IN FLIGHT
Affected Engine:
1) Condition Lever - FUEL CUT OFF
2) Propeller Lever - FEATHER
3) Firewall Shutoff Valve - CLOSED
4) Fire Extinguisher (if installed) - ACTUATE (if required)
5) Auto Ignition - OFF
6) Generator - OFF
7) Prop Sync - OFF
8) Electrical Load - MONITOR
ENGINE FIRE ON GROUND
Affected Engine:
1) Condition Lever - FUEL CUT OFF
2) Firewall Shutoff Valve - CLOSED
3) Ignition and Engine Start - STARTER ONLY
4) Fire Extinguisher (if installed) - ACTUATE (if required)
ENGINE FAILURE DURING GROUND ROLL
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1) Power Levers – IDLE
2) Brakes - AS REQUIRED
3) Operative Engine - MAXIMUM REVERSE
WARNING
EXTREME CARE MUST BE EXERCISED WHEN USING SINGLE-ENGINE REVERSING ON SURFACES WITH REDUCED TRACTION.
If Insufficient Runway Remains for Stopping:
4) Condition Levers - FUEL CUT OFF
5) Firewall Shutoff Valves - CLOSED
6) Master Switch - OFF (Gang bar down)
ENGINE FAILURE AFTER LIFT-OFF
1) Power - MAXIMUM ALLOWABLE
2) Airspeed - MAINTAIN (take-off speed or above)
3) Landing Gear - UP
NOTE
If the autofeather system (if installed) 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.
7) Propeller Lever (inoperative engine) - FEATHER (or verify FEATHER if autofeather is installed)
8) Airspeed- VYSE (after obstacle clearance altitude is reached)
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9) Flaps - UP
10) Clean-up (inoperative engine): a) Condition Lever - FUEL CUT OFF
b) Propeller Lever - FEATHER
c) Firewall Shutoff Valve - CLOSED
d) Auto Ignition - OFF
e) Autofeather (if installed) - OFF
f) Generator – OFF
11) Electrical Load - MONITOR
ENGINE FAILURE IN FLIGHT BELOW AIR MINIMUM CONTROL SPEED
1) Power - Reduce as required to maintain directional control.
2) Nose - Lower to accelerate above VMCA.
3) Power (operative engine) - AS REQUIRED.
4) Failed Engine - SECURE (See 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.
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ENGINE OUT GLIDE
1) Landing Gear – UP
2) Flaps - UP
3) Propellers - FEATHERED
4) Airspeed - 135 KNOTS
WARNING
DETERMINE THAT PROCEDURES FOR RE-STARTING FIRST AND SECOND FAILED ENGINES ARE INEFFECTIVE BEFORE FEATHERING SECOND ENGINE
PROPELLER.
PILOT TIP
The Glide Ratio is 2.0 nm for each 1000 feet of altitude.
ABNORMAL ENGINE SYSTEM PROCEDURES
AIR START
WARNING
AIRSTART USING THE STARTER ASSIST PROCEDURES MAY MOMENTARILY CAUSE THE LOSS OF ATTITUDE DISPLAY ON ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) EQUIPPED AIRPLANES, AND LEAD TO
PREMATURE SYSTEM FAILURES. IF FLIGHT CONDITIONS DO NOT PERMIT THE TEMPORARY LOSS OF ATTITUDE REFERENCE, CONDUCT AIRSTART
USING THE NO STARTER ASSIST PROCEDURES.
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CAUTION
THE PILOT SHOULD DETERMINE THE REASON FOR ENGINE FAILURE BEFORE ATTEMPTING AN AIR START. DO NOT ATTEMPT AN AIR START IF N1
INDICATES ZERO. 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 OVERTEMPERATURE CONDITION.
STARTER ASSIST
1) Cabin Temp Mode - OFF
2) Vent Blower - AUTO
3) Aft Blower (if installed) - OFF 4) Radar - STANDBY or OFF
5) Windshield Heat - OFF
6) Power Lever - IDLE
7) Propeller Lever (inoperative engine) - LOW RPM
8) Condition Lever - FUEL CUT OFF
9) Firewall Shutoff Valve - OPEN
10) Generator (inoperative engine) – OFF
NOTE
If conditions permit, retard operative engine ITT to 700°C or less to reduce the possibility of
exceeding ITT limit. Reduce electrical load to minimum consistent with flight conditions.
1) Ignition and Engine Start - ON, IGNITION ON annunciator - ILLUMINATED
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2) Condition Lever - LOW IDLE
3) ITT and N1 - MONITOR (1000°C MAXIMUM)
4) Ignition and Engine Start - OFF (N1 above 50%)
5) Propeller Lever - AS REQUIRED
6) Power Lever - AS REQUIRED
7) Generator - ON
8) Auto Ignition - ARM
9) Prop Sync – ON
10) Cabin Temp Mode – AUTO
NO STARTER ASSIST (Windmilling Engine and Propeller)
1) Power Lever - IDLE
2) Propeller Lever - FULL FORWARD
3) Condition Lever - FUEL CUT OFF 4) Engine Ice Vane (inoperative engine) RETRACTED
5) Firewall Shutoff Valve - OPEN
6) Generator (inoperative engine) - OFF
7) Airspeed - 140 KNOTS MINIMUM
8) Altitude - BELOW 20,000 FEET
9) Auto Ignition - ARM (IGNITION ON annunciator - ILLUMINATED)
10) Condition Lever - LOW IDLE
11) ITT and N1 - MONITOR (1000°C MAXIMUM)
12) Power - AS REQUIRED (after ITT has peaked)
13) Generator – ON
14) Prop Sync – ON
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ONE-ENGINE-INOPERATIVE APPROACH AND LANDING
1) Approach Speed - CONFIRM
2) Fuel Balance - CHECK
3) Pressurization - CHECK
4) Cabin Sign - NO SMOKE & FSB
When it is certain that the field can be reached:
5) Flaps - APPROACH
6) Landing Gear - DN
7) Propeller Lever - FULL FORWARD
8) Airspeed - 10 KNOTS ABOVE NORMAL LANDING APPROACH SPEED
9) Interior and Exterior Lights - AS REQUIRED
10) Radar - AS REQUIRED
11) Surface Deice - CYCLE (as required)
When it is certain there is no possibility of a Go-Around:
1) Flaps - DN
2) Airspeed - NORMAL LANDING APPROACH SPEED
3) Perform 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) Landing Gear - UP
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3) Flaps – UP
4) Airspeed - INCREASE TO BLUE LINE
LOW OIL PRESSURE INDICATION
Oil pressure values between 60 and 85 psi are undesirable and should only be tolerated for the
completion of the flight. In this situation, the engine should be operated at a reduced power
setting not exceeding 1100 foot-pounds torque. Oil pressure values below 60 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.
CHIP DETECT (L or R CHIP DETECT Annunciator)
Illumination of a CHIP DETECT annunciator indicates possible metal contamination in the
engine oil supply. Illumination of a CHIP DETECT annunciator is not in itself cause for an
engine to be shut down. Engine parameters should be monitored for abnormal indications. If
parameters are abnormal, a precautionary shutdown may be made at the pilot's discretion. After
illumination of a CHIP DETECT annunciator, cause of the malfunction should be determined
and corrected prior to the next flight.
EXPANDED ENGINE SYSTEM PROCEDURES
ENGINE STARTING (EXTERNAL POWER)
Never connect an external power source to the airplane unless the battery is indicating a charge
of at least 20 volts. If the battery voltage is less than 20 volts, the battery must be recharged, or
replaced with a battery indicating at least 20 volts, before connecting external power. Only use
an external power source fitted with an AN-type plug.
NOTE
When an external power source is used, it must be set lo 28.0 to 28.4 volts and be capable of
producing 1000 amperes momentarily and 300 amps continuously. The battery should be ON to
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absorb transient voltage spikes present in some auxiliary power units. An EXT PWR annunciator is provided to alert the crew when an external DC power plug is connected to the airplane.
1) Avionics Master Switch - Confirm OFF
2) Left and Right Generator Switches - CONFIRM OFF
3) Battery - ON
4) External Power Source - TURN OFF, then CONNECT TO AIRPLANE
5) External Power Source - TURN ON
6) Voltmeter - 28.0 TO 28.4 VOLTS
7) Propeller Levers - FEATHER
8) Right ignition and Engine start - on (R FUEL PRESS Annunciator - EXTINGUISHED)
9) Right Condition Lever - LOW IDLE (al12o/o N1 or above)
10) ITT and N1 - MONITOR (1000°C maximum)
If no ITT rise is observed within 10 seconds after moving the condition lever to low idle, move
the condition lever to fuel cut off, allow 60 seconds for fuel to drain and starter to cool, then
follow engine clearing procedures.
11) Right Oil Pressure - CHECK
12) Right ignition and Engine Start - OFF (at 50% N1or above)
13) Left ignition and Engine Start - ON (L FUEL PRESS Annunciator - EXTINGUISHED)
14) Left Condition Lever - LOW IDLE (at 12% N1 or above)
15) ITT and N1 MONITOR (1000'C maximum)
16) Left Oil Pressure - CHECK
17) Left ignition and Engine Start - OFF (at 50% N1 or above)
18) External Power Source - TURN OFF, DISCONNECT, SECURE DOOR
19) Left and Right Generators - RESET, (HOLD for 1 sec) THEN ON
20) Propeller Levers - FULL FORWARD
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No Light Start
1) Condition Lever CUT-OFF
2) Ignition/Start Switch OFF
Allow 60 seconds for fuel to drain and starter cooling; then conduct engine clearing procedures.
ENGINE CLEARING
The following procedure is used to clear an engine at any time it is deemed necessary to remove
internally trapped fuel and vapor, or if there is evidence of a fire within the engine. Air passing
through the engine serves to purge fuel, vapor, or fire from the combustion section, gas generator
turbine, power turbines and exhaust system.
1) Condition Lever - FUEL CUT OFF
2) Ignition and Engine start - STARTER ONLY (for a maximum ol40 seconds)
3) Ignition and Engine Start – OFF
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ENGINE SYSTEM
QUESTIONS
1) What does the term “free-turbine” refer to?
2) N1 refers to RPM of what section of the engine?
