B200 Pilot Training Manual

8/20/2019 B200 Pilot Training Manual

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SUPER KING AIR 200/B200PILOT TRAINING MANUAL



Raytheon Learning CenterFlightSafety International9720 East Central AvenueWichita, KS 67206-2595(316) 612-5300

(800) 488-3747

Long Beach Learning CenterFlightSafety InternationalLong Beach Municipal Airport4330 Donald Douglas DriveLong Beach, CA 90808(562) 938- 0100

(800) 487-7670

Atlanta Learning CenterFlightSafety International1010 Toffie TerraceAtlanta, GA 30354(678) 365-2700(800) 889-7916

Lakeland Learning CenterFlightSafety InternationalLakeland Airport2949 Airside Center Dr.Lakeland, FL 33811(863) 646 5037

Toledo Learning CenterFlightSafety InternationalToledo Express Airport11600 West Airport Services Rd.Swanton, OH 43558

(419) 865-0551(800) 497-4023

Houston Learning CenterFlightSafety InternationalWilliam P. Hobby Airport7525 Fauna at Airport Blvd.Houston, TX 77061

(713) 393-8100(800) 927-1521

Paris Learning CenterFlightSafety InternationalBP 25Zone d’Aviation d’AffairesBuilding 404, Aeroport du Bourget

Le Bourget, CEDEXFRANCE+33 (1) 49-92-1919

Courses for the Super King Air 200 and other King Air products are taught at the followingFlightSafety Learning Centers:



CONTENTS

SYLLABUS

Chapter 1 AIRCRAFT GENERAL

Chapter 2 ELECTRICAL POWER SYSTEMS

Chapter 3 LIGHTING

Chapter 4 MASTER WARNING SYSTEMChapter 5 FUEL SYSTEM

Chapter 6 AUXILIARY POWER UNIT

Chapter 7 POWERPLANT

Chapter 8 FIRE PROTECTION

Chapter 9 PNEUMATICS

Chapter 10 ICE AND RAIN PROTECTION

Chapter 11 AIR CONDITIONING

Chapter 12 PRESSURIZATION

Chapter 13 HYDRAULIC POWER SYSTEMS

Chapter 14 LANDING GEAR AND BRAKES

Chapter 15 FLIGHT CONTROLS

Chapter 16 AVIONICS

Chapter 17 MISCELLANEOUS SYSTEMS

Chapter 18 WEIGHT AND BALANCE/PERFORMANCE

GENERAL PILOT INFORMATION

APPENDIX



FOR TRAINING PURPOSES ONLY

NOTICE

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

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

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



CHAPTER 1AIRCRAFT GENERAL

CONTENTS

Page

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

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

AIRPLANE SYSTEMS........................................................................................................... 1-2

Electrical Power System .................................................................................................. 1-2

Lighting............................................................................................................................ 1-4

Master Warning System................................................................................................... 1-5

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

Powerplants...................................................................................................................... 1-6

Fire Protection.................................................................................................................. 1-8

Bleed-Air System............................................................................................................. 1-8

Ice and Rain Protection .................................................................................................... 1-8

Air Conditioning and Heating.......................................................................................... 1-9

Pressurization................................................................................................................. 1-10

Landing Gear and Brakes............................................................................................... 1-11

Flight Controls ............................................................................................................... 1-13

Pitot and Static Systems................................................................................................. 1-13

Oxygen System.............................................................................................................. 1-15

SUPER KING AIR 200/B200 PILOT TRAINING MANUAL






ILLUSTRATIONSFigure Title Page

1-1 Simplified Electrical System.................................................................................... 1-2

1-2 Electrical Panel......................................................................................................... 1-3

1-3 External Power Socket ............................................................................................. 1-3

1-4 Overhead Light Control Panel (BB-1632 and After) ............................................... 1-4

1-5 Cabin Lights Control Switch (BB-1439, 1444 and After) ....................................... 1-4

1-6 Exterior Lights Control Switches ............................................................................. 1-5

1-7 Fuel Control Panels .................................................................................................. 1-6

1-8 Engine Control Levers.............................................................................................. 1-71-9 Bleed-Air Valve Control........................................................................................... 1-8

1-10 Ice Protection Switches—Pilot’s Subpanel .............................................................. 1-8

1-11 Windshield Wiper Control Switch............................................................................ 1-9

1-12 Cabin Pressurization Controller ............................................................................. 1-11

1-13 Landing Gear Control Panel................................................................................... 1-12

1-14 Manual Extension Controls.................................................................................... 1-12

1-15 Parking Brake Handle ............................................................................................ 1-13

1-16 Flight Control Surfaces .......................................................................................... 1-14

1-17 Trim Tab Controls and Indicators .......................................................................... 1-14

1-18 Flap Control Lever ................................................................................................. 1-14

1-19 Pitot Tubes.............................................................................................................. 1-15

1-20 Static Ports ............................................................................................................. 1-15







INTRODUCTION

This pilot training manual covers all systems on the Super King Air 200 and B200. Chapter 1 provides a general overview of the systems and the structural makeup of the airplane.Throughout this manual there are boxed warnings, cautions, and notes. As indicated inthe Aircraft Fl ight Manual , they are defined as follows: Warnings —Operating proce-dures, techniques, etc., which could result in personal injury or loss of life if not care-fully followed; Cautions —Operating procedures, techniques, etc. , which could resultin damage to equipment if not carefully followed; Note —An operating procedure, tech-nique, etc., which is considered essential to emphasize.

CHAPTER 1AIRCRAFT GENERAL





AIRPLANE SYSTEMSELECTRICAL POWER SYSTEM

GeneralThe airplane electrical system is a 28-VDCsystem, which receives power from a 24-volt,42-ampere hour lead acid gel cell battery

(34/36-ampere hour nickel-cadmium batteryprior to BB-1632), two 250-ampere starter-generators, or through an external power socket.

DC power is supplied to one of the two oper-ating inverters, which provide 400-hertz, 115-volt and 26-volt AC power for various avionics

equipment. (For BB-2 through BB-1483 the26-volt AC also powers the torquemeters.Prior to BB-225 the fuel flow meters are also26-volt AC powered.)

DistributionS o m e m a j o r D C b u s e s a r e a s f o l l o w s(Figure 1-1):

1. Hot Battery Bus

2. Main Battery Bus

3. Left Generator Bus

4. Right Generator Bus

5. Isolation Bus



STARTRELAY

G C U

VOLT / LOADMETER

R/HGEN LINECONTACTOR

R/H STARTER/ GENERATOR

VOLT / LOADMETER

G C U

L/HGEN LINE

CONTACTOR

L/H STARTER/ GENERATOR

STARTRELAY

ISOLATION BUS

MAIN BATT BUSAVIO

AVIO

HOT BUS

SHUNT

BATTRELAY

BATTERY

OFF

BATTSWITCH

ON



6. No. 1 Dual Fed Bus




10. The avionics buses

A hot battery bus is powered by the battery,regardless of the posi tion of the BAT switch.

This bus supplies the engine fire extinguish-ers, firewall shutoff valves, entry and cargol i g h t s , c l o c k s , m o d i f i c a t i o n s , g r o u n dCOMMunications, RNAV memory to older avionics, and standby boost pumps prior to BB-1096. It also powers the battery relay which,in turn, allows power through to the main bat-tery bus, provided that the battery switch is ON(Figure 1-2).

The generators are controlled by GEN 1 and

GEN 2 switches, located under the same gangbar as the BAT switch. Early King Air air-planes do not have the GEN RESET position.Some airplanes have the reset function, butthey are not placarded. When reset is incor-porated (BB-88 and after), the switch mustbe held in GEN RESET for a minimum of one

The four dual-fed buses are powered by either generator bus through a 60-amp limiter, a 70-amp diode, and a 50-amp circuit breaker. Thosefour buses supply most of the DC-poweredequipment.

The inverters are powered directly from thegenerator buses and are controlled by the IN-VERTER selector switch (Figure 1-2).

External PowerAn external power socket is located on theunderside of the right wing, outboard of theengine nacelle (Figure 1-3). The airplane willaccept DC power from a ground power unit(GPU) provided the polarity is correct, and theGPU voltage is below 32 volts. The BAT

switch must be positioned to ON in airplanesBB-364 and subsequent. Prior to BB-364, theGPU can energize the airplane without thebattery switch on and there is no overvoltageprotection (i.e., more than 32 volts).



Figure 1-2. Electrical Panel



LIGHTINGInteriorAn overhead light control panel (Figure 1-4)controls all the cockpit and instrument lights.

Cabin lighting is controlled by an interior lightswitch on the copilot’s subpanel, labeledBRIGHT–DIM–OFF. (Prior to BB-1444, except

1439, it is labeled START/BRIGHT–DIM–OFF)(Figure 1-5). This switch controls the cabin over-head fluorescent lights. Also, individual readinglights at each passenger station can be turned onor off by individual switches adjacent to the lights.

The CABIN SIGN switch is adjacent to the in-terior light switch.

A baggage area light switch is located just in-side the airstair door.

A single switch located just forward of theairstair door at floor level, controls the thresh-



MAX GEAR EXTENSION 181 KNOTS

AIRSPEEDS (IAS)

OPERATION LIMITATIONSTHIS AIRPLANE MUST BE OPERATED AS A NORMAL CATEGORY AIRPLANE IN COMPLIANCE WITH

THE OPERATING LIMITATIONS STATED IN THE FORM OF PLACARDS, MARKINGS AND MANUALSNO ACROBATIC MANEUVERS INCLUDING SPINS ARE APPROVED

THIS AIRPLANE APPROVED FOR VFR, IFR, & DAY & NIGHT OPERATION AND IN ICING CONDITIONS

CAUTION

STALL WARNING IS INOPERATIVE WHEN MASTER SWITCH IS OFFSTANDBY COMPASS IS ERRATIC WHEN WINDSHIELD ANTI-ICE AND/OR AIR CONDITIONING IS ON

DO NOT OPERATE

ON DRY GLASS

WINDSHIELD WIPERS

OFF

PARK SLOW

FAST

OFF

MASTER

PANEL

LIGHTS

ON

OVERHEAD

FLOODLIGHTS

OFFBRT

INSTRUMENT

INDIRECTLIGHTS

OFFBRT

AVIONICS

PANEL

LIGHTS

OFFBRT

ENGINE

INSTRUMENT

LIGHTS

OFFBRT

PILOT

FLIGHT

LIGHTS

OFFBRT

OVERHEAD

SUB PANEL

& CONSOLE

LIGHTS

OFFBRT

SIDE

PANEL

LIGHTS

OFFBRT

COPILOT GYRO

INSTRUMENT

LIGHTS

OFFBRT

COPILOT

FLIGHT

LIGHTS

OFFBRT

40 608020

% LOAD40 60

8020

% LOAD400 410390

FREQ0

8020 0

Figure 1-5. Cabin Lights Control Switch(BB-1439, 1444 and After)



old light, an aisle light, understep lighting, and

the exterior entry light. These three lights turnoff automatically when the airstair door isclosed and the handle is in the LOCK position.

The control switches for exterior lights arelocated on the pilot’s right subpanel, as seenin Figure 1-6.

MASTER WARNING SYSTEM

GeneralThe flight crew receives automatic indicationof system operation through the annunciator system. There are two annunciator panels lo-cated on the instrument panel. There are alsotwo master warning and two master cautionflashers.

Annunciator SystemThe warning annunciator panel is located in thecenter glareshield. It contains red indicators,

each of which represents a fault requiring thepilot’s immediate attention and action. At thesame time, red MASTER WARNING flasherson the glareshield directly in front of eachpilot begin flashing. The MASTER WARNINGflashers can be extinguished by depressing ei-h f h li h Th d li h h

A caution/advisory annunciator panel is lo-

cated on the center subpanel (amber indicatorsfor cautions and green for advisory). An amber caution illumination requires the pilot’s im-mediate attention to a fault but does not requireimmediate reaction. There are also two amber MASTER CAUTION f l asher s on theglareshield, just inboard of the red MASTERWARNING flashers. These operate the sameway as the MASTER WARNING flasher.

Two additional caution lights are on the fuelpanel which do not illuminate the MASTERCAUTION flasher.

The green advisory lights indicate functionalconditions, not faults; no master advisoryflashers are associated with the advisory lights.

FUEL SYSTEM

GeneralThe airplane fuel system consists of two sep-arate tank systems, one for each engine, con-nected by a common crossfeed line. Each of the tank systems is further divided into a main

and an auxiliary system.

Each main system consists of a nacelle tank,two wing leading-edge tanks, two box sec-tion bladder tanks, and an integral wing tank,all of which gravity feed into the nacelle tanks.The filler for this family of tanks is located ontop of the wing, near the wingtip.

The auxiliary fuel system consists of an aux-iliary tank, located in the wing inboard of theengine nacelle. It is filled separately throughan overwing filler, and employs an automaticfuel transfer system to supply the fuel to themain system.



Figure 1-6. Exterior Lights ControlSwitches



Each engine drives a high-pressure fuel pump

and a low-pressure boost pump. In addition,an electrically-driven low-pressure standbyboost pump is in the bottom of each nacelletank. The standby boost pump serves threefunctions:

1. To serve as backup for the engine-drivenfuel boost pump.

2. To pump aviation gasoline when flyingabove 20,000 feet.

3. To p u mp f u e l d u r in g c ro s s f e edo p e r a t i o n .

If the electric standby boost pump fails, cross-feed will not be possible from that side.

If aviation gasoline is used, a limitation of

150 hours of operation per engine before over-hauls must be observed.

There are two firewall shutoff valves, eachcontrolled by a red switch guarded to theOPEN position on the fuel control panel(Figure 1-7).

The fuel quantity is measured by a capaci-

tance system, which reads out in pounds on theleft and right fuel gages (Figure 1-7). A switch

between the gages allows the pilot to monitor

MAIN or AUXILIARY fuel levels.

POWERPLANTS

GeneralThe Super King Air is powered by two Prattand Whitney turbopropeller PT6A engines,

each rated at 850 SHP. They each have a three-stage, axial-flow, single-stage centrifugal flowcompressor (rpm indicated as N1) which isdriven by a single-stage reaction turbine. Thepower turbine is a two-stage reaction turbinecounter rotating with the compressor turbine.A pneumatic fuel control schedules fuel flow.Propeller speed remains constant within thegoverning range for any given propeller con-

trol lever position.

An accessory gearbox, mounted at the rear of the engine, drives the fuel pumps, fuel control,oil pump, refrigerant compressor (right en-gine), starter-generator, and the N1 tachome-ter transmitter.

Engine instruments are grouped at the left

center of the instrument panel.





Engine ControlsThere are three sets of controls on the pedestal(Figure 1-8):

1. Power levers provide control of enginepower from FULL REVERSE throughTAKEOFF power. Increasing N1 rpmresults in increased engine power.

2. Propeller levers operate springs to repo-sition the primary governor pilot valve,effecting an increase or decrease in pro-peller rpm.

3. Condition levers have three positions:

• FUEL CUTOFF

• LOW IDLE

• HIGH IDLE

Ground Fine (Beta)/ReversingWhen the power levers are lifted aft over theIDLE detent, they control the blade angle of thepropellers in Ground Fine (Beta) mode. This pro-vides a near zero thrust setting. For BB-1439,1444 and subsequent, to select reverse the power

levers need to be lifted over a second gate. Prior

to BB-1444 except 1439, reverse can be se-lected by continuing to move the power leversaft of the beta position into a red- and white-la-beled zone on the power quadrant.

Propeller reversing on unimprovedsurfaces should be accomplished

carefully to prevent propeller ero-sion from reversed airflow and industy or snowy conditions to preventobscured vision.

Condition levers, when set to HIGH IDLE,keep the engine operating at a minimum of 70% N1 for quicker reversing response due to

less spool up time.

Power levers should not be movedover either gate when the engines arenot running, or with engines runningand the propeller feathered, becausethe reversing system will be damaged.

CAUTION

CAUTION





FIRE PROTECTION

There are two fire-detection systems. On BB-1439, 1444 and subsequent the system consistsof a temperature sensing cable for each engine.Prior to BB-1444, except 1439, the systemuses three detectors incorporated into eachengine nacelle. Each system has red warningannunciator readouts and a test function. Theoptional engine fire-extinguisher system adds

an extinguisher cylinder within each engine na-celle. When the system is installed, glareshieldcontrol switches and additional positions onthe test switch are added (one for each extin-guisher cartridge). There are two portable fireextinguishers installed: one under the copilot’sseat, and the other near the entrance door.

BLEED-AIR SYSTEMGeneralEach engine compressor supplies bleed air for the pressurization and pneumatic systems.The bleed air used for pressurization is routedfrom the engine to a flow control unit then intothe pressure vessel. This same air is condi-tioned for environmental use.

