King Air 200 – The Training Workbook OF CONTENTS CHAPTER 1: AIRCRAFT - GENERAL 11 INTRODUCTION TO THE KING AIR 200 AND B200 11 OBJECTIVES

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

King Air 200 – The Training Workbook 11

FOR TRAINING PURPOSES ONLY www.KingAirTraining.com

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

12 King Air 200 – The Training Workbook


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.



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.



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.



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



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



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.



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.



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.



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



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.



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



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.



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.



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



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



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



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.



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)



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



AIRSPEED LIMITATIONS



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.



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



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.



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.



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 ________.



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.



CHAPTER 2

ELECTRICAL SYSTEM

OBJECTIVES


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



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



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.



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



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.



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



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



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



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



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.



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%



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








The following steps should be considered mandatory in any emergency:



1) Fly the airplane.

2) Identify the emergency.

3) Complete the appropriate checklist.


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.



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)



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:



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



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



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?



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?



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.



CHAPTER 3

ANNUNCIATOR SYSTEM

OBJECTIVES


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



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.



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



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)



CHAPTER 4

FUEL SYSTEM

OBJECTIVES


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



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.



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



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



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



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



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.



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.



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



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.



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)



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










The following steps should be committed to memory and considered mandatory in any

emergency:





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



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



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.


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.



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



c) 3 hours

d) 4 hours



CHAPTER 5

ENGINE SYSTEM

OBJECTIVES


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



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



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.

:



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.



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



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.



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.



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



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.



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.



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



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



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.



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



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.



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



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



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.



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



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 unsafe; 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.



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



EMERGENCY ENGINE SYSTEM PROCEDURES The pilot in command of an aircraft is directly responsible for and is the final authority as to the








The following steps should be committed to memory and considered mandatory in any

emergency:





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



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:


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



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)



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


4) Conduct Air Start Procedures.

NOTE

The propeller will not unfeather without engine operating.



ENGINE OUT GLIDE

1) Landing Gear – UP

2) Flaps - UP

3) Propellers - FEATHERED

4) Airspeed - 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.



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


7) Propeller Lever (inoperative engine) - LOW RPM


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



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)


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



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



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



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



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.


2) Ignition and Engine start - STARTER ONLY (for a maximum ol40 seconds)

3) Ignition and Engine Start – OFF



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.



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?



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


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.



CHAPTER 6

PROPELLERS

OBJECTIVES


1) Identify the major components of the propeller system.

2) Describe the operation of the propeller governor, overspeed governor and the fuel topping governor.

3) Explain onspeed, overspeed and underspeed conditions.

4) Describe the feathering process.

5) Explain the use of "Beta".

6) Explain the autofeather system and describe its operation.

7) Understand 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



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.



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



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



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,



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.



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.



PILOT TIP

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

PROPELLER SYNCHROPHASER

The Type I propeller synchrophaser automatically matches the right slave propeller and

maintains the blades of one propeller at a predetermined position relative to the blades of the

other propeller. To prevent the right propeller from losing excessive RPM if the left propeller is

feathered while the synchrophaser is on, the synchrophaser is limited to approximately ±30 RPM

from the manual prop control setting. Normal governor operation is unchanged but the

synchrophaser will continuously monitor propeller RPM and reset the governor as required. A

magnetic pickup mounted in each propeller overspeed governor transmits electric pulses to a

transistorized control box. The control box converts any pulse rate differences into correction

commands, which are transmitted to an actuator motor. The motor then trims the right propeller

governor through a flexible shaft to exactly match the left propeller. A toggle switch, installed on



the instrument panel, turns the system on. With the switch off, the actuator automatically runs to

the center of its range of travel before stopping to assure that when next turned on the control

will function normally. To operate the system, synchronize the propeller in the normal manner

and turn the synchrophaser on. The right propeller RPM and phase will automatically be adjusted

to correspond with the left. To change RPM, adjust both propeller controls at the same time. This

will keep the right governor setting within the limiting range of the left propeller. If the

synchrophaser is on but is unable to adjust the right propeller to match the left, the actuator has

reached the end of its travel. Turn the synchrophaser switch off (allowing the actuator to run to

the center of its range and the right propeller to be governed by the propeller lever), synchronize

the propellers manually and turn the synchrophaser switch on.

The Type II propeller synchrophaser system automatically matches the RPM of both propellers

as a result of maintaining a specific phase relationship between the blades of the left and right

propellers. The control box senses pulses which are generated by pickups mounted at identical

locations on both engines. Ferrous metal targets, mounted on the propeller spinner bulkheads,

provide the pulse reference for the pickups. Adjusting the RPM's of the propellers is

accomplished by the control box with correction commands to each propeller governor. The

governor servo can increase but never decrease the speed set by the propeller control lever. The

RPM of one propeller will follow the changes in RPM of the other propeller over the

predetermined holding range of the governor (approximately 25 RPM). This limited holding

range prevents either propeller from losing more than a limited RPM if the RPM of the other

propeller is manually reduced, such as in power changes or propeller feathering, while the

synchrophaser is on. The synchrophaser system is controlled through a toggle switch placarded

PROP SYNCH-ON-OFF. To operate the system, synchronize the propellers in the normal

manner and turn the synchrophaser on. To change RPM, adjust both propellers at the same time.