3) The PT6A engine power section consists of:
a) One compression stage and four turbine stages.
b) A two-stage reaction turbine.
c) A two-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) 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 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.
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9) The Pratt & Whitney PT6A-41 0r 42 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) True or False: Compressor bleed valves are designed to prevent compressor stalls at reduced power.
12) What is another name for bleed air?
13) What is the approximate power turbine to propeller gear reduction ratio?
14) True or False: The power turbine and N1 shafts turns in opposite direction.
15) At what speed is the compressor turning, at 100% N1?
16) What are the following engine limits for the engine during takeoff?
ITT – 42 ________ -41 ________
TORQUE-42 ________ -41 ________
Np -42 ________ -41________
N1 -42 ________ -41________
17) The Low Idle ITT limit of the engine is -42___, -41___, °C.
18) On a hot day while awaiting take-off clearance, you see the ITT above the Low Idle limit.
What should you do?
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19) True or False: Illumination of a CHIP DETECT annunciator indicates a positive metal
contamination in the engine oil supply.
20) True or False: 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) When is it best to check oil level and service it, if required?
25) What caution is there regarding the addition of oil to your engine?
26) List the starter limitations.
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CHAPTER 6
PROPELLERS
OBJECTIVES
After completing this chapter, you will be able to:
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 emergency procedures.
GENERAL
The King Air 200 utilizes a three or four blade propeller. Serial numbers BB-2 through BB-1438
have a three bladed Hartzell or McCauley prop while later models have a four bladed prop
installed. 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
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engine fails in flight, the propeller will not feather because of the windmilling 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 on the King Air 200 or 8.7 psi on the B200
at power settings of 90 percent N1 or greater. Low pitch propeller position is determined by a
mechanically monitored hydraulic stop.
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 operated in three conditions controlled by a propeller governor. They are:
1) Onspeed
2) Overspeed
3) 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 selected RPM.
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PROPELLER GOVERNOR
The King Air is equipped with three propeller governors. They are the primary governor, the
over-speed governor and the fuel topping governor.
PRIMARY GOVERNOR
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. The primary governor can maintain any selected propeller speed from approximately
1600 RPM to 2000 RPM. Assume an aircraft is in normal cruising flight with the propeller
turning 1700 RPM. If 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 1700 RPM, or "onspeed." Likewise, if the airplane moves 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
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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 1700 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 is 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 actual propeller RPM is below the selected governor 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 1800 RPM but the propeller is turning at less than 1800 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 18º blade angle. But bringing the power lever aft of idle progressively
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repositions the stop to blade angles less than 18°. Keep in mind that just because the low pitch
stop has been moved back to smaller angles than 18°, 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 18º and 5º blade
angle is referred 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). 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.
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 2000 RPM, the overspeed
governor is set at 2080 RPM. As a runaway propeller's speed reaches 2080 RPM, the overspeed
governor will begin increasing blade angle to a higher pitch, to prevent the RPM from continuing
its rise. From a pilot's point of view, a propeller tachometer stabilized at approximately 2080
RPM would indicate failure of the primary governor and proper operation of the overspeed
governor. A test switch can reset this point of the overspeed governor down to approximately
1870 RPM for 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. When the
propeller reaches 2120 RPM, the fuel topping governor limits the fuel flow to the gas generator,
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reducing N1 RPM, which in turn prevents the propeller RPM from exceeding approximately
2200 RPM. The fuel-topping governor vents air pressure from the Fuel Control Unit, which
results in a fuel flow reduction. 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, placarded L AUTOFEATHER and R
AUTOFEATHER.
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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 takeoff 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
full reverse. 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.
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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
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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.
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PROPELLER LIMITATIONS
PROPELLER ROTATIONAL SPEED LIMITS
Transients not exceeding 5 seconds-2200 RPM
Reverse-1900 RPM
All other conditions-2000 RPM
PROPELLER ROTATIONAL OVERSPEED LIMITS
The maximum propeller overspeed limit is 2200 RPM and is time-limited to five seconds.
Sustained propeller overspeeds faster than 2000 RPM indicate failure of the primary governor.
Flight may be continued at propeller overspeeds up to 2080 RPM, provided torque is limited to
1800 foot-pounds. Sustained propeller over-speeds faster than 2080 RPM indicate failure of both
the primary governor and the secondary governor, and such overspeeds are unapproved.
PROPELLER EMERGENCY PROCEDURES NONE
PROPELLER ABNORMAL PROCEDURES NONE
PROPELLER EXPANDED PROCEDURES OVERSPEED GOVERNOR/RUDDER BOOST TEST
1) Rudder Boost Switch ON
2) Propeller Levers FULL FORWARD
3) Propeller Test Switch HOLD TO TEST
4) Left Power Lever 1,800 RPM
5) Left Overspeed Governor/Rudder Boost CHECK (1,870 ± 40)
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6) Left Power Lever IDLE
7) Right Power Lever 1,800 RPM
8) Right Overspeed Governor/Rudder Boost CHECK (1,870 ± 40)
9) 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.
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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) True or False: The prop control levers should be full forward prior to selecting reverse.
4) The overspeed governor is reset to what RPM for testing?
5) True or False: 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) Onspeed
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) When using maximum reverse power at HI IDLE and full increase RPM, you would expect a
maximum propeller RPM of:
a) 2000 RPM
b) 1900 RPM
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c) 2080 RPM
d) 1600 RPM
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CHAPTER 7
PRESSURIZATION AND ENVIRONMENTAL SYSTEMS
OBJECTIVES
After completing this chapter, you will be able to:
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.
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 6.5 psid the cabin can be at sea level with the aircraft at
15,600 feet. With the cabin at 10,000 feet, the aircraft can climb to nearly 35,000 feet before
maximum differential is reached. Although the King Air's pressure vessel is designed to
withstand a normal maximum differential of 6.5 psid, the minimum allowable differential is 0.
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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. The bleed
air control valve is energized by a bleed air switch on the copilot's subpanel.
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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 lines
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 aft of the forward pressure bulkhead. 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. If the air in the heated air duct becomes excessively hot, an overtemperature switch
located in the ducting illuminates the DUCT OVERTEMP caution annunciator.
When the DUCT OVERTEMP
annunciator light comes on, operation of
the temperature and air controls should
lower the temperature. If this fails, the
bleed air bypass valves should be checked
for proper operation. 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. On some airplanes, a solenoid-operated air-balance valve is installed between the
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main heat duct and the two forward floor outlets. The valve is normally closed and limits the
amount of air going to the two forward floor outlets, thereby permitting a balanced flow of air to
the rear of the cabin. When the vent blowers operate, the air-balance valve opens, permitting an
increased flow of air to the two forward floor outlets. When an aft vent blower is installed, an air
check valve between the blower output duct and the heated air duct permits the blower output air
to circulate into the heated air ducting. At cruise power, the heating capacity of the system is
sufficient to maintain cabin temperatures in excess of 65°F at ambient temperatures of -65°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, a duct temperature
sensor, and two heat exchanger bypass valves. The cabin temperature mode switch has four
positions; MANUAL HEAT, MANUAL COOL, OFF and AUTO. The forward 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 deenergize 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
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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.
RADIANT HEAT PANELS
Optional radiant heat panels may be used where additional
heat is required. The radiant heating system on airplanes BB-2
through BB- 449 consists of two heating panels bonded to the
forward and aft headliner. The heating panels are controlled
manually by a single on/off switch on the subpanel. Thermal
switches mounted in the panels provide overheat protection.
The radiant heating system consists of five heating panels
installed above the windows in the service panels. The system
is controlled by an on/off switch on the subpanel. Overheat
protection is provided by a thermostat and a 194° thermal fuse located on the back of each heat
panel. For ease of service, each heating panel is attached to the service panel with six strips of
Velcro tape.
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PILOT TIP
When the airplane is connected to an auxiliary power unit, Radiant Heat can be used to warm the cabin prior to engine start.
ELECTRIC HEAT
An optional electric heat system is used to preheat the interior of the airplane prior to engine
operation and is not designed to supplement engine bleed air heat. The electric heat system
should be powered through a ground power unit, as the ship's battery cannot power the system.
Electric heat is normally operated when cold weather makes it necessary to heat the cabin area
prior to the boarding of passengers. The system is designed so that it only operates when the
airplane is on the ground and the ambient temperature inside is at or below 60° F. Once on, the
thermostatically controlled system will continue to provide heat until a thermostat signals the
electric heat relay that duct temperature has reached approximately 118° F, at which time the
electric heat magnetically held switch releases to turn the electric heat off.
NOTE
Manually holding the electric heat switch in the ON position will not override the electric heat
control relay to operate the electric heat system.
Control of the electric heat system is separate from the automatic and manual temperature
controls for bleed air heat. A control switch, placarded ELEC HEAT on the right inboard
subpanel, energizes the heater power relays for the forward and aft electric heaters. The aft vent
blower switch, placarded AFT BLOWER ON, is located next to the ELEC HEAT switch. The
forward electric heat circuit is enabled when the cabin temperature mode switch is set to the
MAN HEAT position. The aft electric heat circuit is enabled when MAN HEAT is selected and
the AFT BLOWER switch is set to ON. The vent blowers that distribute cool air also distribute
the heat produced by the electric heaters. Overheat sensors cutoff power to the electric heaters if
duct temperature reaches approximately 118°F or above.
FRESH AIR VENTILATION
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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 ambient air obtained (through a check valve) from the condenser section in the nose of the
airplane. During pressurized operation, cabin pressure forces the check valve closed. During the
unpressurized mode, a spring holds the check valve open, so that the forward blower can draw
this air into the cabin. The ambient air then mixes with recirculated cabin air, goes through the
forward blower, through the forward evaporator, (if it is operating, the air will be cooled), into
the mixing plenum, into both the ceiling-outlet and the floor-outlet duct, and into the cabin
through all the ceiling and floor outlets. 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 open or closed.