The bleed air used for the pneumatic systemis tapped off prior to the flow control unit andis routed through a shutoff valve to a regula-tor. This pneumatic air is then used for surfacedeice, rudder boost, door seal, bleed-air warn-ing system, the flight hour meter, brake deice(if installed) and the landing gear hydraulic

reservoir ( if installed). Through the use of aventuri, vacuum suction is developed for flightinstruments, pressurization controller opera-tion and deice boots. One engine can supplysufficient bleed air for all associated systems.

Bleed Air Warning

BLEED AIR FAIL light on the warning an-

nunciator panel to illuminate. When bleed-air failure is indicated, the appropriate BLEEDAIR VALVE switch, on the copilot’s subpanelshould be placed to the INSTRument andENVIRonment OFF position (Figure 1-9).

ICE AND RAIN PROTECTION

Ice ProtectionIce protection is accomplished either pneu-matically or electrically. Pneumatic ice pro-tection uses engine bleed air for surfacedeicing of wing and horizontal stabilizer lead-ing edges , and ho t b rakes , i f i n s t a l l ed .Electrical heating elements are used for wind-

shield heating, fuel vent heat, propeller deic-ing, pitot mast heat, and stall warning vane heat(Figure 1-10).

The engine uses two types of anti-ice protection.To protect the air inlet, some of the hot engineexhaust gases are scooped up and directed intothe air inlet lip. To protect the engine, ice vanes



Figure 1-9. Bleed-Air Valve Control



are used which are moved into the airstream.

These cause a slight deflection in the enteringairflow, introducing a turn in the airstream. Theaccelerated moisture particles continue on to thedischarge port, rather than entering the engine.On BB-1439, 1444 and subsequent a secondelectric actuator is employed as a backup. Prior to BB-1444 except 1439, if the electric ice vanecontrols do not work, mechanical extensionhandles may be used. Operation of the vanes are

displayed either by green L or R ENG ANTI-ICE advisory lights (normal operation) or byamber L or R ENG ICE FAIL caution lights, in-dicating a possible malfunction. (Prior to BB-1444 except 1439, these annunciators are labeledL or R ICE VANE EXT and L or R ICE VANE,respectively.)

An optional brake deice system allows a flow

of hot bleed air to the brakes. If installed, op-eration is controlled by a switch on the ICEpanel (Figure 1-10) and indicated by a greenBRAKE DEICE ON advisory annunciator light.

Rain ProtectionThere are dual, two-speed, electric windshieldwipers, controlled by a switch on the overheadlight control panel. The PARK position on thecontrol switch sets the wipers to the inboardposition (Figure 1-11).

AIR CONDITIONING ANDHEATING

GeneralCabin air conditioning is provided by a re-frigerant-gas-vapor cycle refrigeration sys-tem. The compressor is mounted on the rightengine accessory pad. The refrigerant is routedto the airplane nose where the condenser coil,receiver-dryer, expansion and bypass valves,and evaporator are located.

The compressor is deenergized any time theengine speed is below 62% N1. An attempt touse air conditioning when N1 is below theabove values, will result in illumination of the green AIR COND N1 LOW advisory lighton the annunciator panel. High or low refrig-erant pressure switches will also trip the sys-tem and illuminate the reset switch light in thenose gear wheel well. (Prior to BB 729, itopens a fuse or a circuit breaker in the rightwing area next to the hot battery bus).

The forward vent blower sends recirculatedcabin air through the evaporator for air-con-ditioning output. The output from the ceilingoutlets will always be cool. Cool a ir also en-ters the floor-level duct, but is mixed withwarm environmental bleed air if either BLEEDAIR valve is open. Therefore, the lower duct,discharging pressurized air, will always bewarmer than the overhead “eyeball” ducts.

An optional aft evaporator and blower may beinstalled. Refrigerant will flow through bothevaporators as long as the system is operating,but additional cooling for the aft outlets willoccur only when the aft blower is operating.



DO NOT OPERATE

ON DRY GLASS

WINDSHIELD WIPERS

OFF

PARK SLOW

FAST



The cabin is heated by engine compressed

bleed air. After the airplane is airborne, am-bient air valves open and allow ambient ai r tomix with the bleed air for increased density.Pilot and copilot volume of air is controllableby respective air knobs on each subpanel. ACABIN AIR knob varies the volume of air di-rected into the cockpit or into the cabin flowducting. A DEFROST AIR control knob di-rects warm air to the windshield.

Unpressurized VentilationVentilation is provided through the bleed-air system during either pressurized or unpres-surized flight. Fresh air can also be providedby ram air but only during unpressurized flight.

Electric Heating (BB-1439, 1444and Subsequent)An optional electric heating system is avail-able for ground operation only. A ground power unit must be used prior to engine starting or generator power after engine starting in order to use electric heating system. It is for groundoperation only and is used in conjunction with

either manual heat or automatic temperaturecontrol mode. A green advisory light on the an-nunciator panel is provided to indicate power is being supplied to the unit. Both the ventblower and aft blower must be operating whenusing the electric heater.

Radiant Heating (Prior to BB-1444, Except 1439)An optional radiant heating system is an over-head heated panel system, which can be pow-ered by a ground power unit for cabin heatingprior to engine start, or it can use airplanepower to supplement the heating system inflight. It should be used only in conjunctionwith the manual temperature control mode.

PRESSURIZATION

GeneralThe pressurization system is designed to pro-vide a normal working pressure differential(psid) when flying at altitude. Table 1-1 presentsthe pressure differentials on the 200 and B200.

Bleed air from the engine compressor sectionis used to supply airplane pressurization.Engine bleed is mixed with ambient air toform a suitable mixture. The flow control unitand BLEED AIR VALVE switches, as seen inFigure 1-9, control the mixture. If this switchis positioned to ENVIRonmental OFF or

INSTrument and ENVIRonmental OFF, thebleed-air valve will be closed. When posi-tioned to OPEN, air is routed through a heatexchanger and then into a mixing plenum. Itmixes with recirculated air, is routed to the out-let ducts, and is introduced into the cabin.



FLIGHT ALTITUDE CABIN ALTITUDE

ALTITUDES ARE IN FEET

200 (6.0 ± 0.1 psid) B200 (6.5 ± 0.1 psid)

Table 1-1. CABIN ALTITUDES



The outflow valve, located on the aft pres-

sure bulkhead, controls the amount of pres-surized air in the airplane. The pressure andrate of cabin pressure changes are controlledby vacuum-operated modulation of the outflowvalve.

Also, a vacuum-operated safety valve ismounted adjacent to the outflow valve. Itserves four purposes:

1. To provide positive pressure relief if the outflow valve malfunctions.

2. To allow depressurizat ion when thepressure switch is moved to the DUMPposition.

3. To maintain an unpressurized state whileon the ground with the left landing gear

safety switch compressed.

4. To prevent negative differential.

When the BLEED AIR switches are OPEN, air used for pressurization enters the airplane,with or without ambient air, depending on theposition of the landing gear safety switch (onthe ground, no ambient flow), and temperature.For pneumatic flow packs (prior to BB-1180),

the use of ambient air is also dependent on am-bient pressure.

An adjustable cabin pressurization controller is located on the pedestal (Figure 1-12).

The CABIN ALT selector knob can be used to

select a desired cabin pressure altitude be-tween -1,000 feet and 15,000 feet. The se-lected pressure altitude will be reflected on theouter scale of the indicator. The inner scaleshows the highest ambient pressure altitudethat the airplane can fly in order to maintainthe selected CABIN ALT. A rate control se-lector knob, placarded RATE–MIN–MAX canselect between 200 and 2,000 feet per minute

of change of cabin altitude. These controlsdirect the action of the outflow valve.

The CABIN PRESS–DUMP–TEST switch islocated next to the cabin pressurization con-troller. When selected to DUMP, the safetyvalve opens, relieving all accumulated cabinpressure. In TEST, the valve is closed, by-passing the left landing gear safety switch for

a ground pressurization test.

LANDING GEAR AND BRAKES

GeneralThe retractable tricycle landing gear is ex-tended or retracted by a 28-volt motor andgearbox or by an electrically-driven hydraulicpump (airplane Serial Nos. BB-1193 and sub-sequent). The LDG GEAR CONTROL HAN-DLE on the pilot’s right subpanel controls thesystem. A solenoid-operated lock preventsthe handle from being raised when the air-plane is on the ground. This can be bypassedby the red DOWN LOCK REL button just tothe left of the control handle.

Individual gear position is indicated by threegreen lights adjacent to the handle. The gear han-dle contains two red lights, which illuminatewhen the gear is in transit or not properly locked.Two versions of the control panel are found inFigure 1-13. On airplanes with the hydrauli-





Manual Extension (HydraulicGear)Manual extension of the gear on these air-planes requires pulling the LANDING GEARRELAY circuit breaker and placing the land-ing gear switch handle in the DN position. Ahydraulic hand pump, located on the floor be-tween the pilot’s right foot and pedestal (Figure1-14), is then operated until three green gear

position indicator lights are observed.

Manual Extension (ElectricGear)The landing gear can be manually extended bypulling the LANDING GEAR RELAY circuitbreaker and placing the landing gear switch

handle in the DN position. Pulling up andturning the emergency engage handle (Figure1-14) positions an emergency drive gear tothe gearbox. A continuous-action ratchet isthen pumped to lower the gear. The system maybe reverted to electrical operation by reposi-



PRIOR TO BB-453 (SUBSEQUENT MODELSHAD THE GEAR DOWN INDICATORLIGHTS IN A CUBE ARRANGEMENT)

BB-1439, 1444 AND AFTER

Figure 1-13. Landing Gear Control Panel

Fli htS f t



tioning both handles on the floor and resetting

the circuit breaker.

Warning SystemDuring flight, a warning horn and red lightsin the landing gear handle warn the crew of im-proper landing gear position relative to flapand/or power lever position. They also acti-vate when the gear handle is up while on theground.

Nosewheel SteeringThe rudder pedals control nosewheel steeringwhile the gear is down. Both the nosewheelsteering and rudder deflection receive inputsfrom rudder pedal motion, but in varying pro-portions depending on the speed that thewheels are rolling. When the wheel brakesare applied during rudder pedal deflection,there is even greater steering effect. Duringnose gear retraction, it is mechanically self-centered and receives no further rudder pedalsteering force.

Brake SystemDual hydraulic brakes are operated by de-pressing either the pilot’s or copilot’s toe por-tion of the rudder pedals. Both sets of pedalsoperate the brakes. Prior to BB-666, the ini-tial pressure from a set of pedals will positiona shuttle valve in the braking system. Brakeoperation from the opposite side can then onlybe accomplished by moving the shuttle valve.

A parking brake (Figure 1-15) can be actuatedto lock the pressure within the brake lines.The airplane may be designed to permit park-ing brake operation either in conjunction withpilot brake pressure only, or with pressurefrom either set of brakes

cal stabilizer. Interconnected conventionalcontrol columns within the cockpit controlthe ailerons and elevators. Rudder pedals arealso connected so that either the pilot or copi-

lot can operate the rudder. There are dual flapson each wing. Rudder, elevator, and ailerontrim are adjustable with controls mounted onthe center pedestal. The flight control sur-faces are illustrated in Figure 1-16.

OperationThe flight controls are cable operated and re-

quire no power assistance. Flaps and optionalelectric elevator trim are electrically driven.A pneumatic rudder boost system assists in di-rectional control when one engine has failed.

Rudder, elevator, and aileron trims are adjustable with controls on the center pedestal.Elevator trim is manual or optionally electri-cal. There is a position indicator on each

pedestal tab control (Figure 1-17).

A lever on the control pedestal (Figure 1-18)controls the two flaps installed on each wing.A wing flap percentage indicator is locatedon the pedestal next to the cabin climb ratei di t



Figure 1-15. Parking Brake Handle

Fli htS f t





ELEVATORS

TRIM TABS

RUDDER

TRIM TAB

AILERON

TRIM TAB

FLAPS

FLAPS

GROUND ADJUSTABLE TAB

AILERON

Figure 1-16. Flight Control Surfaces

FlightSafety



Pitot SystemA heated pitot tube is located on each s ide of the lower portion of the nose. The pilot’s air-speed indicator uses input from the left pitotmast, while the copilot’s input is from theright mast (Figure 1-19).

Static SystemThe normal static system provides separateinput for pilot and copilot instruments. Eachhas a port on each side of the aft fuselage,which is not heated (Figure 1-20).

If the pilot’s static system is plugged, an al-ternate air tube obtains static air from insidethe unpressurized rear fuselage. This systemis selected by moving the PILOT’S STATICAIR SOURCE valve handle, located on theright side panel, to the ALTERNATE position(Figure 1-21).

The pilot’s airspeed, vertical speed,and altimeter indications change whenthe alternate static air source is in use.

Super King Air B200The masks and oxygen duration chart are basedon a flow rate of 3.9 liters per minute (LPM-NTPD) per mask. When using the diluter-de-mand crew mask in the 100% mode, each maskcounts as two masks at 3.9 LPM-NTPD.

Super King Air 200The masks and oxygen duration charts are basedupon 3.7 standard liters per minute (SLPM) per mask. The only exception is the diluter-demandcrew mask when used in the 100% mode. Whencomputing oxygen duration, each diluter-de-mand mask used in the 100% mode, is countedas two masks at 3.7 SLPM.

WARNING



Figure 1-19. Pitot Tubes

Figure 1-20. Static Ports

FlightSafety



Manual Plug-in SystemEarly Super King Air 200s employ a constant-flow, plug-in system. All masks for crew andpassengers are stored in the seat area and areremoved and plugged into available recepta-cles as needed.

Autodeployment System

When the autodeployment system is installedfor the passengers, the crew normally has di-luter-demand masks, which are one-hand,quick-donning masks.

Oxygen supply is controlled by a push-pullhandle, placarded PULL ON-SYStem READYand is located on the left side of the pedestal(Figure 1-22). (Prior to BB-1444, except 1439

they are overhead in the cockpit — Figure 1-22). When pushed in, no oxygen is availableanywhere in the airplane. It should be pulledout prior to engine start to ensure availableoxygen when needed. The primary oxygensystem delivers oxygen to the two crew masks,to the first-aid outlet in the toilet area, and tothe passenger oxygen system shutoff valve.

The passenger system is the constant-flow type.If the oxygen system line has been charged(oxygen in the supply bottle and SYStemREADY handle pulled) when the cabin altitudeexceeds approximately 12,500 feet, the oxy-gen pressure will automatically open the maskstorage doors and allow the passenger masks todrop out. Oxygen will flow to the mask when a

further pull on the lanyard by the passenger

pulls the pin out of the valve. A green PASSOXYGEN ON light on the advisory annuncia-tor panel will indicate that the passenger maskshave dropped out of the overhead.

If the oxygen supply line is charged, oxygenis available at the first-aid station. The cover must be opened and the valve turned on.

In the event that oxygen pressure fails to openthe passenger oxygen shutoff valve automat-ically, the pilot has a PASSENGER MAN-UAL OVERRIDE handle on the right side of the pedestal (prior to BB-1444, except 1439,it is next to the SYStem READY handle on theoverhead panel). It will open the valve man-ually, and all other operations will be the sameas in the automatic mode.

AIRPLANESTRUCTURES

GENERAL

The Super King Air is 43 feet 9 inches long

from the nose to the aft most point of the hor-izontal stabilizer (Figures 1-23 and 1-24). Theairplane sections consist of the:

• Fuselage

• Wings

• Empennage





FlightSafety





43' 10" (1)43' 9" (2)

29.60" (3)29.85" (4)

WING AREA303.0 SQUARE FEET

18' 5"

54' 6"

98.5" DIA (3)98" DIA (4)

CONFIGURATIONS:

(1) STANDARD LANDING GEAR

(2) HIGH FLOTATION LANDING GEAR

(3) HARTZELL PROPELLER

(4) MCCAULEY PROPELLER

14' 11.5" (1)14' 11.4" (2)

14.50"(1), (3)14.04"(2), (3)14.75"(1), (4)14.29"(2), (4)

14' 10" (1)14' 6" (2)

FlightSafety



The fuselage is composed of the:

• Nose section

• Cockpit

• Cabin

• Foyer and aft cabin

• Aft fuselage

The wing is built as a center section and twooutboard wing assemblies.

The empennage is composed of a vertical sta-bilizer with a high T-tail horizontal stabilizer.

FUSELAGE

The nose section is an unpressurized equipmentstorage area, separated from the cockpit areaby the forward pressure bulkhead (Figure 1-25).

The cockpit is separated from the cabin by asliding door for privacy and to prevent lightspilling between compartments. A typical in-strument panel is shown in Figure 1-26.

Various configurations of passenger chairsand couches may be installed. All passenger chairs are placarded FRONT FACING ONLYor FRONT OR AFT FACING. Only chairs somarked may be installed facing aft. All aft-fac-ing chairs and al l forward-facing chairs

equipped with shoulder harnesses have ad-

justable headrests.