This will keep the setting within the holding range of the system. If the synchrophaser is on, but

will not synchronize propellers, the propeller speeds are not within the limits required for the

system to assume control. Turn the synchrophaser off, synchronize the propellers manually, and

then turn the synchrophaser on.



PROPELLER 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)



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.



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



c) 2080 RPM

d) 1600 RPM



CHAPTER 7

PRESSURIZATION AND ENVIRONMENTAL SYSTEMS

OBJECTIVES


1) Identify the components in the pressurization system.

2) Explain the operation of the pressurization system.

3) Recognize pressurization system emergencies.

4) Identify the components in the environmental system.

5) Explain the operation of the heating and air conditioning system.

6) Explain the operation of the emergency oxygen system.

INTRODUCTION

This chapter describes the operation of the pressurization and environmental systems.

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.



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.



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



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



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.



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



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,



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



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



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



Each flow control unit consists of an ejector and an integral bleed air modulating valve, firewall

shutoff valve, and a check valve that prevents the bleed air from escaping through the ambient

air intake. The flow of bleed air through the flow control unit is controlled as a function of

atmospheric pressure and temperature. Ambient air flow is controlled as a function of

temperature only. When the bleed air valve switches on the co-pilot's left subpanel are turned on,

a bleed air shutoff electric solenoid valve on each flow control unit opens to allow the bleed air

into the unit. As the bleed air enters the flow control unit, it passes through a filter before going

to the reference pressure regulator. The regulator will reduce the pressure to a constant value of

18 to 20 psi. This reference pressure is then directed to the various components within the flow

control unit that regulate the output to the cabin. One reference pressure line is routed to the

firewall shutoff valve located downstream of the ejector. A restrictor is placed in the line

immediately before the shutoff valve to provide a controlled opening rate. At the same time, the

reference pressure is directed to the ambient air modulating valve located upstream of the ejector

and to the ejector flow control actuator. A pneumatic thermostat with a variable orifice is



connected to the modulating valve. This pneumostat is located on the lower aft side of the

fireseal forward of the firewall. The 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



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



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.



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


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


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



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)


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



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)


d) Audio Speaker - ON


3) Vent Blower - AUTO 4) Aft Blower (if installed) - OFF

5) Avionics Master - OFF

6) Nonessential Electrical Equipment – OFF

If Fire or Smoke Ceases:

7) Individually restore avionics and equipment previously turned off.

8) Isolate defective equipment.

WARNING

DISSIPATION OF SMOKE IS NOT SUFFICIENT EVIDENCE THAT A FIRE HAS BEEN EXTINGUISHED. IF IT CANNOT BE VISUALLY CONFIRMED THAT NO

FIRE EXISTS, LAND AT THE NEAREST SUITABLE AIRPORT.

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



9) Cabin Pressure - DUMP

10) Land at the nearest suitable airport.

NOTE

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

ENVIRONMENTAL SYSTEM SMOKE OR FUMES

1) Oxygen

a) Oxygen System Ready - PULL ON (Verify)

b) Crew (Diluter Demand Masks) - DON MASKS (100% position)


d) Audio Speaker – ON


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.



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


d) Audio Speaker - ON

e) Passenger Manual Drop Out - PULL ON

f) Passengers - PULL LANYARD PIN, DON MASK

2) Power Levers - IDLE


4) Flaps - APPROACH


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:



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




2) Oxygen System Ready - PULL ON

3) Crew Diluter Demand Masks - DON MASK, CHECK FIT AND OPERATION, AND STOW

4) Oxygen Duration - DETERMINE



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?



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



CHAPTER 8

LANDING GEAR, TIRES AND BRAKE SYSTEM

OBJECTIVES


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.



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



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



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



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.



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



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



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



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.



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.



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



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



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.



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


NONE



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:



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:



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.)


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)




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.


10) The gear warning horn will sound when the gear is not down and:

a) Either power lever is reduced to a certain setting.



b) The wing flaps are extended beyond the approach setting.

c) The hydraulic system pressure falls below 1,500 psi.

d) Both a and b.

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?



CHAPTER 9

PNEUMATIC AND VACUUM SYSTEM

OBJECTIVES


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



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



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



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.



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


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



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



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.



CHAPTER 10

ANTI-ICE SYSTEM

OBJECTIVES


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



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.



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



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,



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



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



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.



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.



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.



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



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.



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



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



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.



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



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



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



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?



CHAPTER 11

FLIGHT CONTROLS OBJECTIVES


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.



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.



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.



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.



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



through a gearbox mounted on the forward side of the rear spar. The motor incorporates a

dynamic braking system which helps to prevent overtravel of the flaps. The gearbox drives four

flexible drive shafts connected to a jackscrew actuator at each flap. 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.



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



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



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


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



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



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.


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




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.