COOLING - DESCRIPTION AND OPERATION
The King Air 200 air-conditioning system is similar to a home or automotive system. The air-
conditioner system consists of five major components. They are the evaporator(s), condenser,
expansion valves, compressor and receiver/dryer. During operation, the belt-driven compressor,
located on the right engine, compresses the refrigerant gas to a high pressure, high temperature
vaporized gas. The gas is routed through a condenser coil, located in the nose of the fuselage,
where cooling air drawn through the condenser by a blower removes heat from the gas, thereby
condensing it to a liquid. The liquid then passes through the receiver/ dryer, located to the left of
the condenser, where any moisture or foreign material is removed from the Freon. From here the
liquid refrigerant flows to the expansion valve where it is metered into 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 which is drawn over the
evaporator cooling fins by the ventilation blower which also distributes heated or cooled air to
the cabin. The forward evaporator and forward vent blower are located in the right nose keel
section. An optional aft evaporator and aft vent blower, for additional cooling capacity, are
located under the center aisle floorboard aft of the wing main spar. If the optional evaporator and
vent blower are installed, the forward vent blower distributes air to the forward overhead outlets,
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the crew compartment outlets and the
forward floor outlets. The aft
evaporator and vent blower will
supply air to the aft overhead outlets,
the rear floor outlets and the toilet
compartment (if installed). If only the
forward evaporator and vent blower
are installed, air will be supplied to
all outlets. The air conditioning system with only the forward evaporator is rated at 18,000 BTU.
The combined rated output of both forward and aft evaporators is 32,000 BTU at 70% N1
turbine speed.
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.
AUTOMATIC OPERATION
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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
refrigerant compressor clutch and condenser blower. The clutch 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 refrigerant
compressor clutch 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. These switches are attached to the refrigerant lines
under the right leading edge of the wing center section. When the switches are actuated on early
model 200's, a fuse located in the right side of the wing center section will blow; on later model
200's, a reset switch located in the nose wheel well will de-energize the system. When the fuse is
blown or the reset switch opened, both the condenser blower and the compressor are shut down.
The vent blower will remain in operation to provide cabin air circulation. When a pressure
switch is actuated, the system should be thoroughly checked before being returned to service;
however, when a service facility is not readily available and air conditioning is required, the reset
switch on the late model 200's may be depressed to actuate the system. It may be assumed that
the circuit at the switch is closed when the light on the reset switch button is extinguished.
MANUAL COOL OPERATION
With the cabin temperature mode switch in the MANUAL COOL position, the compressor
clutch 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
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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 unless the manual temperature switch is held in the decrease position for 1 minute.
FORWARD EVAPORATOR FREEZE PROTECTION
An automatic hot gas bypass valve, located in the refrigerant plumbing in the front evaporator
section, operates to prevent freeze-up of the evaporator by routing the hot refrigerant gas around
the expansion valve. This maintains a constant evaporator temperature just above freezing. A
33°F thermal switch is installed in the forward evaporator section to operate the bypass valve.
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.
Bleed air comprises as much as 10 pounds of air flow on cold days and as little as 6 pounds 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
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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
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connected to the modulating valve. This pneumostat is located on the lower aft side of the
fireseal forward of the firewall. The bimetallic sensing discs of the thermostat are inserted into
the cowling intake. These discs sense ambient temperature and regulate the size of the thermo-
stat orifices. Warm air will open the orifice and cold will restrict it until, at 30ºF, the orifice will
be completely closed. 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, a cabin pressure controller, safety and preset solenoids.
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 6.1 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 6.1 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 by 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
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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 12,500 feet, the cabin altitude warning pressure switch closes and the
warning annunciator light labeled ALT WARN will illuminate.
OXYGEN SYSTEM
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The system consists of an oxygen bottle mounted in the aircraft tail section, oxygen mask
compartments in the cabin ceiling, a mask compartment in front of the toilet and a first aid mask
over the cabin door. The pilot has two controls in the cockpit overhead panel; one to arm the
system, labeled "PULL ON - SYSTEM READY," and a manual override control knob as a
backup. In addition there are crew mask outlets in the cockpit.
When the system is "armed," oxygen pressure regulated down to 70 psi is sent to a solenoid in
the forward cabin ceiling. Next to the solenoid is a cabin pressure sensing switch which upon
sensing a cabin above 12,500 feet will open the oxygen solenoid. The 70 psi pressure is then sent
to pressure activated plungers in each mask compartment to drop the doors. When the masks fall
out, they must be pulled to remove the pin from the oxygen flow valves in the mask
compartment. On aircraft before BB-450, the cabin barometric pressure switch will turn on the
cabin fluorescent lights and cabin signs, and a pressure switch on the single mask in front of the
toilet will turn on the "PASS OX ON" annunciator light. On aircraft after BB-450, the pressure
switch on the single oxygen mask illuminates the cabin signs, fluorescent lights, and the "PASS
OX ON" annunciator light.
The manual override system mechanically opens the oxygen solenoid to insure mask deployment
should the automatic mode malfunction.
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PILOT TIP
The oxygen bottle is fully charged when it reads 1850 psi at 15º C.
PRESSURIZATION AND ENVIRONMENTAL SYSTEMS LIMITATIONS
CABIN DIFFERENTIAL PRESSURE GAGE
Green Arc (Approved Operating Range) 0 to 6.6 psi Red Arc (Unapproved Operating Range) 6.6 psi to end of scale
EMERGENCY PRESSURIZATION AND ENVIRONMENTAL SYSTEMS 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.
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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 System Ready - PULL ON (verify)
2) Crew (Diluter Demand Masks) - DON MASKS
3) Mic Selector - OXYGEN MASK
4) Audio Speaker - ON
5) Passenger Manual Drop Out - PULL ON
6) Passengers - PULL LANYARD PIN, DON MASK
7) Oxygen Duration - CONFIRM
8) First Aid Oxygen - AS REQUIRED
a. Oxygen Compartment - PULL OPEN
b. ON/OFF Valve - ON
c. Mask - DON
PRESSURIZATION LOSS (ALT WARN Annunciator)
1) Oxygen
a) Oxygen System Ready - PULL ON (verify)
b) Crew - DON MASK
c) Mic Selector - OXYGEN MASK
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d) Audio Speaker – ON
e) Passenger Manual Drop Out - PULL ON
f) Passengers - PULL LANYARD PIN, DON MASK
2) Descend as required.
3) Range - DETERMINE FOR FINAL CRUISE ALTITUDE
4) Oxygen Duration - CONFIRM
AUTO-DEPLOYMENT OXYGEN SYSTEM FAILURE
(ALT WARN Annunciator Illuminated, PASS OXY ON Annunciator Not Illuminated)
1) Passenger Manual Drop Out - PULL ON
2) First Aid Mask (if required) - DEPLOY MANUALLY 3) Oxygen Control Circuit Breaker - PULL
4) Passenger Manual Drop Out - PUSH OFF
HIGH DIFFERENTIAL PRESSURE (Cabin Differential Pressure Exceeds 6.6 PSI)
1) Bleed Air Valves - ENVIR OFF
2) Oxygen (Crew and Passengers) - AS REQUIRED
3) Descend - AS REQUIRED
WARNING
ADEQUATE OXYGEN PRESSURE IS NOT PROVIDED TO THE PASSENGERS FOR SUSTAINED FLIGHT AT CABIN ALTITUDES ABOVE 34,000 FEET. THE HIGHEST
RECOMMENDED CABIN ALTITUDE FOR SUSTAINED FLIGHT IS 25,000 FEET.
SMOKE AND FUME ELIMINATION
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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 Ready - PULL ON (Verify)
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) 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:
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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 System Ready - PULL ON (Verify)
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 - HI
4) Left Bleed Air Valve - ENVIR OFF
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 - ENVIR OFF
8) If smoke decreases, continue operation with right bleed air off.
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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) Oxygen - CREW REQUIRED (passengers as required)
a) Oxygen System Ready - PULL ON (verify)
b) Crew (Diluter Demand Masks) - DON MASKS
c) Mic Selector - OXYGEN MASK
d) Audio Speaker - ON
e) Passenger Manual Drop Out - PULL ON
f) Passengers - PULL LANYARD PIN, DON MASK
2) Power Levers - IDLE
3) Propeller Levers - FULL FORWARD
4) Flaps - APPROACH
5) Landing Gear - DN
6) Airspeed - 181 KNOTS MAXIMUM
ABNORMAL PRESSURIZATION AND ENVIRONMENTAL SYSTEMS PROCEDURES
DUCT OVERTEMPERATURE
1) Vent Blower - HIGH
2) Cabin and Cockpit Air - PUSH IN (to increase airflow to cabin) If Condition Persists:
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3) Cabin Temp Mode - MAN HEAT
4) Manual Temp - DECREASE (for 60 seconds)
If Condition persists, the Right Bypass Valve May Be Inoperative, Preventing Both Valves from moving to the Colder Position.
5) Left Bleed Air Valve - ENVIR OFF
If the DUCT OVERTEMP Annunciator Does Not Extinguish after 2 Minutes:
6) Oxygen - AS REQUIRED
7) Right Bleed Air Valve - ENVIR OFF Descend as required.
PRESSURIZATION AND ENVIRONMENTAL SYSTEMS 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 SYSTEM PREFLIGHT INSPECTION
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1) Passenger Manual Drop Out - PUSH OFF
2) Oxygen System Ready - PULL ON
3) Crew Diluter Demand Masks - DON MASK, CHECK FIT AND OPERATION, AND STOW
4) Oxygen Duration - DETERMINE
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PRESSURIZATION AND 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.
5) What is the function of the by-pass valves?
6) What controls radiant heat?
7) What is the normal allowable max differential pressure for the Model 200?
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) List the memory items on the Loss of Pressurization Checklist.
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 provides overheat protection for the radiant heat panels?
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15) True or False: With the cabin at 10,000 feet, the aircraft can climb to nearly 35,000 feet
before maximum differential is reached.
16) In what position should the condition levers be for a pressurization test?
a) High
b) Low
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CHAPTER 8
LANDING GEAR, TIRES AND BRAKE SYSTEM
OBJECTIVES
After completing this chapter, you 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 200 utilizes two types of landing gear systems depending on serial number of the
aircraft. BB-2 through BB-1192 use the
mechanical landing gear system. Aircraft BB-
1193 and after, utilize a hydraulic system.
Both systems are controlled by a handle
placarded LDG GEAR CONTROL - UP - DN
on the pilot's 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.