Before takeoff and landing, the head-rest should be adjusted as required toprovide support for the head and neckwhen the passenger leans against the

seatback.

Couches, if installed, are not adjustable.

The cabin is separated from the foyer by an-other sliding door to provide privacy for thetoilet, which is located in the foyer. When thetoilet is not in use, seat cushions convert the

position to another passenger seat.

The aft cabin area may have one or two op-tional folding seats installed. When these seatsare not needed, they may be folded against thecabin sidewall, and the entire aft cabin areamay be utilized for baggage storage.

Webs should secure baggage andother objects in order to prevent shift-ing in turbulent air.

CAUTION

CAUTION



AFT

CABIN

FOYERCABINCOCKPIT

FlightSafety



Items stowed in this area are easily accessi-ble in flight. An optional curtain can be closedto separate the aft cabin from the foyer. Alatching compartment door may be installedin place of the curtain.

DOORS

Cabin DoorThe cabin door is located on the left side of the fuselage, in the foyer area. The cabin door is hinged at the bottom, and swings out andd h d (Fi 1 27) A h d li

BB-1444, except 1439) may be installed alongthe other side of the steps, giving support toboth sides of the door.



Figure 1-26. Cockpit Layout (Typical)

FlightSafety



Only one person at a time should beon the door stairway.

The plastic handrail is utilized when closingthe door from the inside. The door is closedagainst an inflatable rubber seal around theopening. When the weight of the airplane isoff the landing gear, pneumatic air is used toinflate the door seal through a 4-psi regulator.

The door-locking mechanism can be operatedby either the outside or inside door handle,which rotates simultaneously. A release but-ton (Figure 1-28) is adjacent to each handle andmust be held depressed before the handle canbe rotated. The handle system necessitates atwo-hand operation, thereby ensuring a de-liberate action. The release button also in-corporates a pressure-sensing diaphragm, sothat if there is a pressure differential betweenthe inside and outside, the pressure on the re-lease button must be proportionally increasedto prevent inadvertently opening the door while pressurized.

Never attempt to check or unlock the door inflight. If the CABIN DOOR light is on (amber in the 200, red in the B200), or if the pilot sus-pects door security, direct all occupants to re-main seated with seatbelts secured, descend asnecessary, and depressurize the airplane. After the airplane has landed and stopped, and the

cabin has been depressurized, a crewmember

can then check the door security.

When closing the door from inside the air-plane, pull up on the handrail until the airstair door reaches the door frame. Rotate the door handle up as far as possible, pulling inward onthe door. The door should seal; then rotate thehandle down to lock the door (Figure 1-28).Positive locking may be checked by attempt-

ing to rotate the handle without depressingthe release button. It should not move. A plac-ard is located beneath the folded step justbelow the door handle. The placard showshow to check the locks in the inspection portwindows near each corner of the door (Figure1-29). A green stripe painted on each of thefour latch bolts should be aligned with its re-spective black pointer (Figure 1-30).

CAUTION


g yinternational

Figure 1-29. Placard and Inspection Port

FlightSafetyi t ti l



Cargo Door (200C and B200C)A large, swing-up cargo door, hinged at the top,provides access for loading and unloadinglarge cargo. The airstair door is an integral partof the cargo door and should be closed andlatched when the cargo door is opened.

The cargo door latches can be operated onlyby the use of two handles, both located insidethe airplane. The handle in the upper part of the door controls the rotating latches in the for-ward and aft sides, while the handle in thelower, forward part of the door actuates four pin-lug latches along the bottom of the door.

Once the latches are retracted initial pres-

come manually, until the door is almost closed.

When the door is almost closed, the gas springovercenter mechanism will redirect springpressure toward the closed position, assist-ing the latching cycle.

The door closes against a rubber seal, to main-tain the pressure vessel integrity. The seal isnot inflated by pneumatic bleed air, but rather allows cabin-pressurized air to seep into holes

on the inside. This allows for greater sea lingwhen there is a high pressure differential.

Emergency ExitThe emergency exit window, placarded EXIT-PULL (Figure 1-31) is located at the forwardright side of the passenger compartment. Itcan be released from the inside by using a

pull-down handle, or from the exterior (if itis unlocked) by a flush-mounted, pull-outhandle (Figure 1-31). It is a plug-type exit,which is removed completely from the frameand taken into the cabin. The exit can belocked from the inside, but can be openedfrom the inside even when it is locked. For BB-415 and after, the locking mechanism isactivated by pulling out a handle below the

door release handle (Figure 1-31). Prior air-craft and BL-1 and after have a key next tothe door release handle that can lock/unlockthe door. This key cannot be removed whenthe door is locked.

This door must be unlocked prior to takeoff for exterior opening in case of emergency.

CABIN WINDOWS

Each cabin windowpane is composed of asheet of polyvinyl butyral between two trans-parent sheets of acrylic plastic. It is stressedto withstand the cabin pressure differential


g yinternational

Figure 1-30. Latch Bolt




Polarized TypeTwo dust panes are inboard of the cabin win-dow each composed of polarized film. Theinboard pane may be rotated to permit lightregulation.

Do not look directly at the sun , eventhrough polarized windows, becauseeye damage could result.

When the airplane is to be parked inareas exposed to intensive sunlight,the polarized windows should be ro-

Shade TypeA single sheet of tinted acrylic plastic servesas a dust pane. The shade is mounted in the win-dow frame, inboard of the cabin window dustpane. It can be moved along detents in a track.

CONTROL LOCKS

The flight and engine controls are mechani-cally locked by a U-shaped clamp and two pinswithin the cockpit, as seen in Figure 1-32. Thepins lock the primary flight controls and the U-

shaped clamp fits around the engine controllevers. A pin is inserted through the controlcolumn to lock the ailerons and elevator. A sec-ond pin is inserted through a hole in the floor,which locks the rudder bellcrank. All locksmust be installed and removed together to pre-l d ii fl i i h h i l

CAUTION

WARNING


g yinternational

Figure 1-31. Emergency Exit Release Handles




Before starting engines, remove thelocks.

Remove the control locks before tow-

ing the airplane. If towed with a tugwhile the rudder lock is installed,serious damage to the steering link-age can result.

CAUTION

WARNING


international

Figure 1-32. Control Locks




CHAPTER 2

ELECTRICAL POWER SYSTEMS

CONTENTS

Page

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

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

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

Battery.............................................................................................................................. 2-2

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

Ground Power .................................................................................................................. 2-5

Controls and Indicators .................................................................................................... 2-8

Distribution ...................................................................................................................... 2-8

Operation ....................................................................................................................... 2-10

Avionics Master Switch ................................................................................................. 2-12

AC Power Inverters ....................................................................................................... 2-12

Controls and Indicators.................................................................................................. 2-12

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

Operation ....................................................................................................................... 2-15

LIMITATIONS ...................................................................................................................... 2-22

Generator Limits (250 Amperes)................................................................................... 2-22

Starters ........................................................................................................................... 2-22





ILLUSTRATIONS

Figure Title Page

2-1 Electrical Component Location................................................................................ 2-2

2-2 Battery Cooling (Nickel Cadium) ............................................................................ 2-3

2-3 Battery Control Circuit............................................................................................. 2-3

2-4 Volt-Loadmeters-Battery Ammeter.......................................................................... 2-4

2-5 BATTERY CHG Annunciator.................................................................................. 2-4

2-6 Generator.................................................................................................................. 2-4

2-7 Generator Switches................................................................................................... 2-5

2-8 Generator Control Circuit......................................................................................... 2-6

2-9 Ground Power Connector......................................................................................... 2-7

2-10 External Power Circuit ............................................................................................. 2-7

2-11 MASTER SWITCHES............................................................................................. 2-8

2-12 Lights and Meters..................................................................................................... 2-8

2-13 Electrical Distribution .............................................................................................. 2-92-14 Circuit-Breaker Panels—Pilot’s ............................................................................. 2-10

2-15 Circuit-Breaker Panels—Copilot’s......................................................................... 2-11

2-16 Avionic Power Distribution.................................................................................... 2-13

2-17 Typical Avionics Bus Distribution (EFIS Equipped Aircraft) ............................... 2-14

2-18 Inverters.................................................................................................................. 2-15

2-19 Volt-Frequency Meter ............................................................................................ 2-15

2-20 Inverters Control Circuit ........................................................................................ 2-16





2-24 Electrical System—Super King Air B200 (BB-734, 793, 829, 854-870,

874-891, 894, 896-911, 913-1438, 1440-1443, BL 37-138).................................. 2-202-25 Electrical System—Super King Air 200 (B-2, 6-733, 735-792, 794-828, 830-853

871-873, 892, 893, 895, 912, BL-1-36) ................................................................. 2-21

TABLESTable Title Page

2-1 Limitations—Ground Operations........................................................................... 2-22

2-2 Fuel Control Circuit-Breaker Panel ....................................................................... 2-23

2-3 Right Side Circuit-Breaker Panel........................................................................... 2-24

2-4 Pilot’s Right Subpanel Circuit-Breaker Switches .................................................. 2-28





INTRODUCTION

The primary electrical system on the airplane is a 28-VDC generator system. It is usedfor inverter input and, through the distribution system, for powering the electronicequipment and landing gear. The DC system consists of generation, distribution, stor-age, control, and monitoring of DC power. The AC system consists of the inverters, power

distribution, control, and monitor ing of AC power.

A section on specific limitations, a circui t-breaker table, and a series of questions con-clude this chapter.

# 1 S E R

V O

S Y S T E M

B A T T

H O T

B A T O F F

A C

G E N

# 1 D C

G E N

# 1 E N G

O I L

P L

CHAPTER 2ELECTRICAL POWER SYSTEMS





DC POWER

BATTERYFor BB-1632 and subsequent, a single, 24-volt,42 ampere-hour sealed lead acid gel cell batteryis located in the right wing center section for-ward of the main spar. Prior to BB-1632, a sin-

l 24 l 34/36 h i k l d i

generators are not on. Power to the main busfrom the battery is routed via the battery relay,which is controlled by the BAT ON–OFF

switch on the pilot’s left subpanel.

For aircraft BB-1632 and subsequent, the bat-tery ammeter (Figure 2-4) provides a directreading of the charge or discharge rate of thebattery (–60 amps to +60 amps). The charge

h ld b 0 10 f k ff


STARTER–GENERATOR

INVERTER

INVERTER

BATTERY

EXTERNALPOWER

CONNECTORSTARTER–

GENERATOR

PRINTEDCIRCUIT BOARDS

Figure 2-1. Electrical Component Location





Figure 2-2. Battery Cooling (Nickel Cadium)

I

SOLATION

BUS

M

AIN

BATTERY

B

US

BATTERYRELAY

BATTERYSWITCH

H

OT

BATTERY

BU

S

BATTERY

S

H

U

N

T

TOBATTERYCHARGESENSOR




Following a battery-powered engine start, thebattery recharge current is very high and causesillumination of the BATTERY CHG annunci-ator, thus providing an automatic self test of the detector and the battery. As the batteryapproaches a full charge and the charge cur-rent decreases to a satisfactory level, the an-nunciator will extinguish. This will normallyoccur within a few minutes after an enginestart, but it may require a longer time if the bat-tery has a low state of charge initially beforeengine start, or if it is exposed to low or hightemperatures. In flight this alerts the pilot thatconditions may exist that could eventuallydamage the battery. If the BATTERY CHGannunciator illuminates, the pilot should turnthe battery switch to OFF. If the annunciator remains on after the BAT switch is moved tothe OFF position during the check, a mal-function is indicated in either the battery sys-tem or charge current detector, in which casethe airplane should be landed as soon as prac-ticable. This system is designed for continu-ous monitoring of the battery condition.

GENERATORS

Two 30-volt, regulated to 28.25 ± .25 volts,250-ampere starter-generators connected in par-

allel provide normal DC power (Figure 2-6).Either one of the generators can supply the en-tire electrical load.

NOTE

Optional 300-ampere starter-gener-ators are available and installed onsome airplanes.

Starter power to each starter-generator is pro-vided from the main battery bus through astarter relay. The start cycle is controlled bya three-position switch for each engine la-beled IGNITION AND ENGINE START.

When placed to the ON (up) position, theswitch becomes mechanically locked and mustbe pulled out to reposition. When held to thedown pos ition, labeled STARTER ONLY, theassociated engine will motor, but ignition willnot occur. When released, the spring-loadedswitch will move to the center position, whichis labeled OFF.


40 608020

0 3010 20

0

PUSH

FORVOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 608020

0 3010 20

0

PUSH

FORVOLTS

100

DC VOLTS

% LOAD

Beechcraft

400 410420380

390

100 130110 120

PUSH

FORVOLTS

AC VOLTS

FREQ

Beechcraft

-60

0

+60BATT AMPS

Figure 2-4. Volt-Loadmeters-Battery Ammeter





During an engine start, the starter-generator,

drives the compressor section of the enginethrough the accessory gearing. The starter-generator, in the start mode, could initiallydraw approximately 1,100 amperes, and thendrop rapidly to about 300 amperes as the en-gine reaches 20% N1. When the engine reachesapproximately 35%, it drives the starter. After the condition lever is set to high idle (ap-proximately 70%), the generator can be turned

on.

The generator operation is controlled by indi-vidual generator switches located on the pilot’sleft subpanel under the MASTER SWITCHgang bar with the BAT switch. As shown inFigure 2-7, the switches are labeled GEN 1and GEN 2. In order to turn the generator on,the control switch must be held upward in the

GEN RESET position (Figure 2-7) for a min-imum of one second, then released to the ONposit ion. (Prior to BB-88, the generator switches do not have the reset position.)

Figure 2-8 shows that power to the bus systemfrom the generators is protected by Generator

Control Units (GCU). For BB-88 and after, theGCU operates a line contactor relay to protectthe generator. Prior to BB-88, reverse-currentprotection is provided by a unit in line with thegenerator output.

The generators are controlled by individual

loadmeter (Figure 2-8) on the overhead panel

which reads in percent of the generator’smaximum continuous capacity. Normally,this value is 250 amps; therefore, a loadme-ter reading of .5, or 50%, is equal to 125amps of generator output.

NOTE

The generators will drop off the line

if underexcitation, overexcitation,overvoltage, or undervoltage condi-tions exist.

GROUND POWER

For ground operation, a ground power recep-tacle, located under the right wing outboardof the nacelle, is provided for connecting aground power unit (Figure 2-9). A relay inthe external power circuit will close only if:

1. The ground power source polarity iscorrect.

2. The BAT SWITCH is on.

3. The GPU voltage is not greater than 32volts (BB-364 and subsequent).

NOTE

Prior to BB-364, the battery switchdoes not have to be on to applyground power (Figure 2-10).

For starting, an external power source capa-

ble of supplying up to 1,000 amperes (300amperes maximum continuous) should beused. A caution light on the caution advisoryannunciator panel labeled EXT PWR is pro-vided to alert the operator when a groundpower plug is connected to the airplane. Some

li i l d i h

Figure 2-7. Generator Switches






I S O L A T I O N

B U S

M A I N

B A T T E R Y B U S

VOLTLOAD

METER

SHUNT

ISOLATION LIMITER

ISOLATIONLIMITER

RIGHTSTARTRELAY

BATTERYSWITCH

L GEN LINECONTACTOR

HOTBATTERY

BUS

LEFT GEN CONTROL

LEFTSTARTER

GEN

RIGHT GEN BUS

VOLTLOAD

METER

SHUNT

RIGHT GEN CONTROL

RIGHTSTARTER

GEN

R GEN LINECONTACTOR

OFF

BATTERY RELAY

BATTERYRELAY

SHUNT

BATTERYCHARGEMONITOR

BATTERY

LEFTSTARTRELAY

I S O L A T I O N

B U S

M A I N

B A T T E R Y B U S

BATTERY

LEFT GEN CONTROL

VOLTLOAD

METER

SHUNT

ISOLATION LIMITER

LEFTSTARTER

GEN

VOLTLOAD

METER

SHUNT

ISOLATION LIMITER

RIGHTSTARTRELAY


BATTERYSWITCH

LEFTSTARTRELAY

HOTBATTERY

BUS

BATTERYRELAY

S H U N T

LEFT GEN BUS

REVERSECURRENT

PROTECTION

RIGHT GEN CONTROL

RIGHTSTARTER

GEN

REVERSECURRENT

PROTECTION

RIGHT GEN BUS

LEFT GEN BUS

BATTERY RELAY





Never connect an external power source to the airplane unless a batteryindicating a charge of at least 20 voltsis in the airplane. If the battery volt-age is less than 20 volts, the battery

must be recharged, or replaced with

a battery indicating at least 20 volts,before connecting ground power.

Observe the following precautions when usinga ground power source:

1. Use only a ground power source that isnegatively grounded. If polarity of thepower source is unknown, determine

the polarity with a voltmeter before con-necting the unit to the airplane.