WARNING

OPERATION OF THE ELECTRIC TRIM SYSTEM SHOULD OCCUR ONLY BY MOVEMENT OF PAIRS OF SWITCHES. ANY MOVEMENT OF THE ELEVATOR TRIM WHEEL WHILE ACTUATING ONLY ONE SWITCH DENOTES A SYSTEM

MALFUNCTION. IF A MALFUNCTION OF THE ELECTRIC TRIM SYSTEM IS INDICATED, ELECTRIC TRIM MUST BE DISENGAGED AND TRIM CHANGES

MADE WITH MANUAL TRIM ONLY.



FLIGHT CONTROLS

QUESTIONS 1) 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.



16) The wing flaps are:

a) Fowler

b) Split

c) Plain



CHAPTER 12

PITOT STATIC SYSTEM OBJECTIVES


1) Identify the major components of the pitot static system.

2) Describe how the pilot and copilot instruments receive pitot and static pressure.

3) Be able to drain the pitot static system.

4) Describe the alternate static source.

PITOT AND STATIC PRESSURE SYSTEM

The pitot and static pressure system provides a source of impact pressure and static air for

operation of selected flight instruments. The pitot portion of the system is comprised of the pitot

mast mounted on each lower side of the nose, the wiring connecting the heating element of the

mast into the electrical system and the tubing between the mast and the airspeed indicators. The

impact pressure entering the masts is transmitted to the dual airspeed indicators mounted on the

instrument panel through separate tubing routed along each upper side of the nose compartment.

Since the pitot mast is the lowest point in each line from

the airspeed indicators, the resultant natural drainage

eliminates the need for drain valves. Two circuit breaker

switches on the left inboard subpanel control the heating

elements that prevent the pitot openings in the mast from

becoming clogged with ice. The static portion of the system

includes two static ports on each side of the fuselage aft of the aft pressure bulkhead. Lines

connect the static ports to the instruments in the crew compartment and an alternate line supplies

static air for the pilot's instruments should the fuselage static ports become obstructed. The static

lines are routed from the static ports to the top center of the fuselage and immediately over to the

right side of the fuselage. They are then routed forward along the fuselage beneath the windows



to the rate-of-climb indicator, altimeter and airspeed indicator at the instrument panel. The static

line drain valves are located behind the access door located in the lower right flight compartment

wall adjacent to the instrument panel. The static lines should be drained any time the aircraft has

been exposed to rain, either on the ground or during flight. Should abnormal or erratic instrument

readings indicate that the normal static source is restricted; the alternate air source may be

utilized. This alternate system supplies static air from the interior of the aft fuselage. The

alternate static air line is routed through the aft pressure bulkhead forward along the right side of

the fuselage to the static air selector valve. This selector valve is located below the copilot's

circuit breaker panel adjacent to the instrument panel. The static air selector valve is held in the

normal position by a clip. The alternate air source is selected by raising the clip and moving the

toggle from NORMAL to ALTERNATE. The pilot's instruments then function on the alternate

air source.

OUTSIDE AIR TEMPERATURE

The outside air temperature indicator is installed in the pilot's overhead panel or the pilot's left

sidewall panel. The indicator dial is on the inside of the compartment with the stem of the

instrument protruding through the skin of the airplane. The instrument is hermetically sealed

against dust and moisture.

The instrument consists of a bimetal element which is attached to the staff and pointer. A hollow

stainless steel stem encloses the element. A sunshield is installed over the stem for protection.

PITOT STATIC SYSTEM LIMITATIONS NONE

PITOT STATIC SYSTEM EMERGENCY PROCEDURES

NONE

PITOT STATIC SYSTEM ABNORMAL PROCEDURES



PILOT'S ALTERNATE STATIC AIR SOURCE

THE PILOT'S ALTERNATE STATIC AIR SOURCE SHOULD BE USED FOR

CONDITIONS WHERE THE NORMAL STATIC SOURCE HAS BEEN OBSTRUCTED.

When the airplane has been exposed to moisture and/or icing conditions (especially on the

ground), the possibility of obstructed static ports should be considered. Partial obstructions will

result in the rate of climb indication being sluggish during a climb or descent. Verification of

suspected obstruction is possible by switching to the alternate system and noting a sudden

sustained change in rate of climb. This may be accompanied by abnormal indicated airspeed and

altitude changes beyond normal calibrated differences.

Whenever any obstruction exists in the Normal Static Air System, or when the Alternate

Static Air System is desired for use:

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.



PITOT STATIC SYSTEM

QUESTIONS

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

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

indicators.

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

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

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

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



CHAPTER 13

OXYGEN SYSTEM OBJECTIVES


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



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



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.



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.



Auto Deployment Oxygen System Installation



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





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


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)


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


OXYGEN SYSTEM ABNORMAL PROCEDURES NONE



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?



CHAPTER 14

POWER SETTINGS AND PROFILES



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.



CAUTION



PILOT TIP






CAUTION



PILOT TIP




CAUTION



PILOT TIP








Documents

King Air 200 – The Training Workbook OF CONTENTS CHAPTER 1: AIRCRAFT - GENERAL 11 INTRODUCTION TO THE KING AIR 200 AND B200 11 OBJECTIVES