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Visual indication of landing gear position is provided by individual green GEAR DOWN
annunciators placarded NOSE -L –R on the pilot's right subpanel. The annunciators may be
checked in flight by pressing the annunciator. 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. 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
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gear steering stop block is installed to warn the pilot if tow limits have been exceeded. The
maximum nose wheel turn angle is 48° left and right. When ground handling the airplane, do not
use the propellers or control surfaces as hand holds to push or move the airplane.
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
50 degrees to the right or left of center,
allowing the airplane to be turned within a
39'10" 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
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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 takeoff 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 - MECHANICAL LANDING GEAR
The landing gear is operated by a split-field series wound motor, mounted on the forward side of
the center section main spar. One field is used to drive the motor in each direction. To prevent
over-travel of the gear, a dynamic brake relay simultaneously breaks the power circuit to the
motor and makes a complete circuit through the armature and the unused field winding. The
motor then acts as a generator and the resultant electrical load on the armature stops the gear.
The main gear actuators are driven by torque shafts that carry torque from the gear box. The nose
gear actuator is driven by Duplex chain that attaches to a sprocket on the gearbox torque shaft. A
spring loaded friction clutch between the gear box and the torque shaft protects the motor in the
event of mechanical malfunction. The main gears are held in the down-lock position by a hook
and lock plate arrangement on each main gear drag brace. The nose gear is held in the down-lock
position by the slight over center positioning of the nose gear drag brace. The drag brace is
locked in position by the actuator. The jackscrew in each actuator holds the main and nose gears
in the retracted position. An alternate extension jack mounted between the pilot and copilot seats
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provides a means of landing gear extension in the event of a landing gear motor or electrical
system malfunction.
Manual landing gear extension is provided through a separate, chain-drive system. To engage the
system, pull the LDG GEAR RELAY circuit breaker, located to the left of the landing gear
control handle on the pilot's right subpanel, and ensure that the landing gear control handle is in
the DN position. Pull up on the alternate engage handle (located on the floor) and turn it
clockwise until it stops. This will electrically disconnect the motor from the system and lock the
alternate drive system to the gear box.
With the alternate drive locked in, the chain is driven by a continuous-action ratchet, which is
activated by pumping the alternate extension handle located adjacent to the alternate engage
handle. As many as 50 full strokes may be required to fully extend the landing gear. Stop
pumping when all three green gear-down annunciators are illuminated. Further movement of the
handle could damage the drive mechanism and prevent subsequent electrical gear retraction.
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.
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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.
After a practice manual extension of the landing gear, the gear may be retracted electrically.
The landing gear control lever on the pilot's inboard subpanel controls the landing gear. A safety
switch on the right main gear torque knee opens the control circuit when the strut is compressed.
The safety switch also activates a solenoid-operated down-lock hook on the landing gear control
handle located on the pilot's right subpanel. This mechanism prevents the landing gear control
handle from being raised when the airplane is on the ground. The hook automatically unlocks
when the airplane leaves the ground. In the event of a malfunction of the down-lock solenoid, the
down lock can be released by pressing downward on the red down-lock release button. The
release button is located just left of the landing gear handle. The landing gear control handle
should never be moved out of the DN detent while the airplane is on the ground. Moving the
gear handle out of the DN position while the aircraft is on the ground will cause the landing gear
warning horn to sound intermittently and the red gear-in-transit lights in the landing gear control
handle to illuminate (provided the MASTER SWITCH is ON). To prevent accidental retraction
of the landing gear while the airplane is on the ground, a safety switch mounted on each of the
main gears cuts power to the control circuit when the shocks are compressed.
CAUTION
NEVER RELY ON THE SAFETY SWITCH TO KEEP THE GEAR DOWN.
THE LANDING GEAR CONTROL SWITCH MUST BE IN THE DOWN POSITION.
WARNING SYSTEM MECHANICAL LANDING GEAR SYSTEM
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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 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. The lights in the
landing gear control handle cannot be canceled. 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 canceled.
DESCRIPTION AND OPERATION- HYDRAULIC LANDING GEAR
The nose and main landing gear assemblies are operated by a hydraulic power pack in the left
wing center section forward of the main spar. The two main components of the power pack are
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the motor and the hydraulic pump. Installed on the hydraulic pump housing are a pressure switch
and a low fluid filter. To prevent pump cavitation, an engine bleed air pressure of 18 to 20 psi is
plumbed to the power pack and hydraulic fill reservoirs. Three separate hydraulic lines are
routed from the power pack to each of the actuators and supply hydraulic pressure for each of the
landing gear modes which include retraction, extension, and emergency extension. A landing
gear control switch on the pilot's inboard subpanel controls the landing gear. A solenoid-operated
down lock latch prevents the switch from being actuated while the airplane is on the ground.
This latch can be overridden by depressing the red down lock-release switch. To prevent
accidental retraction of the landing gear, a safety switch mounted on each main gear cuts power
to the control circuit whenever the shock struts are compressed.
CAUTION
NEVER RELY ON THE SAFETY SWITCH TO KEEP THE GEAR DOWN WHILE
TAXIING. THE LANDING GEAR CONTROL SWITCH MUST BE IN THE DOWN
POSITION DURING ALL GROUND OPERATIONS.
When the landing gear handle is moved to the down position, the power pack down solenoid
routes hydraulic fluid to the extend portion of the system. As the actuator piston moves 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
actuator pistons are positioned to
fully extend the landing gear, an
internal mechanical lock in the
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nose gear actuator will lock the actuator piston to hold the nose gear in the down position. The
main gears are held by a mechanical locking system. In this position, the internal locking
mechanism in the nose gear actuator will actuate the actuator down lock switch to interrupt
current to the pump motor. The motor will continue to run until all three landing gears are down
and locked. 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 using the
hydraulic power pack; however, the landing gear can be extended using the emergency extension
hand pump. A sensing unit mounted on the motor end of the power pack provides the circuitry to
illuminate the low-fluid light. The optically operated sensing unit has a self-test circuit. The
integral self-test circuit is energized by a switch on the instrument panel and tests the sensing
unit's internal circuitry. Manual landing gear extension is provided through a manually powered
hydraulic system. 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.
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A 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 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 and pumped through the power pack hand pump pressure port to the actuators. The
pressure exerted on the secondary extend port of the actuators shifts the shuttle valves, allowing
the fluid to enter the extend side of 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 returned 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 dump valve of the
pump to allow the return fluid to flow into the primary reservoir. As many as 80 full strokes may
be required to fully extend the landing gear. 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 fully down position prior to placing the pump handle in the securing clip.
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When the pump handle is stowed, an internal relief valve is actuated to relieve the hydraulic
pressure in the pump. After a practice manual extension of the landing gear, the gear may be
retracted hydraulically.
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.
WARNING SYSTEM HYDRAULIC LANDING GEAR SYSTEM
The landing gear warning system is provided to warn the pilot that the landing gear is not down
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 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. The lights in the landing gear control
handle cannot be canceled. The landing gear warning system will be rearmed if the power levers
are advanced sufficiently. With the flaps beyond APPROACH position, the warning horn and
landing gear switch handle lights will be activated regardless of the power settings, and neither
can be canceled.
A fill reservoir is located just inboard of the LH nacelle and forward of the front spar. It contains
a cap and dipstick assembly which is marked HOT/FILL, COLD/FILL, to check system fluid
level.
TIRES
The airplane utilizes a pair of 18x5.5 8 ply tires on each main gear assembly. However, an
optional 10-ply-rated tire can be used. If one main tire becomes deflated, it should be possible to
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conclude operation in a safe and normal manner on the other tire. A 22x6.75-10, 8-plyrated tire
is installed on the nose gear. As an option, the standard main gear can be replaced with a high
flotation gear. The main difference in this gear is that larger, low pressure 22x6.75-10 8 ply tires
are utilized. The larger footprint (per gear average of 40.5 sq. in. on the high float versus 24.5 sq.
in. on the standard gear) and lower ground contact pressure (per gear average of 72 P.S.I. on the
high float gear versus 119 P.S.I. on the standard gear) of the high flotation landing gear make it
more desirable for rough/soft field operations.
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 standard main wheel
tires (18x5.5) to 96 psi. Inflate the optional high flotation main wheel tires (22x6.7510) to 62 psi.
Both the standard and high flotation configuration nose wheel tires should be inflated to between
55 and 60 psi.
HYDRAULIC BRAKE SYSTEM
The dual hydraulic brakes are operated by depressing the pilot's or copilot's rudder pedals.
Airplanes prior to BB-666 are equipped with a shuttle valve adjacent to each set of pedals. The
shuttle valve permits the changing of braking action from one set of pedals to the other so
whoever brakes first has control. The dual brakes on airplanes BB-666 and after are plumbed in
series so that if both crew members apply pedal force, the resulting total force is applied to the
brakes. The pilot's master cylinders are plumbed through the copilot's master cylinders, thus
allowing either set of pedals to perform the braking action and eliminating the need for shuttle
valves. The depression of either set of pedals compresses the piston rod in the master cylinder
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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 disc
brake assemblies on the main landing gear. This pressure forces the brake pistons to press against
the 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 pedals 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 handle 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 brake pedals are depressed and the parking brake control is
pushed in. Most aircraft are equipped with automatic brake adjusters. The automatic brake
adjusters reduce brake drag, thereby allowing unhampered roll. Airplanes with the high flotation
landing gear and brakes are not equipped with the automatic brake adjusters and cannot be
reworked to accept them.
Brake system servicing is limited
primarily to maintain the hydraulic fluid
level in the reservoir mounted in the
upper LH corner of the aft bulkhead of
the nose baggage compartment. A dip
stick is provided for measuring the fluid
level. When the reservoir is low on fluid,
add a sufficient quantity of MIL-H-5606
hydraulic fluid to fill the reservoir to the
full mark on the dipstick.
Each wheel cylinder (except those
airplanes equipped with optional brake deice) is provided with a means of conveniently checking
brake wear. The distance between the piston housing and the lining carrier will increase with
lining wear. When the distance exceeds 0.250 inch (as indicated by the accompanying
illustration) the brakes should be replaced. This check should be accomplished with brake
pressure applied.
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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.
LANDING GEAR, TIRES AND BRAKE SYSTEM LIMITATIONS
LANDING GEAR CYCLE LIMITS
Landing gear cycles (1 up - 1 down) are limited to one every 5 minutes for total of 6 cycles
followed by a 15 minute cool-down period.