2. Before connecting a ground power unit,turn off the avionics master power switchand the generator switches, and turn thebattery switch on.

Voltage is required to energize theavionics master power relays to re-move the power from the avionicsequipment. Therefore, never applyground power to the airplane without

CAUTIONCAUTION


Figure 2-9. Ground Power Connector

ISOLATIONLIMITER

EXT POWERCONNECTOR

HOT BATTERYBUS

BATTERYRELAY

EXTERNALPOWER M

A I N

B A T T E R Y B U S

I S O L A T I O N

B U S

EXT POWERRELAY

BATTERY

SHUNT


OFF

ON

BATTERYRELAY

ISOLATIONLIMITER

EXT POWERCONNECTOR

HOT BATTERYBUS

EXT POWERRELAY

M A I N

B A T T E R Y B U S

I S O L A T I O N

B U S

BATTERYRELAY

BATTERYSWITCH

BATTERYRELAY

BATTERY

EXTERNAL POWERPLUG ENGAGED

SENSOR

S H U N T

EXT POWERSENSEISOLATIONLIMITER

EXT POWERCONNECTOR

HOT BATTERYBUS

EXT POWERRELAY

M A I N

B A T T E R Y B U S

I S O L A T I O N

B U S

BATTERYRELAY

BATTERYSWITCH

BATTERYRELAY

BATTERY

EXTERNAL POWERPLUG ENGAGED

SENSOR

S H U N T





first applying battery voltage. If the

battery is removed from the airplaneor if the battery switch is to be placedin the OFF position, turn each indi-vidua l radio and other avionicsequipment off.

3. After the external power plug is con-nected and power is applied, leave thebattery on during the entire ground

power operation to protect transistor-ized equipment against transient volt-age spikes.

The battery may be damaged if ex-posed to voltages higher than 30 voltsfor more than two minutes.

Only use a ground power source fitted with anAN2552-type plug. If uncertain of the polar-ity, check it with a voltmeter to ensure that itis a negative-ground plug. Connect the posi-tive lead to the larger center post of the re-ceptacle, and connect the negative-groundlead to the remaining large post. The small post

is the polarizing pin; it must have a positivevoltage applied to it in order for the externalpower relay to close.

CONTROLS AND INDICATORS

Electrical control switches are convenientlylocated on the pilot’s left subpanel (Figure 2-

11). The battery switch and the two generator switches are positioned under a hinged flap la-beled MASTER SWITCH, commonly referredto as the gang bar. When this flap is depressed,the battery and both generators are switched off.

El i l i di i i h h

For NiCad batteries, in the event of an exces-sive battery charge rate, the amber BATTERYCHG light comes on.

The generator loadmeters indicate generator amperage in percent of 250 amps per genera-tor and the associated meter button must bepressed to indicate bus voltage.

DISTRIBUTIONThe battery is connected to a hot battery bus(Figure 2-13) which powers threshold lights,the fire extinguishing system, firewall shut-off valves, the battery relay, ground com-

i i ili DC b (if i ll d)

CAUTION


Figure 2-11. MASTER SWITCHES

L DC GEN R DC GEN

40 608020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 608020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

Figure 2-12. Lights and Meters





isolation limiters (current limiters), connects

the left and right generator buses together.

When the battery, generators, or GPU are pro-viding power, the isolation bus, L generator bus, and R generator bus function as one unit,as long as both current limiters are not open.There are four subbuses fed by both the left andright generator buses. They are labeled No. 1through No. 4 DUAL FED BUS. Each subbus

is fed from either side through a 60-ampere cur-

rent limiter, a 70-ampere reverse current diode,

and a 50-ampere circuit breaker which is ac-cessible to the crew. There are eight of these 50-amp feeder breakers. Four are located on thecopilot’s side panel for the No. 1 and No. 2 sub-buses, and on the fuel panel circuit breaker busfor the No. 3 and No. 4 subbuses. Of those itemswith paired circuits such as the left and rightlanding lights, the distribution will be such thatthe left circuit is on the No. 1 or No. 3 dual fed

bus and the right is on the No. 2 or No. 4.


DUAL FED SUB-BUS #1

DUAL FED SUB-BUS #2

STARTRELAY

G C U

VOLT / LOADMETER

R/HGEN LINECONTACTOR

R/H STARTER/ GENERATOR

325A325AL / H

GE

VOLT / LOADMETER

G C U

L/HGEN LINE

CONTACTOR

L/H STARTER/ GENERATOR

STARTRELAY

ISOLATION BUS

MAIN BATT BUSAVIONICS

#1

R / H

GE

AVIONICS

#2

HOT BUS

SHUNT

BATTRELAY

BATTERY

OFF

BATTSWITCH

ON











AVIONICS MASTERPOWER CB

AVIONICS MASTERPOWER SWITCH

5A

NUMBER 1DUAL FED BUS

ON

OFF

RIGHTGENERATOR

BUS

40A

30A

NUMBER 2NUMBER 1

30A

40A

LEFTGENERATOR

BUS





AVIONICSMASTER

RIGHT MULTI FCTN PRCSR

COMM NO 1

NAVNO 1

COMPASS NO 1

ADF NO 1

RMI NO 2

XPNDR NO 1

DME NO 1

EFIS FANS NORMAL

DTU

STEREO

FMS

RADIO ALTM

AP SERVO

FCS POWERPILOT TURN & SLIP

PILOT ALTM & AIR DATA

PITCH TRIM

OUTSIDE AIR TEMP

CVR

AURAL WARN

PILOT AUDIO

ALT ALERT

LEFT MULTI FCTN PRCSR

PILOT EADI

DSPL PRCSR

COPLT TURN & SLIP

COPLT ENCD ALTM

CABIN AUDIO

COPILOT AUDIO

PILOT EHSI

ELEK DSP

COMM NO 2

NAVNO 2

COMPASS NO 2

EFIS AUX BAT

ADF NO 2

RADAR

MULTI FCTN DSPL

RMI NO 1

XPNDR NO 2

DME NO 2

EFIS FANS STBY

OFF

ON

AVIONICSMASTERSWITCH

LIMITER

40A

30A

AVIONICSNO 1

AVIONICSBUS NO 1RELAY

AVIONICSBUS NO 1

AVIONICSBUS NO 2

ISOLATION BUS& BATTERY BUS*

R/HGENERATORBUS*

NO 1 DUAL FEDELECTRICALBUS*

NO 2 DUAL FEDELECTRICALBUS*

L/HGENERATORBUS*

AVIONICS ANN





Inverter operation is controlled by an IN-VERTER select switch (Figure 2-20) on thepilot’s left subpanel. Selection of either in-verter activates the inverter power relay andsupplies inverter input power. Only one in-

t t t ti

DISTRIBUTION

The inverter system described here is the s tan-dard installation. The circuit diagram in ATAchapter format 24-20 of the Wiring Diagram Manual provides a circuit routing of the DCand AC power for the standard airplane in-strumentation. Due to the wide variety of cus-tomer-requested avionics options installed inthe airplane, the avionics diagrams are sup-plied with each airplane to provide the avion-ics portion of the AC power system. Thesewiring diagrams will show any modifications,which have been made to the standard instal-lation (Figures 2-21 through 2-25).

OPERATION

Turn the INVERTER select switch to either in-verter position, note that the INVERTER

Figure 2-18. Inverters

Figure 2-19. Volt-Frequency Meter

2 - 1 6

INVERTER NO 1



F OR

T RAI NI N G P URP O S

E S

ONL Y

S UP E

R KI N G

A I R

2 0 0 / B 2

0 0 P I L O T T R A I NI N G M

A N U A L

F

l i gh t S af e t y

i n t er n a t i on al

No.1 INVERTER ON LINE

AC VOLT/FREQ

METER

POWERRELAY

LIMITER

LIMITERPOWER

RELAY

DC GROUND

DC GROUND

DC POWER

DC POWER

115 VAC

115 VAC

26 VAC

26 VAC

AC COMMON

AC COMMON

5A

NO 1

OFF

NO 2INVERTERSELECTSWITCH

5A

115 VACSELECT

RELAY

26 VACSELECT

RELAYL/H GEN

BUS

R/H GEN

BUS

115 VACBUS

AC COMBUS

26 VACBUS

5A

5A

AC TESTJACK(BLUE)

AC POWERRETURNS

FROM

SYSTEMS

1A

1A

1A

2A

1A

2A

1A

1A

1A

1A

50A

50A

1A

1A

B200 AC Power System

5A

ANN IND

AVIONICS JUNCTION BOX

INVERTER No 2

No. 2 INVERTERCONTROL

No. 1INVERTERCONTROL

28VDC

5A

10A

10A

5A

INVERTERSELECTRELAY

INVERTERWARNING

RELAY

VG POWER & REF TO

OTHER SYSTEMS

RADAR REF FROM VG

FOR STABILIZATION

AP REF FROM VG

AP YAW RATE GYRO POWER

COMPASS 2 REF TO RMI

NO. 1 & MPU

ADF 1 REF SIGNAL FOR RMI

NO. 1 & NO. 2 & COPILOT EHSI

COMPASS 1 REF TO RMI NO. 2

ADF 2 REF SIGNAL FOR RMI


NAV 1 REF SIGNAL FOR RMI


NAV 2 REF SIGNAL FOR RMI


FMS FOR HDG & AUTOPILOT

COMPASS 1 REF TO DPU,

MPU, & UNSIK

NOTE: * BB-2-1448, 1450-1457, 1463: NO. 1 INVERTER CONTROL POWERS BY DUAL FED BUS NO. 1 (NOT GEN BUS)

NO. 2 INVERTER CONTROL POWERS BY DUAL FED BUS NO. 2 (NOT GEN BUS)