Maximum Landing Gear Operating Speed
182 164
181 163
Do not extend or retract landing gear above the speeds given.
Maximum Landing Gear Extended Speed WE
182 181 Do not exceed this speed with landing gear extended.
LANDING GEAR, TIRES AND BRAKE SYSTEM ABNORMAL PROCEDURES
BOLD TYPE INDICATES MEMORY ITEMS!
NONE
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LANDING GEAR, TIRES AND BRAKE SYSTEM EMERGENCY PROCEDURES
HYDRAULIC FLUID LOW (HYD FLUID LOW Annunciator)
If HYD FLUID LOW annunciator illuminates during flight, attempt to extend the landing gear
normally upon reaching destination. If the landing gear fails to extend, follow LANDING GEAR
MANUAL EXTENSION procedures.
LANDING GEAR MANUAL EXTENSION (HYDRAULIC SYSTEM)
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 subpanel) – PULL
2) Landing Gear Control – DN
3) Alternate Extension Handle - PUMP UP AND DOWN UNTIL THE THREE GREEN
GEAR-DOWN ANNUNCIATORS ARE ILLUMINATED. WHILE PUMPING, DO
NOT LOWER HANDLE TO THE LEVEL OF THE SECURING CUP DURING THE
DOWN STROKE AS THIS WILL RESULT IN THE LOSS OF PRESSURE. If all three
green gear-down annunciators are illuminated:
4) Alternate Extension Handle - STOW
5) Landing Gear Controls - DO NOT ACTIVATE (The Landing Gear Control and the
Landing Gear Relay Circuit Breaker must not be activated. The landing gear should be
considered UNSAFE until the system is cycled and checked with the airplane on jacks.)
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 UNTIL MAXIMUM
RESISTANCE IS FELT.
7) Alternate Extension Handle - DO NOT LOWER. LEAVE AT THE TOP OF THE UP
STROKE.
Prior to Landing:
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8) Alternate Extension Handle - PUMP UNTIL MAXIMUM RESISTANCE IS FELT. DO
NOT STOW.
After Landing:
9) Alternate Extension Handle - CONTINUE PUMPING, WHEN CONDITIONS PERMIT,
TO MAINTAIN HYDRAULIC PRESSURE UNTIL THE GEAR CAN BE
MECHANICALLY SECURED. DO NOT STOW HANDLE. DO NOT ACTIVATE
THE LANDING GEAR CONTROL OR THE LANDING GEAR RELAY CIRCUIT
BREAKER. THE LANDING GEAR SHOULD BE CONSIDERED UNLOCKED
UNTIL THE SYSTEM IS CYCLED AND CHECKED WITH THE AIRPLANE ON
JACKS.
LANDING GEAR MANUAL EXTENSION (MECHANICAL SYSTEM)
If the landing gear fails to extend after placing the Landing Gear Control down, perform
the following:
1) Airspeed - ESTABLISH 130 KNOTS
2) Landing Gear Relay Circuit Breaker (pilot's subpanel) – PULL
3) Landing Gear Control – DN
4) Alternate Engage Handle - LIFT AND TURN CLOCKWISE TO THE STOP TO
ENGAGE.
5) Alternate Extension Handle - PUMP UP AND DOWN UNTIL THE THREE GREEN
GEAR-DOWN ANNUNCIATORS ARE ILLUMINATED. ADDITIONAL PUMPING
WHEN ALL THREE ANNUNCIATORS ARE ILLUMINATED COULD DAMAGE
THE DRIVE MECHANISM AND PREVENT SUBSEQUENT ELECTRICAL GEAR
RETRACTION.
If all three green gear-down annunciators are illuminated:
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6) Alternate Extension Handle - DO NOT STOW (Proceed to step 8.) If one or more
green gear-down annunciators do not illuminate for any reason and a decision is made to
land in this condition:
7) Alternate Extension Handle - CONTINUE PUMPING UNTIL MAXIMUM
RESISTANCE IS FELT, EVEN THOUGH THIS MAY DAMAGE THE DRIVE
MECHANISM.
8) Landing Gear Controls - DO NOT ACTIVATE (The Landing Gear Control and the
Landing Gear Relay Circuit Breaker must not be activated. The landing gear should be
considered UNSAFE until the system is cycled and checked with the airplane on jacks.)
LANDING GEAR, TIRES AND BRAKE SYSTEM
EXPANDED PROCEDURES
BRAKE DEICE CHECK
1) Power Levers __________________________________1,800 RPM (NOTE ITT) 2) Brake Deice Switch ___________________________ ON (DEICE LIGHT ON) 3) Left and Right ITT______________________________ SLIGHT INCREASE 4) Brake Deice Switch _____________ OFF (ITT RETURN TO VALUE IN STEP 1)
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LANDING GEAR, TIRES AND BRAKE 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) 115 K
2) What is the tire pressure for the mains? For the nose gear tire?
3) Prior to serial number B666, who controls how much brake force is applied?
a) The pilot.
b) The co-pilot.
c) The pilot who applied brakes first.
d) The pilot who applies the most force to the brake pedals.
4) True or False: Brake wear can be checked during preflight.
5) Where is the brake fluid reservoir located?
6) When could you not silence the landing gear warning horn with the horn silence button?
7) If manually extending the landing gear, when would you stop pumping? Why?
8) Where is the landing gear relay control circuit breaker located?
9) 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.
10) The gear warning horn will sound when the gear is not down and:
a) Either power lever is reduced to a certain setting.
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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.
11) The emergency landing gear extension system utilizes:
a) A hand crank located behind the pilot's seat.
b) A hand pump and release mechanism located in the cockpit.
c) A nitrogen blow-down bottle.
d) A mechanical drop-down release.
12) True or False: Once the gear has extended manually, it can be retracted normally.
13) Airspeeds for the landing gear:
a) Maximum gear extended speed __ KCAS
b) Maximum gear extension speed __ KCAS
c) Maximum gear retraction speed __ KCAS
14) Is the parking brake hydraulic or mechanical?
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CHAPTER 9
PNEUMATIC AND VACUUM SYSTEM
OBJECTIVES
After completing this chapter, you will 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
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 bleed air warning system, the autopilot and the surface deicer
system. 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
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incorporates a safety valve that will limit pressure to 3 psi higher than that setting as a safety
feature in the event of regulator failure. From the pressure regulator valve, lines are routed to the
various aircraft systems that utilize pneumatic pressure.
VACUUM SYSTEM - DESCRIPTION AND OPERATION
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
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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 airplane 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
plumbed 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-WARNING SYSTEM - DESCRIPTION AND OPERATION
This system provides a visual warning of a rupture in a bleed-air or pneumatic line. The warning
provides sufficient time to shut down the bleed-air firewall-shutoff valve on the affected side
before the heat from the rupture has time to damage the structure, skin or adjacent components.
The bleed- air lines from the engine to the cabin are shielded with oven insulation and foil tape to
retain the bleed-air heat in the system and to protect nearby components. The bleed-air and
pneumatic lines that run through the nacelles, center section, and fuselage are accompanied in
close proximity by the bleed-air warning tubes. When the heat from a ruptured bleed-air or
pneumatic line comes into contact with the plastic warning line, the warning line will melt and
burst (at approximately 204° F), releasing 17 to 22 psi of internal pressure and triggering the
applicable pressure switch. When the pressure at the switch drops to 1 to 2 psi, the switch closes
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and illuminates the appropriate red BL AIR FAIL warning annunciator in the warning
annunciator panel. The two pressure switches are mounted beside the pedestal under the copilot's
floorboard. One switch monitors the warning system for the LH side of the airplane and the other
switch monitors the system for the RH side of the airplane. The two switches and associated
tubing are pressurized by air tapped off the deice manifold. The bleed-air warning lines have a
clearance of one to four inches between the warning tubes and pneumatic lines.
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PNEUMATIC AND VACUUM SYSTEM 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 AND VACUUM SYSTEM EMERGENCY PROCEDURES
BOLD TYPE INDICATES MEMORY ITEMS!
BLEED AIR LINE FAILURE (L or R BL AIR FAIL Annunciator)
Warning annunciators should be monitored during engine start procedure. Either engine will
extinguish both annunciators upon starting.
Illumination of a warning annunciator in flight indicates a possible rupture of a bleed air line aft
of the engine firewall.
1) Bleed Air Valve (affected engine) - INSTR & ENVIR OFF position
2) Engine Instruments – MONITOR
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NOTE
The bleed air warning annunciator will not extinguish after closing the Bleed Air Valve.
PNEUMATIC AND VACUUM SYSTEM ABNORMAL PROCEDURES
NONE
PNEUMATIC AND VACUUM SYSTEM EXPANDED PROCEDURES
VACUUM/PNEUMATIC PRESSURE CHECK (1,800 RPM)
1) Left Bleed-Air Switch INST/ENVIRO OFF
2) Pneumatic/Vacuum Gage PNEU 12-20/VAC 4.3-5.9 psi
3) Right Bleed-Air Switch INST/ENVIRO 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
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PNEUMATIC AND VACUUM SYSTEM
QUESTIONS
1) What is the purpose of the Bleed Air Failure warning lights?
2) What is the procedure if a Bleed Air Failure light illuminates in flight?
3) True or False: The Bleed Air Failure light will remain illuminated after closing the bleed air
switch.
4) How is the vacuum source created?
5) True or False: The cabin pressurization control system uses valves to operate the controller
and outflow.
6) The Bleed air warning line will melt and burst at approximately:
a) 204ºC
b) 204ºF
c) 300ºF
d) 250ºC
7) Normal gyro suction is ____ psi.
8) Normal pneumatic pressure is ____ psi.
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CHAPTER 10
ANTI-ICE SYSTEM
OBJECTIVES
After completing this chapter, you will 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 the 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
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The pneumatic deice boots on the wings and on the horizontal stabilizer 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 stabilizers. 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 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 6 seconds for the wings and 4 seconds for the tail, 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 in the inflated position 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.
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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. 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
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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.
AIR INTAKES
INERTIAL ICE SEPARATION SYSTEM (BB-1443 AND PRIOR)
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 a bypass door
in the lower cowling at the aft
end of the air duct. The inertial
ice vane and bypass door are
extended and retracted
simultaneously through a
linkage system connected to an
electric actuator. The actuator is
energized through a 3-position
switch placarded ICE
RETRACT, VANE -EXTEND -
in the pilot's outboard subpanel.
A mechanical backup system is
provided which may be actuated
by pulling the T-handles
(placarded ICE VANE EMERGENCY MANUAL - PULL - LEFT ENG - RIGHT ENG) just
below the left subpanel.
When the ice vane switch is placed in the RETRACT position, the inertial ice vane and bypass
door retract out of the airstream. When the vane is fully extended, micro switches on the vane
linkage will illuminate the green L ICE VANE EXT and R ICE VANE EXT annunciators in the
caution/advisory annunciator panel. When the ice vane switches on the subpanel are actuated,
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they will energize a 15-second time-delay circuit. If full extension of the ice vanes is not attained
in the 15 seconds, the amber L ICE VANE or R ICE VANE annunciators in the caution/advisory
annunciator panel will illuminate, signaling a malfunction of the power actuator system. Full
extension must then be accomplished with the manual override control. Once the manual
override system has been operated, the electrical actuator will not actuate the linkage to the ice
vane until the mechanical override has been manually disengaged.
CAUTION
TO AVOID DAMAGE TO THE LINKAGE, THE OVERRIDE ASSEMBLY MUST BE RESET BEFORE THE SYSTEM IS OPERATED ELECTRICALLY.
The ice vane and bypass door should be either fully extended or fully retracted. There are no
intermediate positions. In the retracted (non-icing) position, the annunciator lights will be off.
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.
DUAL-MOTOR INERTIAL ICE SEPARATION SYSTEM (BB-1444 AND
AFTER)
An inertial ice separation system is installed in each engine air inlet to prevent moisture particles
from entering the engine inlet plenum 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 plenum. The heavy ice-laden air is then discharged
overboard through a bypass door in the lower cowling at the aft end of the air duct. The inertial
ice vane and bypass door are extended or retracted simultaneously through a linkage system
connected to an electric dual-motor actuator. The dual-motor actuator is controlled with two
switches for each of the left and right engine systems. The ACTUATOR switch is in the MAIN
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position except when the ACTUATOR STANDBY position is used to actuate the backup motor
because the main motor is inoperable. Power is applied to the motor by placing the ENGINE
ANTI-ICE LEFT and RIGHT switches in the ON position to extend or OFF position to retract.
During non-icing conditions, the inertial ice vane and bypass door are in the retracted positions.
In icing conditions, the inertial ice vane and the bypass door are fully extended by the main
actuator motor. When the doors are fully extended, micro switches on the inertial ice vane
linkage will illuminate the green L ENG ANTI-ICE or R ENG ANTI-ICE annunciators in the
caution/advisory annunciator panel. When the control switches on the subpanel are actuated, a 33
second time-delay circuit is energized. If full extension of the ice vanes is not attained in the 33
seconds, the yellow L ENG ICE FAIL or R ENG ICE FAIL annunciator in the caution/advisory
annunciator panel will illuminate to signal a malfunction of the main actuator motor. Full
extension must then be accomplished with the standby actuator motor. The inertial ice vanes and
bypass doors should be fully extended or fully retracted. There are no intermediate positions. In
the non-icing position, the annunciator lights will be off.
PILOT TIP
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 hot exhaust gases
to prevent the formation of ice during inclement weather. This system is in
operation any time the engines are running. The anti-ice lip is riveted to
the lower forward cowl assembly. On airplanes BB-1265 and prior, a
scoop in each of the engine exhaust stacks deflects some of the hot
exhaust gases downward into the hollow lip tube that encircles the engine
air intake. The gases are exhausted through an opening at the bottom of
the cowling immediately aft of the air intake. On airplanes BB-1266 and
after, a scoop in the left exhaust stack on each engine diverts some of the hot exhaust gases
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downward through a duct into the hollow lip tube that encircles the engine air intake. The
exhaust is ducted into the right exhaust stack where it is expelled into the atmosphere.
BRAKE DEICE SYSTEM
Engine bleed air is routed by line and hose through a
solenoid-operated shutoff valve to a distributor manifold
that directs hot air to the brakes for deicing during
inclement weather and conditions. The heated air for
brake deicing is supplied by bleed air from the
compressor of each engine. The brake deice system is
plumbed into the bleed air system that provides air for
surface deice and instrument vacuum operation. The engine bleed air is routed to each main gear
wheel well. From there bleed air is routed through a distributor manifold and directed to the
brake for each wheel.
The brake deice system is controlled by an ON-OFF toggle switch mounted on the pedestal
immediately aft of the pressurization controller. When this switch is in the ON position, power
from the airplane electrical system is supplied to open the solenoid shutoff valves in each wheel
well, allowing the hot bleed air to enter the distributor manifold for diffusion through the orifices
to deice the brakes. This action also provides a signal to illuminate the BRAKE DEICE ON
(green) light in the annunciator panel on the pedestal. If the pilot fails to turn the system off after
takeoff, a timing circuit will cycle the deice system off after 10 minutes to shut off the flow of
bleed air to the brakes to prevent damage through overheating. The system cannot be activated
again until the landing gear has been cycled. The brake deice system is the single largest user of
engine bleed air. If an engine failure occurs while brake deice is on, rudder boost may not be
available because of insufficient differential pressure to activate the system.
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PILOT TIP
The brake deice valves may become inoperative if the valves are not cycled at least once a day regardless of weather conditions. Do not leave the system on longer than required to do a
function test if the OAT is above 15ºC.
WINDOWS AND WINDSHIELDS
Electrical heating elements embedded in the
windshield 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 constant
temperature of 95ºF to 105ºF is maintained.
Windshield heat switches are located on the pilot's
subpanel (in-board) and are placarded ICE - WSHLD
ANTI-ICE - NORMAL - OFF - HI - PILOT - COPILOT. Two levels of heat are provided. When
the switches are in the NORMAL (up) position, heat is supplied to the major portion of the
windshields. When they are in the HI (down) position, a higher level of heat is supplied to a
smaller area of the windshields. Each switch must be lifted over a detent before it can be moved
into the HI position. This lever-lock feature prevents inadvertent selection of the HI position
when moving the switches from NORMAL to the OFF (center) position. Controllers with
temperature-sensing units provide for proper heat at the windshield surfaces conditions. Either or
both windshields can be heated at any time since overheating is prevented by thermal sensors.
The heating elements are connected at terminal blocks in the corner of the glass to the wiring
leading to the control switches mounted in the left sub-panel. Five-ampere circuit breakers,
located on a panel on the forward pressure bulkhead, protect the control circuits. The power
circuit of each system is protected by a 50-ampere circuit breaker located in the power
distribution panel under the floor forward of the main spar.
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PILOT TIP
Erratic operation of the magnetic compass may occur while windshield heat is being used. To prolong the life of the windshield, turn on the windshield heat climbing through 10,000' and turn
it off passing 10,000 feet in the descent unless in icing conditions below 10,000. If in icing conditions, the windshield heat should be on.
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.
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Airplane serials BB-991 and prior are equipped with dual heating element deice boots. One
element is for deicing the inner portion of the propeller blade and the other element deices the
outer portion of the deicer blade. Power is cycled by the deicer timer to these heating elements in
the following sequence: RH outboard, RH inboard, LH outboard and LH inboard. Each sequence
has a 34-second duration and completes a full cycle every two minutes and sixteen seconds.
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.
Airplane serials BB-992 and after are equipped with improved single heating element deicer
boots. 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. A manual propeller deicer system is provided as a backup to the automatic system. A
control switch located on the inboard LH subpanel controls the manual override relays. The
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switch on airplane serials BB-991 and prior is placarded PROP-INNER-OUTER. When the
switch is in the outer position, power is supplied to the outer heating elements of both propellers.
When the switch is moved to the inner position, power is supplied to the inner heating elements
of both propellers. The manual over-ride switch on airplane serials BB-992 and after is placarded
PROP-MAN-OFF. When the switch is in the MAN position, power is supplied to the entire deice
surface of both props. The manual override switch is of the momentary type and must be held in
place until the ice has been dislodged from the propeller surface. Because the MANUAL mode
bypasses the timer, the MANUAL deice system must be released after 90 seconds of operation.
The load meters will indicate approximately a 0.5 increase of load when the manual propeller
deicer system is in operation. The propeller ammeter will not indicate any load in the manual
mode of operation.
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.
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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 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.
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.
Each pneumatic fuel control line is protected against ice by an electrically heated jacket. Power
is supplied to each fuel control air line jacket heater by two switches actuated by moving the
condition levers in the pedestal out of the fuel cutoff range. Fuel control heat is automatically
turned on for all flight operations and requires no action by the pilot.
ANTI-ICING SYSTEMS LIMITATIONS
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Minimum Ambient Temperature for Operation of Deicing Boots -40°C
Minimum Airspeed for Sustained Icing Flight -140 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.
Once the manual override system is activated (i.e., anytime the ICE VANE EMERGENCY
MANUAL EXTENSION handle has been pulled out), do not attempt to operate the ice vanes
electrically until the override assembly inside the engine cowling has been properly reset on the
ground. Even after the manual extension handle has been pushed back in, the manual override
system is still engaged.
ANTI-ICE SYSTEM EMERGENCY PROCEDURES
NONE
ANTI-ICE SYSTEM ABNORMAL PROCEDURES
ELECTROTHERMAL PROPELLER DEICE (Auto System)
Abnormal Readings on Deice Ammeter. (Normal Operation: 14 to 18 amps)
1) Zero Amps:
a) Prop Deice - CHECK AUTO
b) If OFF, reposition to AUTO after 30 seconds.
c) If in AUTO position with zero amps reading, system is inoperative: position the switch
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to OFF.
d) Use manual backup system. (No deice ammeter indication - monitor loadmeter)
2) Below 14 amps:
a) Continue operation.
b) If propeller imbalance occurs, increase RPM briefly to aid in ice removal.
3) Over 18 amps:
a) If the Auto Prop Deice circuit breaker switch does not trip, continue operation.
b) If propeller imbalance occurs, increase RPM briefly to aid in ice removal.
c) If the Auto Prop Deice circuit breaker switch trips, use the manual system. Monitor
loadmeter for excessive current drain.
d) If the Prop Deice Control circuit breaker or the Left or Right Prop Deice circuit
breaker trips, avoid icing conditions.