DUAL FED NO. 2 BUS

*

*

LEGEND

28 VDC POWER

115 VAC POWER

26 VAC POWER

GROUND

Figure 2-20. Inverters Control Circuit







FUEL CROSSFEED

RIGHT FIREWALL VALVE

RIGHT STANDBYFUEL PUMP

RIGHT AUX FUEL QTYWARNING & TRANSFER

RIGHT FUEL QUANTITY

RIGHT FUELPRESSURE WARNING

NAV MEMORY

HOT BATTERYBUS

ENTRY LT. CLOCK LT.& EXT POWER SENSE

BATTERY RELAY

L ENG FIRE EXT

R ENG FIRE EXT

MOD

RIGHT FIREWALLSHUTOFF VALVE

LEFT FIREWALLSHUTOFF VALVE

AVIONICSBUS NO. 1 TO AVIONICSMASTER CONTROL CB

AVIONICS NO. 1POWER RELAY

M A I N

B A T T E

R Y B U S

I S O L A T I O N B U S

LANDINGGEAR

MOTORLANDINGGEARRELAY

ISOLATION LIMITER

R GEN CONTROL

SHUNT

V LMETER

RIGHTSTARTERGEN

– +

–++

R GEN LINECONTACTOR

RIGHTSTARTRELAY


BATTERY

SHUNT

EXT POWERRELAY

EXTERNALPOWER

LEFTSTARTRELAY

EXT POWER CONNECTOR

ON

OFFBATTERYSWITCH

BATTERYRELAY

TO INVERTERCONTROL

BLUETESTJACK

INVERTERWARNRELAY

VOLTFREQ

METER

TO INVERTERSELECT SWITCH

LEFT GENERATOR BUS

26 VAC

115VAC

26 VAC

2 6 V A

C B

U S

TO AVIONICS

TO AVIONICS

115VAC

N O . 2 D U A L F E D B

U S

N O . 3 D U A L F E D

B U S


B U

S

2 6 V A

C

1 1 5 V A

C

2 6 V A C

1 1 5 V A C

N O . 1 D U A L F E D B

U S

INVERTERNO. 2

RIGHT START CONTROL

RIGHT IGNITOR POWER

PROPELLER GOVERNOR

MANUAL PROP DEICECONTROL

RIGHT MANUAL PROPDEICE

INVERTERNO. 1

AVIONICS BUS NO. 3

FWD ELECTRIC HEAT

PILOT'S WINDSHIELDANTI-ICE

CONDENSER BLOWER

ON

OFF

TO INVERTERCONTROLSWITCH

SHUNT

ISOLATIONLIMITER

LEFT GEN LINECONTACTOR

L. GEN CONTROL

V LMETER

– + LEFTSTARTERGEN

LEFT TORQUE METER

RIGHT TORQUE METER

YAW RATE

LEFT START CONTROL

RIGHT GEN CONTROL

CIGAR LIGHTER

RUDDER BOOSTCONTROL

CPILOT ILS INDICATOR

RIGHT BLEED AIRCONTROL

CABIN TEMPERATURECONTROL

CABIN READINGLIGHTS

AVIONIC & ENGINEINSTRUMENT LIGHTS

OVHD. SUBPANEL ANDCONSOLE LIGHTS

CABIN LIGHTS &ORDINANCE

R BLEED AIRWARNING

LANDING GEARPOSITION IND

ANNUNCIATORINDICATOR

R FUEL VENT HEAT

WINDSHIELD WIPER

R FUEL FLOWINDICATOR

R OIL PRESSUREINDICATOR

R OIL TEMPINDICATOR

R ENGINE FUELCONTROL HEAT

R ICE VANE CONTROL

R ICE VANE EMER

AUTOFEATHER

R CHIP DETECTOR

FURNISHINGS MASTERCONTROL

STALL WARNING HEAT

TAXI LIGHT

ICE LIGHTS

NAVIGATION LIGHT

RECOGNITION LIGHT

L PILOT HEAT

LANDING GEARCONTROL

R LANDING LIGHT

LEFT IGNITOR POWER

FLAP CONTROLAND INDICATOR

FLAP MOTOR

LEFT MANUAL PROPDEICE

LEFT FIREWALL VALVE

LEFT STANDBYFUEL PUMP

LEFT AUX FUEL QTYWARNING & TRANSFER

LEFT FUEL QUANTITY

LEFT FUELPRESSURE WARNING

LEFT GEN CONTROL

OUTSIDE AIR TEMP

PITCH TRIM

PILOT ILS INDICATOR

LEFT BLEED AIRCONTROL

CABIN PRESSURECONTROL

AUTOMATIC OXYGENCONTROL

CPILOT FLT INSTRLIGHTS

INSTRUMENT INDIRECTLIGHTS

PILOT FLT INSTRSIDE PANEL

OVHD FLOOD LIGHTS

L BLEED AIRWARNING

LANDING GEARWARNING HORN

ANNUNCIATOR POWER

L FUEL VENT HEAT

BRAKE DEICE

L FUEL FLOWINDICATOR

L OIL PRESSUREINDICATOR

L OIL TEMPINDICATOR

L ENGINE FUELCONTROL HEAT

L ICE VANE CONTROL

L ICE VANE EMER

FIRE DETECTION

L CHIP DETECTOR

PROP SYNCHROPHSERPROP BALANCE

BEACON LIGHTS

STROBE LIGHTS

TAIL FLOOD LIGHTS

PROP AUTOMATIC HEAT

L PILOT HEAT

L LANDING LIGHT

YAW DAMPER

AVIONICS MASTERCONTROL

PNEUMATIC SURFACEDEICE

STALL WARNINGSYSTEM





FUEL CROSSFEED

RIGHT FIREWALL VALVE


RIGHT AUX FUEL QTYWARNING & TRANSFER

RIGHT FUEL QUANTITY

RIGHT FUEL

PRESSURE WARNING

NAV MEMORY

HOT BATTERYBUS

ENTRY LT. CLOCK LT.& EXT POWER SENSE

BATTERY RELAY

L ENG FIRE EXT

R ENG FIRE EXT

MOD

RIGHT FIREWALLSHUTOFF VALVE

LEFT FIREWALL

SHUTOFF VALVE

AVIONICS

BUS NO. 1

TO AVIONICS

MASTER CONTROL CB


M A I N

B A T T E R Y B U S

I S O L A T

I O N

B U S

LANDINGGEAR

MOTORLANDINGGEARRELAY

ISOLATION LIMITER

R GEN CONTROL

SHUNT

V LMETER

RIGHTSTARTERGEN

– +

–++

R GEN LINECONTACTOR

RIGHTSTARTRELAY


BATTERY

SHUNT

EXT POWERRELAY

EXTERNALPOWER

LEFTSTARTRELAY

EXT POWER CONNECTOR

ON

OFFBATTERYSWITCH

BATTERYRELAY

RIGHT GENERATOR BUS


BLUETEST

JACKINVERTER

WARNRELAY

VOLTFREQ

METER

TO INVERTERSELECT SWITCH

LEFT GENERATOR BUS

26 VAC

115VAC

26 VAC

2 6 V A C B

U S

TO AVIONICS

TO AVIONICS

115VAC

N O . 2 D U A L

F E D

B U S

N O . 3 D U A

L F E D

B U S


B U S

2 6 V A C

1 1

5 V A C

2 6 V A C

1 1 5 V A C

N O . 1 D U A L F E D B U S

INVERTERNO. 2

RIGHT START CONTROL

RIGHT IGNITOR POWER

PROPELLER GOVERNOR

MANUAL PROP DEICECONTROL

RIGHT MANUAL PROPDEICE

INVERTERNO. 1

FWD ELECTRIC HEAT

PILOT'S WINDSHIELDANTI-ICE

CONDENSER BLOWER

ON

OFF

AVIONICS BUS NO. 3


SHUNT

ISOLATIONLIMITER

LEFT GEN LINECONTACTOR

L. GEN CONTROL

V LMETER

– + LEFTSTARTERGEN

LEFT TORQUE METER

RIGHT TORQUE METER

YAW RATE

LEFT START CONTROL

RIGHT GEN CONTROL

NO. 2 INV CONTROL

CIGAR LIGHTER

RUDDER BOOSTCONTROL

CPILOT ILS INDICATOR



CABIN READINGLIGHTS

AVIONIC & ENGINEINSTRUMENT LIGHTS

OVHD. SUBPANEL ANDCONSOLE LIGHTS

CABIN LIGHTS &ORDINANCE

R BLEED AIRWARNING

LANDING GEAR

POSITION IND


R FUEL VENT HEAT

WINDSHIELD WIPER

R FUEL FLOWINDICATOR

R OIL PRESSUREINDICATOR

R OIL TEMPINDICATOR

R ENGINE FUELCONTROL HEAT

R ICE VANE CONTROL

R ICE VANE EMER

AUTOFEATHER

R CHIP DETECTOR


STALL WARNING HEAT

TAXI LIGHT

ICE LIGHTS

NAVIGATION LIGHT

RECOGNITION LIGHT

L PILOT HEAT

LANDING GEARCONTROL

R LANDING LIGHT

LEFT IGNITOR POWER

FLAP CONTROL

AND INDICATOR

FLAP MOTOR

LEFT MANUAL PROPDEICE

LEFT FIREWALL VALVE


LEFT AUX FUEL QTYWARNING & TRANSFER

LEFT FUEL QUANTITY

LEFT FUEL

PRESSURE WARNING

LEFT GEN CONTROL

NO. 1 INV CONTROL

OUTSIDE AIR TEMP

PITCH TRIM

PILOT ILS INDICATOR



AUTOMATIC OXYGENCONTROL

CPILOT FLT INSTRLIGHTS

INSTRUMENT INDIRECTLIGHTS

PILOT FLT INSTRSIDE PANEL

OVHD FLOOD LIGHTS

L BLEED AIRWARNING

LANDING GEAR

WARNING HORN

ANNUNCIATOR POWER

L FUEL VENT HEAT

BRAKE DEICE

L FUEL FLOWINDICATOR

L OIL PRESSUREINDICATOR

L OIL TEMPINDICATOR

L ENGINE FUELCONTROL HEAT

L ICE VANE CONTROL

L ICE VANE EMER

FIRE DETECTION

L CHIP DETECTOR

PROP SYNCHROPHSERPROP BALANCE

BEACON LIGHTS

STROBE LIGHTS

TAIL FLOOD LIGHTS

PROP AUTOMATIC HEAT

L PILOT HEAT

L LANDING LIGHT

YAW DAMPER



STALL WARNINGSYSTEM





AVIONICSBUS NO. 1

TO AVIONICSMASTERCONTROL CB

ON

LEFT GENCONTROL

LINECONTACTOR

SHUNT

ISOLATION LIMITER

EXT POWER CONNECTOR

OFF


LEFTSTARTER

GEN

RIGHTSTARTER

GEN

LEFTSTARTRELAY

HOT BATTERYBUS

BATTERYSW

CIRCUITEDINTOEXTERNALPOWERRELAY

EXT POWER RELAY

BATTERYRELAY

M A I N

B A T T E R Y B U S

I S O L A T I O N B U S

BATTERY SW

ENTRY LTS &CLOCK LT &EXT PWRSENSE

RNAV MEMORY(OPT)

STEREO (OPT)

L ENG FIRE EXT

RIGHT FIREWALLSHUT OFF VALVE

LEFT FIREWALLSHUT OFF VALVE


REMOVED

FROM HOTBUS ONBB1098 ANDAFTER


R ENG FIRE EXT

THIS LINE OFFUSES CHANGETO 5A CIRCUITBREAKERS ONBB1098 ANDAFTER

BATTERY

S H U N T

BATTERYCHARGESENSOR

VOLTLOAD

METER

VOLTLOAD

METER

SHUNT

RIGHT GENCONTROL

RIGHTSTARTRELAY

ISOLATION LIMITER

LINECONTACTOR

AVIONICSBUS NO 2

TO INVCONTROLDUAL FEDBUS NO. 2

AVIONICS NO. 3POWER RELAY(OPTIONAL)

AVIONICSBUS NO 3

BLUETESTJACK

TO INVCONTROLDUALFED BUSNO. 1

LEFT GEN BUS

RELAYPANEL

INVNO. 1

INVWARN

RELAY

26 VAC

115VAC

VOLTSFREQ.METER

N O 2

D U A L F E D

B U S


B U S

N O . 4 D U A L F E

D B U S

2 6

V A C

1 1 5 V A C

N O 1

D U A L F E D

B U S

S U B P A N E L S

LEFT LANDING LIGHT

LEFT PITOT HEAT

PROP AUTOMATICHEAT SWITCH

TAIL FLOOD LIGHTSSWITCH (OPT)

INVNO. 2

A V I O N I C S

A V I O N I C S

- +

LEFT GEN CONTROL

LEFT RADIANT HEAT

PILOT'SWINDSHIELDANTI-ICE

CONDENSER

BLOWER

NO. 1 INVERTERCONTROL

PITCH TRIM

CIGARETTE LIGHTER


RIGHT GENERATORCONTROL


RUDDER BOOSTCONTROL

COPILOT TURN ANDSLIP


CABIN TEMP CONTROL

CABIN PRESSURELOSS (OPT)

PILOT'S TURN ANDSLIP

ENCODER ALTIMETER(OPT)

YAW DAMPER



AUTOMATIC OXYCONTROL

BRAKE DEICE (OPT)

WINDSHIELD WIPER

LEFT TORQUEMETER

RIGHT TORQUEMETER

YAWRATE

FUELCROSSFEED

RIGHT FUELPRESSUREWARNING

RIGHT FUELQUANTITY

RIGHT AUX FUELQUANTITYWARNING ANDTRANSFER


RIGHT FIREWALLVALVE

LEFT FIREWALLVALVE


LEFT AUX FUELQUANTITYWARNING ANDTRANSFER

LEFT FUELQUANTITY

LEFT FUELPRESS WARN

LEFT STARTCONTROL

LEFT IGNITORPOWER

FLAP CONTROLAND INDICATOR

FLAP MOTOR

LEFT MANUALPROP DEICE

RIGHT STARTERCONTROL

RIGHT IGNITORPOWER

PROPELLERGOVERNOR

RIGHT MANUALPROP DEICE

MANUAL PROPDEICE CONTROL

RIGHT FUEL VENT HEAT


LANDING GEARPOSITION IND

RIGHT BLEED AIRWARNING

FLOURESCENT LIGHTSAND ORD WARNING

(OVHD) SUBPANEL ANDCONSOLE LIGHTS

AVIONICS & ENGINEINSTRUMENT LTS

CABIN READING LIGHTS

AUTOFEATHER

RIGHT CHIP DETECTOR

RIGHT GEN OVERHEAT(OPT)

R ICE VANE CONT

RIGHT OIL PRESSUREWARN (OPT)

RIGHT OILTEMPERATUREINDICATOR

RIGHT OIL PRESSUREINDICATOR

RIGHT FUEL FLOWINDICATOR

STALL WARNING HEAT

RIGHT PITOT HEAT

LANDING GEAR CONTROL

RECOGNITION LIGHT

NAV LIGHT SWITCH

ICE LIGHTS

TAXI LIGHT SWITCH

RIGHT LANDING LIGHTSWITCH

RIGHT ENGINE FUELCONTROL HEAT

LEFT FUEL VENT HEAT

ANNUNCIATORPOWER


LEFT BLEED AIRWARNING

STALL WARNINGSYSTEM

OVERHEAD AND SIDEPANEL LIGHTS

INSTRUMENTINDIRECT LIGHTS

FLIGHT & GYROINSTRUMENT LIGHTS

TAIL FLOOD LIGHT(OPT)

PROPSYNCHROPHASER

LEFT CHIP DETECTOR

LEFT GEN OVERHEAT

FIRE DETECTION

L ICE VANE CONT

LEFT ENGINE FUELCONT HEAT

LEFT OILTEMPERATURE

LEFT OIL PRESSUREWARNING (OPT)

LEFT OIL PRESSURE

LEFT FUEL FLOW


BEACON LIGHTSSWITCH

STROBE LIGHTSSWITCH


+ +

+

–

–

+ –





AVIONICSBUS NO. 1

TO AVIONICSMASTERCONTROL CB

ON

LEFT GENCONTROL

REVERSECURRENTPROTECTION

SHUNT

ISOLATION LIMITER

EXT POWER CONNECTOR

OFF



LEFTSTARTER

GEN

RIGHTSTARTER

GEN

LEFTSTARTRELAY

HOT BATTERY

BUS

SMALLPIN

EXT POWER RELAY

BATTERYRELAY

M A I N

B A T T E R Y B U S

I S O L A T I O N

B U S

BATTERY SW

LEFT FIRE EXT

THRESHOLD LT

RIGHT FIREWALLSHUT OFF VALVE

LEFT FIREWALLSHUT OFF VALVE



RIGHT FIRE EXT

BATTERY

BATTERYCHARGESENSOR

VOLTLOAD

METER

VOLTLOAD

METER

SHUNT REVERSECURRENTPROTECTION

RIGHT GENCONTROL

RIGHTSTARTRELAY

ISOLATION LIMITER

TO ANNUNCIATORADVISORY LIGHT

EXTERNAL

POWER PLUGENGAGED

BB 364AND SUBSEQUENT

AVIONICSBUS NO 2

COPILOT'SWINDSHIELD ANTI-ICE

RIGHT RADIANTHEAT

RIGHT GEN BUS

TO INVCONTROL

NO. 2 DUAL FEDBUS

VENTBLOWER POWER

AFT EVAPORATORBLOWER POWER


AVIONICSBUS NO 3

LEFT FUELQUANTITY

RIGHT FUELQUANTITY

*L FUEL FLOW

LEFT TORQUE

METER

RIGHT TORQUEMETER

*R FUEL FLOW

INVWARNRELAY

BLUETESTJACK

TO INVCONTROLNO. 1 DUALFED BUS

LEFT GEN BUS

INVSELECT

RELAY

INVNO. 1

26 VAC115VAC

FUELCROSSFEED

RIGHT FUELPRESSUREWARNING

RIGHT AUX FUELQUANTITYWARNING ANDTRANSFER


RIGHT FIREWALLVALVE

LEFT FIREWALLVALVE


LEFT AUX FUELQUANTITYWARNING ANDTRANSFER

LEFT FUELPRESSUREWARNING

LEFT STARTERCONTROL

LEFT IGNITORPOWER

FLAP CONTROL& INDICATOR

FLAP MOTOR

LEFT MANUALPROP DEICE

**L FUEL FLOW

N O 2

D U A L F E D

B U S


B U S


B U S

2 6 V A C

1 1 5 V A C

RIGHT STARTERCONTROL

RIGHT IGNITORPOWER

PROPELLERGOVERNOR

RIGHT MANUALPROP DEICE

MANUAL PROPDEICE CONTROL

N O 1

D U A L F E D

B U S

S U B

P A N E L S

FUEL DRAINCOLLECTOR PUMPS

LEFT GENCONTROL

NO 1 INVERTERCONTROL

TRIM TAB

PILOT'S TURN ANDSLIP INDICATOR

OPTIONALALTIMETER

PNEUMATICSURFACE DE-ICE

LEFT FUELVENT HEATER

ANNUNCIATORPOWER


LEFT BLEEDAIR WARNING

STALL WARNINGSYSTEM

OVERHEAD ANDSIDE PANEL LIGHTS

INSTRUMENTINDIRECT LIGHT

FLIGHTINSTRUMENT LIGHT

LOGO LIGHT


LEFT BLEEDAIR CONTROL

PROPSYNCHROPHASER

FIRE DETECTION

LEFT ICE VANE

LEFT ENGINE FUELCONTROL HEAT

LEFT OIL TEMPINDICATOR

LEFT OIL PRESSUREINDICATOR

AVIONICSMASTER CONTROL

STROBE LIGHTS

BEACONLIGHTS SW

PROP DEICEAUTO HEAT SW

LEFT PITOTHEAT SW

LEFT LANDINGLIGHT SW

RIGHT GENCONTROL


RUDDER BOOSTSYSTEM

VACUUM DAMPERSYSTEM

COPILOT'S TURN &SLIP INDICATOR

WINDSHIELD WIPER

RIGHT FUELVENT HEATER


LANDING GEARPOSITION INDICATOR

RIGHT BLEEDAIR WARNING

CABIN FASTEN SEATBELT & NO SMOKINGSIGN AND CHIMES

SUBPANEL ANDCONSOLE LIGHTS

RADIO & ENGINEINSTRUMENT LIGHTS


RIGHT BLEEDAIR CONTROL

RIGHT ENGINE HEATFUEL CONTROL

RIGHT OIL TEMPINDICATOR

RIGHT OIL PRESSUREINDICATOR

STALL WARNINGHEAT

RIGHT PITOTHEAT SW

LANDING GEARCONTROL

NAV LIGHT SW

ICE LIGHT SW

TAXI LIGHT SW

RIGHT LANDINGLIGHT SW

AUTOFEATHER

CHIP DETECTOR

RIGHT ICE VANE

**R FUEL FLOW

INVNO. 2

A V I O N I C S

A V I O N I C S

LEFT RADIANTHEAT

PILOT'SWINDSHIELDANTI-ICE

CONDENSERBLOWER

S H U N T

OPTIONALEQUIPMENT

***





LIMITATIONS

GENERATOR LIMITS (250AMPERES)

Maximum sustained generator load (Table 2-1) is limited as follows:

In Flight:

Sea Level to31,000 feet altitude .................. 1.00 (100%)

Above 31,000 feet altitude ....... 0.88 (88%)

Ground Operation ................... 0.85 (85%)

During ground operation, also observe thelimitations in Table 2-1.

STARTERS

Use of the starter is limited to 40 seconds ON,60 seconds OFF, 40 seconds ON, 60 secondsOFF, 40 seconds ON, then 30 minutes OFF.

INVERTERS

Due to avionics equipment requirements, the115-volt inverter output must be 105-120 VAC,380-420 Hz.

CIRCUIT BREAKERS

Tables 2-3 to 2-4 give circuit breaker titles, val-ues, and the circuits that they control. They aregrouped by panel location.