ELECTROTHERMAL PROPELLER DEICE (Manual System)
On Airplanes Prior to BB-992:
1) To use manual system, hold switch in OUTER position for approximately 30 seconds,
then in INNER position for approximately 30 seconds.
2) Monitor manual system current requirement using the airplane's loadmeters when the
switch is in OUTER or INNER. A small needle deflection (approximately 5%) indicates
the system is functioning.
Airplanes BB-992 and After:
3) To use manual system, hold manual propeller deice switch in MANUAL position for
approximately 90 seconds, or until ice is dislodged from blades. Monitor manual system
current requirement with the airplane's loadmeters when the manual deice switch is in the
MANUAL position. A small needle deflection (approximately 5%) indicates the system
is functioning.
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ENGINE ICE VANE-FAILURE (L or R ICE VANE Annunciator)
1) Ice Vane Control Circuit Breaker - PULL
2) Airspeed - 140 - 160 KIAS
3) Manual Extension Handle - PULL OUT (ICE VANE EXT annunciator Illuminated)
4) Airspeed - RESUME If ICE VANE EXT Annunciator Does Not Illuminate:
5) Exit icing conditions.
6) Manual Extension Handle - PUSH IN (to retract vanes when required)
CAUTION
DO NOT ACTIVATE ICE VANES ELECTRICALLY ONCE THE MANUAL SYSTEM HAS BEEN USED UNTIL THE OVERRIDE LINKAGE HAS BEEN RESET AFTER
LANDING.
NOTE
The ICE VANE fail annunciator will be illuminated any time the position of the ice vane does not match the corresponding switch position. The switch may be repositioned to match the vane
position without damaging the linkage as long as the Ice Vane Control circuit breaker is out.
ANTI-ICE SYSTEM EXPANDED PROCEDURES
BRAKE DEICE CHECK
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1) Power Levers 1,800 RPM (NOTE ITT) 2) Brake Deice Switch ON (DEICE LIGHT ILLUMINATED)
3) Left and Right ITT SLIGHT INCREASE
4) Brake Deice Switch OFF (ITT RETURN TO VALUE IN STEP 1)
ENGINE ICE VANES CHECK
1) Power Levers 1,800 RPM
2) Ice Vane Switches EXTENDED
3) Torque Drop CHECKED
4) Ice Vane Extended Lights ILLUMINATED
5) Ice Vane Bypass Door EXTENDED
6) Ice Vane Switches AS REQUIRED
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ANTI-ICE 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 shatter a cold soaked windshield.
2) Use the inertial separators whenever the temperature is ___ and ___is present.
3) True or False: Use of flaps in icing condition is prohibited.
4) Minimum speed for flight in icing conditions is __K.
5) Brake deice will terminate automatically:
a) 15 minutes after gear retraction.
b) 10 minutes after gear retraction.
c) Does not terminate until switch is turned off.
d) After gear is cycled.
6) True or False: The wing and tail boots sequence at the same time in the CYCLE position.
7) 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.
8) The deice boots should not be cycled if the outside air temperature is below:
a) -50ºC
b) -40ºC
c) -40ºF
d) -30ºC
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9) True or False: Continuous use of the pitot on the ground is recommended.
10) 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.
11) Define icing conditions.
12) Under what conditions is auto ignition required to be armed?
13) Under what conditions might you not want auto ignition to be armed?
14) Describe the working principle of the inertial separators (“ice vanes”).
15) How would you know if the inertial separators have actually lowered?
16) True or False: Damage will occur if windshield heat is used on the ground.
17) What caution should be considered regarding the use of windshield heat?
18) Under what conditions could the stall warning system be inaccurate?
19) On certain aircraft, should the inertial separators be operated electrically after the manual
system has been engaged?
20) How can you check that the propeller deice timer is working correctly?
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CHAPTER 11
FLIGHT CONTROLS OBJECTIVES
After completing this chapter, you will 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.
6) Describe the rudder boost 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.
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Rudder/Trim Control Cables
Elevator/Trim Control Cables
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.
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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. 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.
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PILOT TIP
Do not allow the trim system to move past 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 installed. 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.
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With the tow bar 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 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 King Air is equipped with Fowler type flaps that extend down and aft. The 200 knot
operational speed limit for flaps provides for easy traffic pattern transition. Flaps are selectable
to 3 positions: up, approach (14 degrees), and down (35 degrees). If a go-around is initiated with
flaps fully extended, retraction to either approach or full up positions can be accomplished with a
single switch position selection. 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
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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 jack- screw actuator at each flap. A split flap safety
mechanism for each pair of flaps is provided to disconnect power to the electric motor in the
event of any flap panel to be approximately three to six degrees out-of-phase with the other flaps.
On aircraft BB-2 through BB-1438, the flaps are operated by a sliding switch lever located just
below the condition levers. Flap travel, from 0% (fully up 0°) to 100% (fully down 35°) 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 (40% 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 DOWN 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
40%. 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 BB-1439 and later, all 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. However unlike the earlier models, the flaps cannot be
stopped in between any of the three positions. The flap motor is protected by a 20-ampere flap
motor circuit breaker (placarded FLAP MOTOR) located on the left circuit breaker panel below
the fuel control panel. A 5-ampere circuit breaker placarded FLAP CONTROL is also located on
this panel. This circuit provides power for the flap position indicator and the split-flap safety
mechanism.
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YAW DAMPER
The Yaw Damper system is
designed to provide the pilot with
help in maintaining directional
control and increase ride comfort.
The system can be used at any
altitude but must be operational
above 17,000 feet. The system is
normally incorporated in the
autopilot. Operating instruction can
be found in the Flight Manual
Supplement.
STALL WARNING SYSTEM
The stall warning system provides precise pre-stall warning to the pilot by activating the warning
horn when excessive angles of attack are reached. The activation level of the horn is changed by
the flap position.
STALL WARNING ACTIVATES
5-13 Knots above stall in Clean Configuration
5-12 Knots above stall with Flaps 40%
8-14 Knots above stall at Flaps 100%
The stall warning system consists of the following major components:
1) The lift computer
2) A stall warning horn
3) A squat switch (LH only)
4) A stall warning test switch
5) A five-amp circuit breaker (furnishing power for the system)
6) A lift transducer
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The stall warning horn will not sound when the full weight of the aircraft is on the landing gear
because the landing gear squat switch opens the stall warning horn circuit; consequently, moving
the stall warning vane up during preflight does not sound the warning horn. When the weight of
the aircraft is off the landing gear, the squat switch closes the circuit so that the warning horn can
be actuated by an incipient stall. The system has a heater that can be selected by the pilot prior to
entering icing conditions.
RUDDER BOOST
A rudder boost system is provided to aid the pilot in maintaining directional control in the event
of an engine failure or a large variation of power between the engines. Incorporated into the
rudder cable system are two pneumatic rudder-boosting servos that actuate the cables to provide
rudder pressure to help compensate for asymmetrical thrust. During operation, a differential
pressure valve accepts bleed air pressure from each engine. If the pressure varies between the
bleed air systems, the shuttle valve in the differential pressure valve moves toward the low
pressure side. As the pressure difference reaches a preset tolerance, a switch on the low pressure
side closes, activating the rudder boost system. The system is designed only to help compensate
for asymmetrical thrust. Appropriate trimming is to be accomplished by the pilot. Moving either
or both of the bleed air valve switches on the copilot's subpanel to the INSTR & ENVIR OFF
position will disengage the rudder boost system. The system is controlled by a toggle switch,
placarded RUDDER BOOST - ON - OFF, and located on the pedestal below the rudder trim
wheel. The switch is to be turned ON before flight. A preflight check of the system can be
performed during the run-up by retarding the power on one engine to idle and advancing power
on the opposite engine until the power difference between the engines is great enough to close
the switch that activates the rudder boost system. Movement of the appropriate rudder pedal will
be noted when the switch closes, indicating the system is functioning properly for low engine
power on that side. Repeat the check with opposite power settings to check for movement of the
opposite rudder pedal. The rudder boost system may not operate if the Brake Deice system is
active.
FLIGHT CONTROL LIMITATIONS MANEUVER LIMITS
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The BEECHCRAFT Super King Air B200 and B200C are Normal Category Airplanes.
Acrobatic maneuvers, including spins, are prohibited.
FLIGHT LOAD FACTOR LIMITS
FLIGHT CONTROL EMERGENCY PROCEDURES
BOLD TYPE INDICATES MEMORY ITEMS!
FLIGHT CONTROLS
UNSCHEDULED ELECTRIC ELEVATOR TRIM
1) Airplane Attitude - MAINTAIN (using elevator control)
2) Control Wheel Disconnect Switch - DEPRESS FULLY (2nd level, ELEC TRIM OFF annunciator -ILLUMINATED)
NOTE
Autopilot will disengage when the disconnect switch is depressed.
3) Manually retrim airplane.
4) Elevator Trim - OFF
CAUTION
DO NOT REACTIVATE ELECTRIC TRIM SYSTEM UNTIL CAUSE OF MALFUNCTION HAS BEEN DETERMINED.
UNSCHEDULED RUDDER BOOST ACTIVATION
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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 - INSTR & ENVIR OFF
5) Rudder Trim - AS REQUIRED
6) Perform normal landing.
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
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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 FLAPS-UP LANDING.
6) Landing Gear - DN
7) Lights - AS REQUIRED
NOTE
Under low visibility conditions, landing and taxi lights should be left off due to light reflections.
8) Radar - AS REQUIRED
9) 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:
10) Approach Speed - ESTABLISHED
11) Yaw Damp - OFF
12) Propeller Levers - FULL FORWARD
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13) Power Levers - IDLE
After Touchdown:
14) Power Levers - LIFT AND SELECT REVERSE
15) Brakes - AS REQUIRED
FLIGHT CONTROLS EXPANDED PROCEDURES
OVERSPEED GOVERNOR/RUDDER BOOST TEST
1) Rudder Boost Switch ON
2) Propeller Levers FULL FORWARD
3) Propeller Test Switch HOLD TO TEST
4) Left Power Lever 1,800 RPM 5) Left Overspeed Governor/Rudder Boost CHECK (1,870 ± 40)
6) Left Power Lever IDLE
7) Right Power Lever 1,800 RPM
8) Right Overspeed Governor/Rudder Boost CHECK (1,870 ± 40)
9) Propeller Test Switch RELEASED
Electric Elevator Trim
1) Verify that the ELEV TRIM switch is on.