GENERATOR LOAD MINIMUM GAS GENERATOR RPM – N1

WITHOUT AIR WITH AIR CONDITIONING CONDITIONING

0 to 70% 52% 60%

70 to 75% 55% 60% 75 to 80% 60% 60%

80 to 85% 65% 65%

BB – 1439, 1444 AND SUBSEQUENT

0 to 75% 61% 62%

75 to 80% 61% 62%

80 to 85% 65% 65%

Table 2-1. LIMITATIONS—GROUND OPERATIONS





CIRCUIT BREAKER NAME CAPACITY PROVIDES POWER TO

1. FUEL SYSTEM

A. AUX TRANSFER (L & R) 5 AMP TRANSFER SELECT SWITCH

NO TRANSFER LIGHT

AUX TANK FLOAT SWITCH

MOTIVE FLOW VALVE

B. CROSSFEED 5 AMP CROSSFEED SWITCHCROSSFEED VALVE

AUX FUEL TRANSFER MODULE

C. FIREWALL VALVE (L & R) 5 AMP FIREWALL VALVE SWITCH

FIREWALL VALVE

D. FUEL PRESSURE WARNING(L & R) 5 AMP FUEL PRESS SWITCH

FUEL PRESS WARNING LIGHT

AUX FUEL TRANSFER MODULE

E. FUEL QUANTITY INDICATOR (L & R) 5 AMP INDICATOR POWER

F. STANDBY PUMP (L & R) 10 AMP STANDBY PUMP SWITCH

AUX TRANSFER PCB

G. BUS FEEDERS

NO. 3 (L & R) 50 AMP NO. 3 DUAL-FED BUS

NO. 4 (L & R) 50 AMP NO. 4 DUAL-FED BUS

2. FLAP

A. MOTOR 20 AMP MOTOR RELAY AND MOTOR POWER

B. CONTROL 5 AMP FLAP POTENTIOMETER (POSITION XMTR)

SPLIT FLAP

HOBBS METER

FLAP POSITION INDICATOR

3. PROP

A. GOVERNOR 5 AMP OVERSPEED GOVERNOR TEST SWITCH

B. PROP DEICE

(1) CONTROL 5 AMP MANUAL SWITCH POWER

(2) PROP (L & R) 20 AMP DEICE POWER

25 AMP DEICE POWER (BB-1439, 1444 AND SUBSEQUENT)

Table 2-2. FUEL CONTROL CIRCUIT-BREAKER PANEL






5. ENGINE INSTRUMENTS

(BB-1484, 1486 AND

SUBSEQUENT)

A. ITT (L & R) 5 AMP GAGE POWER

B. TORQUE (L & R) 5 AMP GAGE POWER

C. PROP TACH (L & R) 5 AMP GAGE POWERD. TURBINE TACH (L & R) 5 AMP GAGE POWER

E. FUEL FLOW (L & R) 5 AMP GAGE POWER

F. OIL PRESS (L & R) 5 AMP GAGE POWER

G. OIL TEMP (L & R) 5 AMP GAGE POWER

Table 2-2. FUEL CONTROL CIRCUIT-BREAKER PANEL (Cont)


1. ELECTRICAL DISTRIBUTION

A. NO. 1 BUS FEEDERS (L & R) 50 AMP NO. 1 DUAL FED BUS

B. NO. 2 BUS FEEDERS (L & R) 50 AMP NO. 2 DUAL FED BUS

C. GEN CONTROL (L & R) 10 AMP GENERATOR CONTROL SWITCH

GENERATOR CONTROL PANEL

2. INVERTER CONTROL

(BB-1439, 1444-1448,

1450-1457 AND PRIOR)

A. NO. 1 5 AMP NO. 1 INVERTER CONTROL SWITCH

AND CONTROL RELAY

B. NO. 2 5 AMP NO. 2 INVERTER CONTROL SWITCH

AND CONTROL RELAY

3. ENGINE

A. AUTOFEATHER 5 AMP POWER LEVER ARM SWITCHES

AUTOFEATHER ARM SWITCHES (400 & 200 FT-LB)

B CHIP DETECTOR 5 AMP BB 1 162 CHIP DETECTOR AND LIGHT

Table 2-3. RIGHT SIDE CIRCUIT-BREAKER PANEL






E. FUEL CONTROL HEAT 7.5 AMP L & R FUEL CONTROL HEAT SWITCH

(THROTTLE QUADRANT)

L & R FUEL CONTROL PNEUMATIC LINE HEAT

(PRIOR TO BB-1444, EXCEPT 1439)

F. FUEL DRAIN COLLECTOR PUMP 5 AMP L & R COLLECTOR FLOAT SWITCH AND

COLLECTOR PUMP (BB 2-665)G. FUEL FLOW (L & R) 2 AMP BB 2-224 FUEL FLOW INDICATOR AND

TRANSMITTER (AC)

5 AMP BB 225 – 1483, 1485 FUEL FLOW INDICATOR

AND TRANSMITTER (DC)

H. ENGINE INSTRUMENT POWER 7.5 AMP BB-1484, 1486 AND SUBSEQUENT

I. ICE VANE CONTROL (L & R)

(PRIOR TO BB-1444, EXCEPT 1439) 5 AMP L & R ICE VANE CONTROL SWITCH

ICE VANE SENSE MODULE

ICE VANE ACTUATOR

J. MN ENG ANTI-ICE (L & R)

(BB-1439, 1444 AND SUBSEQUENT) 5 AMP L & R MN ENG ANTI-ICE CONTROL SWITCH

MN ENG ANTI-ICE SENSE MODULE

MN ENG ANTI-ICE ACTUATOR

K. STBY ENG ANTI-ICE (L & R)

(BB-1439, 1444 AND SUBSEQUENT) 5 AMP L & R STBY ENG ANTI-ICE CONTROL SWITCH

STBY ENG ANTI-ICE SENSE MODULE

STBY ENG ANTI-ICE ACTUATOR

L. OIL PRESS (L & R)

(PRIOR TO BB-1486, EXCEPT 1484) 5 AMP OIL PRESSURE INDICATOR AND TRANSMITTER

M. OIL TEMP (L & R)

(PRIOR TO BB-1486, EXCEPT 1484) 5 AMP OIL TEMP INDICATOR AND TRANSMITTER

N. TORQUEMETER (L & R)

(PRIOR TO BB-1486, EXCEPT 1484) 2 AMP TORQUE INDICATOR AND TRANSMITTER AC

O. PROP SYNC 5 AMP PROP SYNCHROPHASER CONTROL BOX

SYNCH CONTROL SWITCH

Table 2-3. RIGHT SIDE CIRCUIT-BREAKER PANEL (Cont)







B. OXYGEN 5 AMP BB 54 AND SUBSEQUENT – PASSENGER O2

MASKS 12,500 FT PRESSURE SWITCH

C. PRESS CONT 5 AMP LEFT SQUAT SWITCH

PEDESTAL PRESS CONTROL SWITCH

SAFETY VALVE DUMP SOLENOID

EVAPORATOR DOOR SOLENOIDCABIN DOOR SOLENOID

D. TEMP CONTROL 5 AMP VENT BLOWER CONTROL SWITCH

LEFT SQUAT SWITCH

AMBIENT AIR VALVES AND PCB

BB 1-450 RADIANT HEAT CONTROL SWITCH

BB 450 AND SUBSEQUENT RADIANT HEAT

POWER CIRCUIT BREAKER

CABIN TEMP MODE SELECTOR SWITCH

5. FLIGHT

A. PITCH TRIM 5 AMP PEDESTAL ELECTRIC ELEVATOR

TRIM SWITCH

TRIM MOTOR

B. RUDDER BOOST 5 AMP PEDESTAL ON/OFF SWITCH

DIFFERENTIAL PRESSURE SWITCH

RUDDER BOOST SOLENOIDS

C. TURN AND SLIP 5 AMP TURN AND SLIP INDICATOR

D. ENCODING ALTIMETER 1 AMP ALTIMETER (ENCODING)

6. LIGHTS

A. AVIONICS AND ENG INST 5 AMP RADIO AND ENGINE INSTRUMENT LIGHTS

PILOT AND COPILOT CLOCK AND MAP LIGHTS

B. FLIGHT INST 7.5 AMP OVERHEAD PANEL AND TERMINAL BOARD

PILOT & COPILOT FLIGHT INST LIGHTS

C. FSB & NO SMOKE CABIN 5 AMP FASTEN SEAT BELT/CABIN NO SMOKING LIGHTS

CABIN FLUORESCENT LIGHTS

CABIN WARNING CHIME







7. WARNING

A. ANNUN INDICATOR 5 AMP ANNUN N1 LOW LIGHT

ANNUN INVERTER OUT LIGHT

ICE VANE PCB

O2 PRESS SWITCH

CABIN ALT WARN PRESS SWITCH (12,500)BATTERY CHARGE MODULE

DUCT OVERTEMP SWITCH

ALTITUDE WARNING LIGHT

B. ANNUN POWER 5 AMP 28V ANNUNCIATOR CONTROL CARD

MASTER WARNING LIGHTS

MASTER CAUTION LIGHTS

CAUTION LEGEND SWITCHC. BLEED AIR WARNING (L & R) 5 AMP BLEED AIR WARN LIGHTS

BLEED AIR WARN PRESS SWITCH

D. LANDING GEAR INDICATOR 5 AMP GREEN GEAR DN LIGHTS

RED GEAR HANDLE LIGHTS

E. LANDING GEAR WARNING 5 AMP GEAR WARNING HORN & FLASHER

GEAR WARNING HORN SILENCE BUTTON & RELAY

F. STALL WARNING 5 AMP POWER TO STALL WARNING LIFT COMPUTER

8. WEATHER

A. BRAKE DEICE 5 AMP LEFT UPLOCK SWITCH

BATTERY CHARGE/ DEICE MODULE

BRAKE DEICE SWITCH

DEICE BLEED AIR VALVES

B. FUEL VENT HEATERS (L & R) 5 AMP HEATER SWITCH

HEATER ELEMENTS

C. SURFACE DEICE 5 AMP SURFACE DEICE SWITCH

DEICE DISTRIBUTOR VALVE

TIME DELAY PCB

D. WINDSHIELD WIPERS 10 AMP OVERHEAD PANEL SWITCH





Table 2-4. PILOT’S RIGHT SUBPANEL CIRCUIT-BREAKER SWITCHES


1. ICE PANEL

A. PITOT (L & R) 7.5 AMP POWER TO PITOT ELEMENTS

B. PROP (AUTO/OFF) 20 AMP POWER TO PROP DEICE AMMETER

AND DEICE TIMER

25 AMP BB-1439, 1444 AND SUBSEQUENT

C. STALL WARN 15 AMP STALL WARNING HEAT CONTROL RELAY2. LANDING GEAR

A. LANDING GEAR RELAY 5 AMP LANDING GEAR CONTROL SWITCH

RVS NOT READY ANNUNCIATOR POWER

B. LANDING GEAR RELAY

(HYDRAULIC GEAR) 2 AMP LANDING GEAR CONTROL SWITCH

HYD FLUID LOW LIGHT

RVS NOT READY ANNUNCIATOR POWER

3. LIGHTS

A. ICE 5 AMP ICE LIGHTS

B. LANDING LIGHTS (L & R) 10 AMP LANDING LIGHTS

C. NAV LIGHT 5 AMP NAV LIGHTS

D. TAIL FLOODLIGHT 15 AMP TAIL FLOODLIGHTS

OVERHEAD PANEL LIGHTS

E. RECOG LIGHTS 15 AMP OR BB 50-177 RECOG LIGHT

RELAY AND LIGHTS (2 BULB)

7.5 AMP BB 178 AND SUBSEQUENT RECOG LIGHT

(1 BULB)

F. TAXI LIGHT 15 AMP TAXI LIGHT

G. BEACON 10 AMP BEACONS

H. STROBE 5 AMP STROBE POWER SUPPLY & STROBE TUBE





1. What is the rat ing for the battery?

A. 28-volt, 24 ampere-hour

B. 24-volt, 34/36 ampere-hour

C. 28-volt, 34/36 ampere-hour

D. 24-volt, 42 ampere-hour

2. Where is the bat tery 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

3. If the amber BATTERY CHG annuncia-

tor illuminates in flight, what initial ac -tio n does the checklist say to do?

A. Turn the battery switch OFF

B. Reduce the electrical load

C. Isolate the battery bus

D. Reset the annunciator

4. What is the individual generator rating?

A. 30-volt, 200-ampere

B. 24-volt, 300-ampere

C. 28-volt, 250-ampere

D. 32-volt, 250-ampere

5. Where are the generator switches located?

A. Under a gang bar on the overheadpanel

B. On the center instrument panel

C. Under a gang bar on the pilot’s leftsubpanel

D On the copilot’s subpanel

6. On airplanes with Serial Nos. BB-88 andsubsequent, how is a generator turned on?

A. Move the switch to OFF, then to ON

B. Hold the switch to RESET for onesecond and release to ON

C. Move the switch to ON

D. Hold the switch to ON for one second

7. When an engine is being started, in whatposition should its GEN switch be?

A. RESET

B. ON

C. OFF

8. When a generator is off line, whatindication is present?

A. An amber DC GEN light is on

B. No indications are present

C. A green DC GEN light is on

D. A red DC GEN light is on

9. Where is the external power connector lo-cated?

A. Under the left wing

B. On the left aft fuselage

C. Under the right wing, outboard of theengine nacelle

D. On the right forward fuselage

10. How much continuous current should theexternal power unit be capable of sup-plying?

A. 100 amperes

QUESTIONS

11 Wh i di i i id d l h 16 H i h ?





11. What indication is provided to alert theoperator 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

12. What is the minimum required battery volt-

age before using an external power unit?A. 28 vol ts

B. 24 vol ts

C. 22 vol ts

D. 20 vol ts

13. Why should the BATtery switch be left onduring use of external power?

A. To charge the battery

B. To energize the generator control units(GCU)

C. To power the volt-loadmeters

D. To protect against voltage spikes

14. To protect the battery from damage, the

external power voltage must not go over what value for more than two minutes?

A. 35 vol ts

B. 30 vol ts

C. 25 vol ts

D. 24 vol ts

15. Prior to BB-1632, if the BATTERY CHGlight does not extinguish in flight whenthe battery control switch is placed toOFF, what action should be taken?

A. Pull the battery circuit breaker

B R h BAT i h

16. How many inverters are there?

A. 1

B. 2

C. 3

D. 4

17. 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, 400 Hz

D. 30-volt and 115-volt, 120 Hz

18. Where is the INVERTER switch located?

A. On the copilot’s right subpanel

B. On the overhead panel

C. On the copilot’s right sidepanel

D. On the pilot’s left subpanel

19. After starting the right engine and turn-ing the right generator on, what should theloadmeter reading decrease to beforestarting the left engine?

A. 25% (.25 )

B. 50% ( .50 )

C. 75% ( .75 )

D. 100% (1.00)

20. 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,

CHAPTER 3





CHAPTER 3

LIGHTING

CONTENTS

Page

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

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

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

Cockpit............................................................................................................................. 3-3

Cabin ................................................................................................................................ 3-5

EXTERIOR LIGHTS .............................................................................................................. 3-7Landing Lights ................................................................................................................. 3-7

Taxi Light......................................................................................................................... 3-7

Wing Ice Lights................................................................................................................ 3-7

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

Recognition Lights........................................................................................................... 3-7

Beacon Lights .................................................................................................................. 3-7

Strobe Lights.................................................................................................................... 3-7

Tail Floodlights ................................................................................................................ 3-7

Airstair Floodlight............................................................................................................ 3-8Under Step Lighting......................................................................................................... 3-8

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

ILLUSTRATIONS





ILLUSTRATIONS

Figure Title Page

3-1 Overhead Lighting Controls..................................................................................... 3-2

3-2 Copilot’s Left Subpanel............................................................................................ 3-2

3-3 Pilot’s Right Subpanel.............................................................................................. 3-3

3-4 Instrument and Panel Lights..................................................................................... 3-3

3-5 Console Lights.......................................................................................................... 3-4

3-6 Overhead Subpanel Lights ....................................................................................... 3-4

3-7 Copilot’s Instrument Lights...................................................................................... 3-4

3-8 OAT Gage................................................................................................................. 3-5

3-9 Free Air Temperature Switch ................................................................................... 3-5

3-10 Fluorescent Light Switch.......................................................................................... 3-5

3-11 Passenger Warning Sign........................................................................................... 3-6

3-12 Reading Lights ......................................................................................................... 3-6

3-13 Threshold, Aisle, and Baggage Lights ..................................................................... 3-6

3-14 Landing and Taxi Lights........................................................................................... 3-7

3-15 Exterior Lights.......................................................................................................... 3-8

3-16 Airstair Floodlight .................................................................................................... 3-8

3-17 Under Step Lighting ................................................................................................. 3-8

CHAPTER 3





INTRODUCTION

The instruments are illuminated either internally or with post-type lights. General cabinlighting consists of overhead fluorescent lights and individual passenger reading lights.A passenger FASTEN SEAT BELT–NO SMOKING sign is provided. Both the airstair and baggage area are illuminated. Exterior lights consist of landing, taxi, ice inspection,navigation, recognition, beacon, strobe, and lights for the area around the airstai r door.

Optional lighting is available to illuminate the vertical tail fin.

GENERAL

EXIT

CHAPTER 3LIGHTING





A FREE–AIR–TEMPERATURE gage is lo-

cated on the left sidewall aft of the fuel panel.For BB-1439, 1444 and subsequent, a digitaldisplay indicates the free air temperature inCelsius. Prior to BB-1444, excluding 1439, ananalog temperature display also indicates the

3 3N3

F O R

0

3 0

6 0

9 0 1 2 0 1 50 1 8 0

2 1 0

2 4 0

2 7 0

3 0 0

3 3 0

COMPASS CORRECTIONCALIBRATE WITH

RADIOON

S T E E

MAX GEAR EXTENSION

MAX GEAR RETRACT

MAX GEAR EXTENDED

MAX APPROACH FLAP

MAX FULL DOWN FLAP

MAX MANEUVERING

181 KNOTS

163 KNOTS

181 KNOTS

200 KNOTS

157 KNOTS

181 KNOTS

AIRSPEEDS (IAS)

OPERATION LIMITATIONSTHIS AIRPLANE MUST BE OPERATED AS A NORMAL CATEGORY AIRPLANE IN COMPLIANCE WITHTHE OPERATING LIMITATIONS STATED IN THE FORM OF PLACARDS, MARKINGS AND MANUALS

NO ACROBATIC MANEUVERS INCLUDING SPINS ARE APPROVEDTHIS AIRPLANE APPROVED FOR VFR, IFR, & DAY & NIGHT OPERATION AND IN ICING CONDITIONS

CAUTIONSTALL WARNING IS INOPERATIVE WHEN MASTER SWITCH IS OFF

STANDBY COMPASS IS ERRATIC WHEN WINDSHIELD ANTI-ICE AND/OR AIR CONDITIONING IS ON

DO NOT OPERATE

ON DRY GLASS

WINDSHIELD WIPERS

OFF

PARK SLOW

FAST

OFF

MASTER

PANEL

LIGHTS

ON

OVERHEAD

FLOOD

LIGHTS

OFFBRT

INSTRUMENT

INDIRECT

LIGHTS

OFFBRT

AVIONICS

PANEL

LIGHTS

OFFBRT

ENGINE

INSTRUMENT

LIGHTS

OFFBRT

PILOT

FLIGHT

LIGHTS

OFFBRT

OVERHEAD

SUB PANEL

& CONSOLE

LIGHTS

OFFBRT

SIDE

PANEL

LIGHTS

OFFBRT

COPILOT GYRO

INSTRUMENT

LIGHTS

OFFBRT

COPILOT

FLIGHT

LIGHTS

OFFBRT

40 608020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 608020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTS

AC VOLTS

FREQ

Beechcraft

08020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

-60

0

+60BATT AMPS

Figure 3-1. Overhead Lighting Controls



A pushbutton switch next to each light con- INTERIOR LIGHTING



p gtrols the individual passenger reading lights

along the top of the cabin.

A switch just inside the airstair door aft of thedoorframe controls a baggage-area light.

A threshold light located forward of the airstair door at floor level, and an aisle light locatedat floor level aft of the spar cover, are con-trolled by a switch next to the threshold light.

When the airstair is open, these lights comeon; when it is closed and locked, they auto-matically extinguish. A flush-mounted flood-light forward of the flaps in the bottom of theleft wing and under the stair lights are also con-trolled by the threshold light switch. They il-luminate the area around the airstair when itis open and the switch is turned on.

Switches for the landing lights, taxi light,ice lights, navigation lights, recognitionlights, beacons, and strobe lights are locatedon the pilot’s right subpanel (Figure 3-3).They are appropriately labeled as to the spe-cific function.

Tail floodlights , if installed, are controlled bya switch located either on the overhead panel

or the pilot’s right subpanel.

INTERIOR LIGHTING

COCKPIT

Overhead FloodlightsThese lights are designed to give general il-lumination for the cockpit area and are con-trolled by a rheostat on the overhead panellabeled OVERHEAD FLOODLIGHTS.

Instrument Indirect LightsThese lights are located under the glareshieldand illuminate the instrument panel.

Pilot Flight LightsThe PILOT FLIGHT LIGHTS rheostat for the

pilot’s flight instrument area controls the in-ternal or eyebrow post lights. The flight lightsare shown in Figure 3-4.

Pilot Gyro Instrument LightsFor EFIS equipped aircraft, the EADI andEHSI intensity are controlled by the EFISdimming rheostats.Figure 3-3. Pilot’s Right Subpanel

Figure 3-4. Instrument and Panel Lights

Avionics Panel Lights Side Panel Lights





Avionics Panel Lights

The AVIONICS PANEL LIGHTS rheostatcontrols the illumination of the internal avion-ics panel lights.

Overhead Subpanel andConsole LightsThe rheostat labeled OVERHEAD SUB-PANEL & CONSOLE LIGHTS controls the

lighting for the overhead subpanel and thethrottle console (Figure 3-5 and Figure 3-6).

Side Panel Lights

A rheostat labeled SIDE PANEL LIGHTS con-trols the lights on the left and right side panels.

Copilot’s Gyro InstrumentLightsTh e CO P I LO T G Y RO I N S TRU MEN TLIGHTS rheostat controls any gyro instru-ments on the copilot ’s flight panel.

Copilot’s Flight LightsThe COPILOT FLIGHT LIGHTS rheostat for the copilot’s flight instrument area controls theexternal eyebrow or post lights (Figure 3-7).

Figure 3-7. Copilot’s Instrument Lights

Figure 3-5. Console Lights

40 608020

0 3010 20

0

PUSH

100

DC VOLTS

% LOAD40 60

8020

0 3010 20

0

PUSH

100

DC VOLTS

% LOAD400 410

420380390

100 130110 120

PUSH AC VOLTS

FREQ0

8020

0 3020

0

PUSH

100

DC VOLTS-60

0

+60BATT AMPS



Master Panel Lights Switch



g

The switch labeled MASTER PANEL LIGHTScontrols power to the overhead light controlpanel, fuel control panel, engine instruments,radio panel, both subpanels, the console, andthe pilot’s and copilot’s instrument lights.These lights can be adjusted individually withthe individual rheostats, but are convenientlyshut off or turned on with this MASTERPANEL light switch (Figure 3-1).

Free Air Temperature SwitchOn BB-1439, 1444 and subsequent aircraft, aseven segment digital display, located on thesidewall, indicates the free air temperature inCelsius. When the adjacent button is depressed,Fahrenheit is displayed (Figure 3-8).

Prior to BB-1444, excluding 1439, the switchlabeled FREE AIR TEMP controls the postlights in the immediate area of the outside air temperature gage. The switch is located either next to the gage on the sidewall panel or on theoverhead lighting control panel (Figure 3-9).

CABIN

Fluorescent LightsThe fluorescent cabin lights are controlled bya switch (Figure 3-10) on the copilot’s sub-panel. The switch positions are BRIGHT– DIM–OFF. (Prior to BB-1444, except 1439,this switch is labeled START/BRIGHT–

DIM–OFF. The switch must be positioned toSTART/BRIGHT until the lights illuminatebefore being moved to DIM.)

Passenger Warning SignA sign to warn passengers not to smoke and/or to fasten their seat belts (Figure 3-11) is con-trolled by a switch on the copilot’s subpanel.

The switch has three positions which are NOSMOKE & FSB–FSB–OFF. In FSB, the FAS-

Figure 3-9. Free Air Temperature Switch

TEN SEAT BELT portion of the sign illumi-Th NO SMOKING d FASTEN SEAT

Threshold and Aisle Lights





nates. The NO SMOKING and FASTEN SEAT

BELT positions are illuminated in the NOSMOKE & FSB position, with accompany-ing chimes.

Reading LightsSwitches next to each light control individualoverhead reading lights (Figure 3-12). Theselights are powered from the No. 2 dual-fed bus.

g

A light at floor level, forward of the airstair door (Figure 3-13) is designed to illuminate thethreshold. Another light, located at floor levelaft of the spar cover, illuminates the aisle. Bothlights are automatically turned on by a switchwhen the door is opened and turned off whenthe door is closed and locked if the adjacentrocker switch is placed to the ON position.

Figure 3-13. Threshold, Aisle, andBaggage Lights

Figure 3-11. Passenger Warning Sign



Baggage Area Light The following described lights are shown inFigure 3 15



A switch just inside and aft of the airstair door-frame controls a baggage area light. This switchis wired to the hot battery bus and does not au-tomatically shut off when the airstair is closed.

Passenger Oxygen SwitchWhen oxygen flows into the passenger oxy-gen system supply line, a pressure-sensitive

switch in the line closes a circuit to illuminatethe green PASS OXYGEN ON annunciator on the caution-advisory annunciator panel.On series beginning with 1979 models, thisswitch will also cause the cabin lights, thevestibule light, and the baggage compartmentlight to illuminate in the full-bright mode, re-gardless of the position of the cabin lightsswitch.

EXTERIOR LIGHTS

LANDING LIGHTS

Two sealed-beam landing lights are mountedon the nose gear (Figure 3-14). An individual

circuit-breaker switch in the lighting group onthe pilot’s right subpanel controls each light.The switches are labeled LANDING and ei-ther LEFT or RIGHT.

Figure 3-15.

WING ICE LIGHTS

The ice inspection lights are mounted on theoutside of each nacelle and illuminate thewing leading edge. A control circuit-breaker switch labeled ICE is located on the pilot’sright subpanel.

NAVIGATION LIGHTS

Navigation lights are located on each wingtipand in the horizontal stabilizer tail cone. Controlis accomplished with a circuit-breaker switchon the pilot’s right subpanel labeled NAV.

RECOGNITION LIGHTS

Lights to be used for recognition purposes areinstalled in each wingtip. These lights arecontrolled with the RECOG switch on thepilot’s right subpanel.

BEACON LIGHTS

A beacon is installed on the top of the verti-

cal stabilizer and another on the bottom of thefuselage just forward of the main gear doors.Control for these lights is incorporated into acircuit-breaker switch labeled BEACON on theright of the pilot’s right subpanel.

STROBE LIGHTS

A strobe light is installed in each wingtip and

also in the tip of the tail cone. Control for these lights is incorporated into a switch on theright of the pilot’s right subpanel and is labeledSTROBE.

Figure 3-14 Landing and Taxi Lights



NAVIGATION LIGHTSRECOGNITION LIGHTS BEACON LIGHT NAVIGATION LIGHT



AIRSTAIR FLOODLIGHT

A flush-mounted floodlight (Figure 3-16) is in-stalled forward of the flaps in the bottom of theleft wing to provide illumination of the areaaround the bottom of the airstair door. It is con-nected to the hot battery bus and is controlledby the threshold light switch and will extin-guish automatically whenever the cabin door is closed.

UNDER STEP LIGHTING

Under each step there is a light to illuminatethe airstair door (Figure 3-17). These lights arealso controlled by the threshold light switchand will extinguish automatically whenever theairstair door is closed.

Figure 3-15. Exterior Lights

WING ICE LIGHTS STROBE LIGHTS TAIL FLOODLIGHTSBEACON LIGHT



QUESTIONS



1. Where are the majority of cockpit light-ing controls?

A. Pilot’s right subpanel

B. Overhead panel

C. Copilot’s left subpanel

D. Pilot’s side panel

2. Where is the baggage-area l ight switchlocated?

A. Just inside and aft of the airstair doorframe

B. Within the baggage compartment

C. On the overhead panel

D. On the pilot’s left subpanel

3. How are the threshold and aisle l ightsturned on?

A. With a switch just aft of the door-frame

B. Automatically, when the batteryswitch is turned off

C. With a switch on the pilot’s right

subpanelD. Automatically, when the airstair door

is opened and the threshold switchturned on

4. Where is the switch for the strobe lightslocated?

A. On the overhead panel

B. On the copilot’s side panelC. On the pilot’s right subpanel

D. On the pilot’s side panel

6. What bus powers the airstair floodlight?

A. No. 1 dual-fed bus

B. Hot battery bus

C. No. 2 dual-fed bus

D. Isolation bus

7. After takeoff how are the landing lightsextinguished?

A. Automatically as the gear doors close

B. Automatically as the airplane lifts off

C. By turning off the LANDING lightswitches

D. By turning off the TAXI light switch

8. Where are the ice lights mounted?A. On the outside of the engine nacelles

B. On the wingroot

C. On the nose

D. On either side of the fuselage

9. What i s the swi tch l abeled MASTER

PANEL LIGHTS used for?A. To control power for all cockpit

panel lights

B. To shut off all cockpit lights

C. To intensify all cockpit panel lights

D. To shut off all lights

Q

CHAPTER 4






CONTENTS

Page

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

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

Dim................................................................................................................................... 4-3

Test .................................................................................................................................. 4-3

GLARESHIELD FLASHERS................................................................................................. 4-3

Master Warning Flashers.................................................................................................. 4-3

Master Caution Flashers................................................................................................... 4-3

WARNING ANNUNCIATOR PANEL (RED) ....................................................................... 4-4

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

Illumination Causes—200................................................................................................ 4-5

Illumination Causes—B200............................................................................................. 4-5

CAUTION-ADVISORY ANNUNCIATOR PANEL (AMBER/GREEN).............................. 4-7

General ............................................................................................................................. 4-7

CAUTION Switch (200 Models Only)............................................................................ 4-7

Illumination Causes—200................................................................................................ 4-9

Illumination Causes—B200........................................................................................... 4-10

QUESTIONS......................................................................................................................... 4-12

ILLUSTRATIONS





Figure Title Page

4-1 Component Locations............................................................................................... 4-2

4-2 MASTER WARNING Flasher and MASTER CAUTION Flasher ......................... 4-3

4-3 Warning Annunciator Panel—200 Aircraft.............................................................. 4-4

4-4 Warning Annunciator Panel—B200 Aircraft

(Prior to BB-1444, Except BB-1439)....................................................................... 4-4

4-5 Warning Annunciator Panel—B200 Aircraft(BB-1439, 1444 and Subsequent) ............................................................................ 4-5

4-6 Caution-Advisory Annunciator Panel—200 Aircraft (Prior to BB-453) ................. 4-7

4-7 Caution-Advisory Annunciator Panel—200 Aircraft (BB-453 and After).............. 4-8

4-8 Caution-Advisory Annunciator Panel—B200 Aircraft(Prior to BB-1444, Except 1439) ............................................................................. 4-8

4-9 Caution-Advisory Annunciator Panel—B200(BB-1439, 1444 and Subsequent) ............................................................................ 4-8

TABLES





Table Title Page

4-1 Warning Annunciator—200 Aircraft........................................................................ 4-5

4-2 Illumination Causes—B200 Aircraft (Prior to BB-1444, Except 1439).................. 4-6

4-3 Illumination Causes—B200 Aircraft (BB-1439, 1444 and Subsequent)................. 4-6

4-4 Caution Advisory Annunciator—200 Aircraft......................................................... 4-9

4-5 Caution Advisory—Prior to BB-1444, Except 1439 ............................................. 4-10

4-6 Caution Advisory—BB-1439, 1444 and Subsequent............................................. 4-11

CHAPTER 4





INTRODUCTION

The master warning system consists of a warning annunciator panel with red readouts cen-trally located in the glareshield, a caution-advisory annunciator panel with amber and greenreadouts located on the center subpanel, and two flasher lights in front of each pilot on theglareshield (one labeled MASTER WARNING (red) and the other MASTER CAUTION(amber). Adjacent to the warning annunciator panel on the glareshield is a PRESS TO TEST

switch, which is used to illuminate the annunciator lights and flashers (Figure 4-1).

GENERAL

TEST




MASTER MASTER



PRESS TO RESET

WARNINGPRESS TO RESET

CAUTION



DIM



The warning annunciators (red), caution an-nunciators (amber), advisory annunciators(green), amber MASTER CAUTION flashersand red MASTER WARNING flashers fea-ture both a bright and a dim mode of illumi-nation intensity. (Prior to BB-1444, except1439, the MASTER WARNING flasher doesnot have a dim mode.) The dim mode will beselected automatically when all the following

conditions are met: a generator is on the line;the MASTER PANEL switch is on; the OVER-HEAD FLOODLIGHTS are off; the PILOTFLIGHT LIGHTS are on; and the ambientlight level in the cockpit (as sensed by a pho-toelectric cell located in the overhead lightcontrol panel) is below a preset value. Unlessall these conditions are met, the bright modewill be selected automatically.

TEST

The lamps in the annunciator system shouldbe tested before every flight, and at any other time the integrity of a lamp is in question.Depressing the PRESS TO TEST button, lo-cated to the right of the warning annunciator panel in the glareshield, illuminates the an-

nunciator lights, both MASTER WARNINGflashers, and both MASTER CAUTION flash-ers. (The yellow NO TRANSFER lights on thefuel panel are not included in this test, sincethey do not affect flashers when a NO TRANS-FER condition exists.) Any lamp that fails toilluminate when tested should be replaced.

GLARESHIELDFLASHERS

MASTER WARNING FLASHERS

tor panel will remain on until the fault is cor-rected. However, the MASTER WARNINGflashers can be extinguished by depressingthe face of e i ther MASTER WARNINGflasher, even if the fault is not corrected. Insuch a case, the MASTER WARNING flash-ers will again be activated if an additionalwarning annunciator illuminates. When awarning fault is corrected, the affected warn-

ing annunciator will extinguish, but the MAS-TER WARNING f l asher s wi l l con t inueflashing until one of the flashers is depressedto reset the circuit.

MASTER CAUTION FLASHERS

When an annunciator-covered fault occurs

that requires the pilot’s attention, the appro-priate amber caution annunciator in the cau-tion-advisory panel illuminates, and bothMASTER CAUTION flashers begin flashing(Figure 4-2). The flashing MASTER CAU-TION lights can be extinguished by pressingthe face of either of the flashing lights to resetthe circuit. Subsequently, when any other cau-tion annunciator illuminates, the MASTER

CAUTION flashers will be activated again.Most illuminated caution annunciators on thecaution-advisory annunciator panel will re-main on until the fault condition is corrected,at which time they will extinguish. The MAS-TER CAUTION fl h ill ti fl h

PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

Figure 4-2. MASTER WARNING Flasher andMASTER CAUTION Flasher

WARNING ANNUNCIATOR





PANEL (RED)GENERAL

If a fault indicated by an illuminated warningannunciator is cleared, the annunciator will au-tomatically extinguish. Figure 4-3 shows

typical 200 airplane warning panels, and Figure4-4 shows a typical B200 airplane warningpanel. Figure 4-5 shows a typical B200, BB-1439, 1444 and subsequent warning p ane l.