2) Check operation of the dual-element thumb switches.
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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.
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FLIGHT CONTROLS
QUESTIONS 1) Is rudder boost required to be operative for flight?
2) What may be the result if rudder boost and brake deice are used at the same time?
3) True or False: The rudder boost system may be tested by advancing the power levers and
turning off one bleed air control switch.
4) Where is the rudder boost switch located?
5) List the maximum flap air speeds:
a) Approach flaps __ KCAS.
b) Full flaps __ KCAS.
6) Explain how to select 60% flaps.
7) In what range could you not select intermediate flaps?
8) Where is the circuit breaker located for flap motor power? How about the control circuit?
9) Refer to the emergency procedures. List the procedures for the flap system.
10) Is any one of the four flap segments different than the others?
11) Where is the aileron trim tab located?
12) Where is the electric trim switch located?
13) True or False: The flaps have no asymmetrical protection.
14) The yaw damper must be operational above what altitude?
15) True or False: The flight controls are hydraulically operated.
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16) The wing flaps are:
a) Fowler
b) Split
c) Plain
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CHAPTER 12
PITOT STATIC SYSTEM OBJECTIVES
After completing this chapter, you 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
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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 SYSTEM LIMITATIONS NONE
PITOT STATIC SYSTEM EMERGENCY PROCEDURES
NONE
PITOT STATIC SYSTEM ABNORMAL PROCEDURES
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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:
1) Pilot's Static Air Source (right side panel) - ALTERNATE
2) 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.
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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?
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CHAPTER 13
OXYGEN SYSTEM OBJECTIVES
After completing this chapter, you 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
A push/pull handle (PULL ON - System READY), located aft of the overhead light control
panel, is used in conjunction with the automatically deployed passenger oxygen system. 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 primary oxygen supply line delivers oxygen to
the two crew oxygen outlets in the cockpit, to the first aid oxygen outlet in the toilet area, and to
the passenger oxygen system shutoff valve. The crew is provided with diluter-demand, quick-
donning oxygen masks. These masks hang on the aft cockpit partition behind and outboard of the
pilot and copilot seats. They are held in the armed position by spring-tension clips, and can be
donned immediately with one hand. The diluter-demand crew masks deliver oxygen to the user
only upon inhalation. Consequently, there is no loss of oxygen when the masks are plugged in
and the PULL - ON - System READY handle is pulled out, even though oxygen is immediately
available upon demand. A small lever on each diluter-demand oxygen mask permits the selection
of two modes of operation: NORMAL and 100%. In the NORMAL position, air from the cockpit
is mixed with the oxygen supplied through the mask. This reduces the rate of depletion of the
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oxygen supply, and it is more comfortable to use than 100% aviators breathing oxygen.
However, in the event of smoke or fumes in the cockpit, the 100% position should be used to
prevent the breathing of contaminated air. For this reason, the selector lever should be left in the
100% position when the masks are not in use. Anytime the primary oxygen supply line is
charged, oxygen can be obtained from the first aid oxygen mask located in the toilet area, by
manually opening the overhead access door (placarded FIRST AID OXYGEN - PULL) and
opening the ON-OFF valve inside the box. A placard (NOTE: CREW System MUST BE ON)
reminds the user that the PULL ON - System READY handle in the cockpit must be pulled out
before oxygen will flow from the first aid oxygen mask. The passenger oxygen system is of the
constant flow type. Anytime the cabin pressure altitude exceeds approximately 12,500 feet, a
barometric-pressure switch automatically energizes a solenoid which opens the passenger
oxygen system shut-off valve. The pilot can open the valve manually anytime by pulling out the
PASSENGER MANUAL Over-RIDE handle, located aft of the overhead light control panel.
Once the passenger oxygen system shut-off valve has been opened (either automatically or
manually), oxygen will flow into the passenger oxygen supply line, if the primary oxygen system
line has been charged (i.e., if the oxygen supply bottle contains oxygen and the PULL ON -
System READY handle in the cockpit is pulled out). When oxygen flows into the passenger
oxygen system supply line, a pressure-sensitive switch in the line closes a circuit to illuminate
the green PASS OXYGEN ON annunciator on the cautionary/ advisory annunciator panel. This
switch will also cause the cabin lights (all fluorescent lights, the foyer light and the center
baggage compartment light) to illuminate in the full bright mode, regardless of the position of the
interior lights switch placarded CABIN LIGHTS - START BRIGHT - DIM -OFF located on the
copilot's left subpanel. The pressure of the oxygen in the passenger oxygen system supply line
then automatically extends a plunger against each of the passenger oxygen mask dispenser doors,
forcing the doors open. The oxygen masks then drop down about 9 inches below the dispensers.
The lanyard valve pin at the top of the oxygen mask hose must be pulled out in order for oxygen
to flow from the mask. The pin is connected to the oxygen mask via a flexible cord; when the
oxygen mask is pulled down for use, the cord pulls the pin out of the lanyard valve. The lanyard
valve pin must be manually reinserted into the valve in order to stop the flow of oxygen when the
mask is no longer needed. The passenger oxygen can be shut off and the remaining oxygen
isolated to the crew and first aid outlets by pulling the OXYGEN CONTROL circuit breaker in
the ENVIRONMENTAL group on the right side panel, providing the PASSENGER MANUAL
O'RIDE handle is pushed in to the OFF position
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AUTO DEPLOYMENT PASSENGER OXYGEN SYSTEM
The auto deployment passenger oxygen system is operated by two push-pull control cables and a
barometric pressure switch. The push-pull control cables are located overhead between the pilots.
On airplanes BB-1444 and after, the push-pull control cables are located on the sides of the
control pedestal. The left control cable operates the oxygen system shutoff valve and places the
system in the ready mode when the knob is pulled. If this handle is pushed in, no oxygen supply
is available anywhere in the airplane. The right cable is the passenger manual-override control to
the shutoff valve that manually turns the passenger oxygen on or off. This valve is normally in
the OFF position and will not be used unless the barometric pressure switch fails to operate when
the cabin depressurizes. The barometric pressure switch automatically releases passenger oxygen
and deploys the passenger oxygen masks when the cabin altitude reaches 12,500 feet. The
released oxygen pressure actuates a plunger in each of the oxygen auto deployment boxes which
causes the dispenser door to open and drop the oxygen masks. After the masks are deployed, the
oxygen valve lanyard pin must be pulled for oxygen to flow to each mask. When the masks are
no longer required, the lanyard pin is reinserted to stop the flow of oxygen. After operation by
the barometric pressure switch, the passenger oxygen can be shut off by pulling the oxygen
control circuit breaker. This will limit the remaining oxygen to the crew and first aid outlets.
OXYGEN CYLINDERS
The Auto Deployment Oxygen System uses steel
oxygen cylinders that are available in four sizes. The
standard system utilizes the 22-cubic-foot cylinder and
the optional systems use the 49-, 64-or the 76-cubic-
foot cylinder. The regulators for these cylinders
provide a constant flow of 200 LPM at a pressure of 70
psi. 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.
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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 PRESSURE-SENSE SWITCH
The oxygen pressure-sense switch is located in the passenger oxygen line in the aft cabin ceiling.
When the passenger manual-override shutoff valve is opened, oxygen pressure is released to the
oxygen mask overhead containers and to the pressure-sense switch. The actuated pressure-sense
switch will illuminate the PASS OXY ON annunciator in the instrument panel advising the crew
that the masks are deployed and oxygen is available to the passengers.
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Auto Deployment Oxygen System Installation
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OXYGEN SYSTEM 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 SYSTEM EMERGENCY 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
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22,000 feet 5 to 10 minutes
12- 18,000 feet 30 minutes or more
1) Oxygen System Ready - PULL ON (verify)
2) Crew (Diluter Demand Masks) - DON MASKS 3) Mic Selector - OXYGEN MASK
4) Audio Speaker - ON
5) Passenger Manual Drop Out - PULL ON
6) Passengers - PULL LANYARD PIN, DON MASK
7) Oxygen Duration - CONFIRM (See OXYGEN SYSTEM in Section IV, NORMAL PROCEDURES for duration tables)
8) First Aid Oxygen - AS REQUIRED
a) Oxygen Compartment - PULL OPEN
b) ON/OFF Valve – ON
c) Mask – DON
AUTO-DEPLOYMENT OXYGEN SYSTEM FAILURE (ALT WARN Annunciator Illuminated, PASS OXY ON Annunciator Not Illuminated)
1) Passenger Manual Drop Out - PULL ON
2) First Aid Mask (if required) - DEPLOY MANUALLY To Isolate Oxygen Supply to the Crew and First Aid Mask:
3) Oxygen Control Circuit Breaker - PULL
4) Passenger Manual Drop Out - PUSH OFF
OXYGEN SYSTEM ABNORMAL PROCEDURES NONE
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OXYGEN SYSTEM
QUESTIONS 1) Why is it unnecessary to remove the oxygen filler valve access plate (on the right rear
fuselage) to check oxygen system pressure?
2) What is the normal system pressure for a full bottle?
3) List some precautions to observe during oxygen purging or 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) True or False: It is acceptable to use medical oxygen if aviator's breathing oxygen is not
available.
7) True or False: If the passenger oxygen masks dropped, the lanyard valve pin at the top of the
oxygen mask hose must be pulled out in order for oxygen to flow from the mask.
8) At what cabin altitude will the passenger masks drop automatically?
9) What is the difference between Normal and 100% on the crew masks?
10) Will pulling the passenger manual over-ride handle turn on the cabin lights?
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CHAPTER 14
POWER SETTINGS AND PROFILES
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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
Reverse is most effective at higher speeds and braking is most effective at lower speeds.
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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
Reverse is most effective at higher speeds and braking is most effective at lower speeds.
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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
Reverse is most effective at higher speeds and braking is most effective at lower speeds.
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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
Reverse is most effective at higher speeds and braking is most effective at lower speeds.
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