PRIOR TO BB-453

BB-453 AND AFTER* OPTIONAL EQUIPMENT

Figure 4-3. Warning Annunciator Panel—200 Aircraft





ILLUMINATION CAUSES—200

Table 4-1 lists legend nomenclatures, colors,

and causes for illumination in 200 aircraft.

ILLUMINATION CAUSES—B200

Table 4-2 and 4-3 list legend nomenclatures, col-

ors, and causes for illumination in B200 aircraft.

NOMENCLATURE COLOR CAUSE FOR ILLUMINATION

FIRE L ENG Red Fire in left engine compartment

ALT WARN Red Cabin altitude exceeds 12,500 feet

FIRE R ENG Red Fire in right engine compartment

L FUEL PRESS Red Fuel pressure failure on left side

INST INV Red The inverter selected is inoperative

R FUEL PRESS Red Fuel pressure failure on right side

L BL AIR FAIL Red Melted or failed plastic left bleed air failure warning line

* A/P TRIM FAIL Red Improper trim or no trim from autopilot trim command

R BL AIR FAIL Red Melted or failed plastic right bleed air failure warning line

L CHIP DETECT Red Contamination is detected in left engine oil

* A/P DISC Red Autopilot is disconnected

R CHIP DETECT Red Contamination is detected in right engine oil

Table 4-1. WARNING ANNUNCIATOR—200 AIRCRAFT

*OPTIONAL EQUIPMENT

Figure 4-5. Warning Annunciator Panel—B200 Aircraft(BB-1439, 1444 and Subsequent)



Table 4-2. ILLUMINATION CAUSES—B200 AIRCRAFT (PRIOR TO BB-1444, EXCEPT 1439)




L ENG FIRE Red Fire in left engine compartment

INVERTER Red The inverter selected is inoperative

DOOR UNLOCKED Red Cabin/cargo door open or not secure


R ENG FIRE Red Fire in right engine compartment

L FUEL PRESS Red Fuel pressure failure on left sideR FUEL PRESS Red Fuel pressure failure on right side

* L OIL PRESS Red Low oil pressure left engine

* L GEN OVHT Red Left generator temperature too high

Table 4-3. ILLUMINATION CAUSES—B200 AIRCRAFT (BB-1439, 1444 AND SUBSEQUENT)

NOMENCLATURE COLOR CAUSE FOR ILLUMINATIONL ENG FIRE Red Fire in left engine compartment

INVERTER Red The inverter selected is inoperative

CABIN/DOOR Red Cabin/cargo door open or not secure


R ENG FIRE Red Fire in right engine compartment

L FUEL PRESS Red Fuel pressure failure on left side

R FUEL PRESS Red Fuel pressure failure on right side

* L OIL PRESS Red Low oil pressure left engine

* L GEN OVHT Red Left generator temperature too high

* A/P TRIM FAIL Red Improper trim or no trim from autopilot trim command

* R GEN OVHT Red Right generator temperature too high

* R OIL PRESS Red Low oil pressure right engine

L CHIP DETECT Red Contamination is detected in left engine oil

L BL AIR FAIL Red Melted or failed plastic left bleed air failure warning line

* A/P FAIL Red Autopilot is disconnected

R BL AIR FAIL Red Melted or failed plastic right bleed air failure warning line

R CHIP DETECT Red Contamination is detected in right engine oil

* Optional equipment

CAUTION-ADVISORY flashing, the pilot can still extinguish the an-nunciator by momentarily moving the spring-





ANNUNCIATOR PANEL(AMBER/GREEN)

GENERAL

If a cautionary fault exists, the appropriateamber light will illuminate. If the fault indicatedby an illuminated caution annunciator is cleared,

the annunciator will automatically extinguish.

The caution-advisory annunciator panel alsocontains green advisory annunciators. There areno master flashers associated with these an-nunciators since they are only advisory in na-ture, indicating functional situations which donot demand the immediate attention or reac-tion of the pilot. An advisory annunciator can

be extinguished only by correcting the condi-tion indicated on the illuminated lens.

CAUTION SWITCH(200 MODELS ONLY)

If the fault indicated by an illuminated cau-tion annunciator is not corrected, and pro-vided the MASTER CAUTION flasher is not

loaded CAUTION toggle switch (if installed)down to the OFF position, then releasing itto the center position. This action will ex-tinguish all illuminated caution annuncia-tors, and will illuminate the green CAUTLGND OFF advisory annunciator in the cau-tion advisory panel; this reminds the pilotthat an uncorrected fault condition exists,but that the caution legends have all been

extinguished. The annunciator(s) previouslyextinguished with the CAUTION switch canagain be illuminated anytime by momentar-ily moving the switch up to the ON position.This action will also extinguish the greenCAUT LGND OFF annunciator. If an addi-tional fault covered by the caution annunci-ators occurs after the caution legends havebeen extinguished with the CAUTION switch,

the appropriate caution annunciator for thenew fault will illuminate, and all previouslyextinguished annunciators will again illu-minate. This switch is not installed in B200airplanes.

Figures 4-6, 4-7, 4-8 and 4-9 show typicalcaut ion advisory annunciator panels in200/B200 aircraft.





Figure 4-7. Caution-Advisory Annunciator Panel—200 Aircraft (BB-453 and After)

Figure 4-8. Caution-Advisory Annunciator Panel—B200 Aircraft(Prior to BB-1444, Except 1439)

ILLUMINATION CAUSES—200

Table 4-4 is a listing of the warning legendl t l d f ill i

nation (starting on the top left and moving toth i ht) f th 200 i ft





nomenclatures, colors, and causes for illumi- the right) for the 200 aircraft.

Table 4-4. CAUTION ADVISORY ANNUNCIATOR—200 AIRCRAFT


L DC GEN Amber Left generator off line

L ICE VANE Amber Left ice vane malfunction. Ice vane has not attained

proper positionRVS NOT READY Amber Propeller levers are not in the high-rpm, low-pitch position

with landing gear extended

R ICE VANE Amber Right ice vane malfunction. Ice vane has not attainedproper position

R DC GEN Amber Right generator off line

CABIN DOOR Amber Cabin door open or not secure

PROP SYNC ON Amber Synchrophaser is turned on with the landing gear extended

EXT PWR Amber External power connector is plugged inBATTERY CHG Amber Excessive charge rate on the battery

DUCT OVERTEMP Amber Duct air too hot

* L AUTOFEATHER Green Autofeather armed with power levers advanced aboveapproximately 90% N1 power lever position

* ELEC TRIM OFF Green Electric trim deengergized by a trim disconnect switch onthe control wheel with the system power switch on thepedestal turned on

FUEL CROSSFEED Green Crossfeed has been selected

AIR COND N1 LOW Green Right engine rpm is too low for air-conditioning load

* R AUTOFEATHER Green Autofeather armed with power levers advancedabove approximately 90% N1 power lever position

L ICE VANE EXT Green Ice vane extended

* BRAKE DEICE ON Green Brake deice has been selected

LANDING LIGHT Green Landing lights on with landing gear up

PASS OXYGEN ON Green Oxygen is available to the passengers

R ICE VANE EXT Green Ice vane extendedL IGNITION ON Green Left starter/ignition switch is in the engine/ignition mode or

left autoignition system is armed and left engine torque isbelow 400 ft-lbs

L BL AIR OFF Green Left environmental bleed-air valve is closed

ILLUMINATION CAUSES—B200

Tables 4-5 and 4-6 list the warning legendsnomenclatures colors and causes for illumi

nation (starting on the top left and moving tothe right) for the B 200 aircraft





nomenclatures, colors, and causes for illumi- the right) for the B-200 aircraft.

Table 4-5. CAUTION ADVISORY—PRIOR TO BB-1444, EXCEPT 1439



HYD FLUID LOW Amber Hydraulic fluid in the landing gear system is low

†*PROP SYNC ON Amber Synchrophaser is turned on with the landing gear extendedRVS NOT READY Amber Propeller levers are not in the high-rpm, low-pitch position

with landing gear extended


DUCT OVERTEMP Amber Duct air too hot

L ICE VANE Amber Left ice vane malfunction. Ice vane has not attainedproper position

BATTERY CHG Amber Excessive charge rate on the battery

EXT PWR Amber External power connector is plugged inR ICE VANE Amber Right ice vane malfunction. Ice vane has not attained

proper position

*L AUTOFEATHER Green Autofeather armed with power levers advancedabove approximately 90% N1 power lever position

*ELEC TRIM OFF Green Electric trim deengergized by a trim disconnect switch onthe control wheel with the system power switch on thepedestal turned on


*R AUTOFEATHER Green Autofeather armed with power levers advancedabove approximately 90% N1 power lever position

L ICE VANE EXT Green Ice vane extended

*BRAKE DEICE ON Green Brake deice has been selected

LDG/TAXI LIGHT Green Landing lights on with landing gear up

PASS OXY ON Green Oxygen is available to the passengers

R ICE VANE EXT Green Ice vane extended

L IGNITION ON Green Left starter/ignition switch is in the engine/ignition mode

or left autoignition system is armed and left engine torqueis below 400 ft-lbs





Table 4-6. CAUTION ADVISORY—BB-1439, 1444 AND SUBSEQUENT






HYD FLUID LOW Amber Hydraulic fluid in the landing gear system is low

RVS NOT READY Amber Propeller levers are not in the high-rpm, low-pitch positionwith landing gear extended


L CHIP DETECT Amber Metal contamination in the left engine oil is detected

DUCT OVERTEMP Amber Duct air too hotR CHIP DETECT Amber Metal contamination in the right engine oil is detected

L ENG ICE FAIL Amber Left engine anti-ice malfunction. Ice vane has not attainedproper position

BATTERY CHG Amber Excessive charge rate on the battery

EXT PWR Amber External power connector is plugged in

R ENG ICE FAIL Amber Right engine anti-ice malfunction. Ice vane has notattained proper position

* L AUTOFEATHER Green Autofeather armed with power levers advanced aboveapproximately 90% N1 power lever position

* ELEC TRIM OFF Green Electric trim deengergized by a trim disconnect switch onthe control wheel with the system power switch on thepedestal turned on


* R AUTOFEATHER Green Autofeather armed with power levers advanced aboveapproximately 90% N1 power lever position

L ENG ANTI-ICE Green Left engine anti-ice vane extended

* BRAKE DEICE ON Green Brake deice has been selected

LDG/TAXI LIGHT Green Landing lights on with landing gear up

PASS OXY ON Green Oxygen is available to the passengers

ELEC HEAT ON Green Cabin electric heat is on

R ENG ANTI-ICE Green Right engine anti-ice vane extended

L IGNITION ON Green Left starter/ignition switch is in the engine/ignition mode orleft autoignition system is armed and left engine torqueis below 400 ft-lbs



R BL AIR OFF Green Right environmental bleed-air valve is closed

R IGNITION ON Green Right starter/ignition switch is in the engine/ignition mode



QUESTIONS



1. How are the MASTER CAUTION flash-ers dimmed?

A. By using the BRT DIM switch

B. With the overhead control rheostats

C. Automatically relative to cockpit lightintensity

D. With the CAUTION switch on the

copilot’s subpanel

2. How can the annunciator lights be tested?

A. By depressing each light legend

B. By moving the CAUTION switch toON

C. With the APPROACH PLATE rheo-stat

D. With the PRESS TO TEST switch

3. To ext inguish a MASTER WARNINGflasher, what action must be taken?

A. Move the CAUTION switch to OFF

B. Depress either MASTER WARNINGflasher

C. Depress the PRESS TO TEST buttonD. Clear the illuminating fault

4. W he n wi ll a r ed a nn un ci at or l ig htextinguish?

A. When the indicated fault is cleared

B. W h e n th e M A S TE R WA R N IN Gflasher is pressed

C. When the RESET button is depressed

D. When the TEST button is depressed

5. To illuminate the CAUT LGND OFF lighton 200 models, what act ion must betaken?

A. Move the CAUTION switch to ON

B. Depress the PRESS TO TEST button

C. Move the CAUTION switch to OFF

D. Depress a MASTER CAUTION flasher

CHAPTER 5

FUEL SYSTEM





FUEL SYSTEM

CONTENTS

Page

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

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

Fuel Routing into the Engine .......................................................................................... 5-2

MAJOR COMPONENT LOCATIONS AND FUNCTIONS ................................................. 5-2

Main and Auxiliary Fuel Systems ................................................................................... 5-2

Auxiliary Fuel Transfer System ...................................................................................... 5-5

Firewall Shutoff Valve .................................................................................................... 5-6

Engine-Driven Boost Pump ............................................................................................ 5-8

Standby Boost Pump ....................................................................................................... 5-9

Firewall Fuel Filter .......................................................................................................... 5-9

Low Fuel Pressure Switch................................................................................................ 5-9

Fuel Flow Transmitter and Gages ................................................................................... 5-9

Fuel Heater..................................................................................................................... 5-10

High-Pressure Engine Fuel Pump.................................................................................. 5-10

FUEL MANIFOLD CLEARING ......................................................................................... 5-10

Fuel Purge System ........................................................................................................ 5-10

Fuel Drain Collector System ......................................................................................... 5-11

Vent System .................................................................................................................. 5-13

Fuel Drains .................................................................................................................... 5-14





.................................................................................................................... 5 14LIMITATIONS ..................................................................................................................... 5-14

Approved Fuel Grades and Operating Limitations ....................................................... 5-14

Approved Fuel Additive ............................................................................................... 5-15

Fueling Considerations ................................................................................................. 5-16

Zero-Fuel Weight .......................................................................................................... 5-16

QUESTIONS......................................................................................................................... 5-17

ILLUSTRATIONS

Figure Title Page





g g5-1 Fuel System Schematic for the Super King Air 200 and B200

(BB-666 and Subsequent) ........................................................................................ 5-3

5-2 Fuel System Schematic for the Super King Air 200 (Prior to BB-666)................... 5-4

5-3 Fuel Tank/Cell Capacities (Super King Air 200 and B200)..................................... 5-5

5-4 Fuel Pressure Warning Lights .................................................................................. 5-65-5 Fuel Control Panel.................................................................................................... 5-7

5-6 Auxiliary Fuel Transfer System ............................................................................... 5-8

5-7 Fuel Flow Gages .................................................................................................... 5-10

5-8 Fuel Purge System.................................................................................................. 5-11

5-9 Fuel Crossfeed System........................................................................................... 5-12

5-10 Fuel Crossfeed Advisory Light .............................................................................. 5-13

5-11 Fuel Temperature (OAT) Versus Minimum Oil Temperature Graph ..................... 5-15

TABLES

Table Title Page

5-1 Drain Locations...................................................................................................... 5-14

CHAPTER 5

FUEL SYSTEM





INTRODUCTION

The airplane fuel system consists of two separate wing fuel systems connected with acommon crossfeed line and solenoid-operated crossfeed valve. Each wing system is fur-ther divided into a main and an auxiliary system. The main system employs a total of 386 gallons of usable fuel; the auxiliary system, 158 gallons . At 6.7 pounds per gallon,these totals convert to 2,586 pounds in the main system and 1,058 pounds in the auxil-iary system. Total usable fuel is 544 gallons, or 3,644 pounds.

GENERAL

h i f l i f l d h h

0

2

4 6

8

10

MAIN

FUEL

LBS X 100

FUEL SYSTEM

Additionally, the fuel system incorporates afully automatic vent system; a capacitancefuel gaging system on each side which providesseparate quantity readings for each main and

ili f l d f l fil

FUEL ROUTING INTO THEENGINE

After exiting the main fuel system, fuel passes





auxiliary fuel system; and a fuel filter systemincorporating a filter bypass to enable fuelfeed to the engine in the event of filter icingor clogging.

A high-pressure fuel pump and a low-pressureboost pump are engine-driven through the ac-cessory drive section. The high-pressure fuelpump delivers fuel to the engine. The engine-driven boost pump delivers low-pressure fuelto the high-pressure fuel pump to prevent cav-itation and ensure continuous flow of fuel. Inthe event that the engine-driven boost pumpfails, the electric standby boost pump shouldbe actuated. The low-pressure standby boostpump is electrically powered and is submergedin the bottom of the nacelle tank.

On SN BB-666 and subsequent, a pneumaticpressure fuel purge system delivers fuel re-maining in the engine fuel nozzle manifoldsat engine shutdown to the combustion cham-ber for burning. On airplanes prior to BB-666,excess fuel remaining in the engine fuel noz-zle manifolds at engine shutdown is returned

to the gravity-feed line from the fuel draincollector system.

A fuel crossfeed system is available for (andlimited to) single-engine operation to cross-feed from the main fuel system. However, if needed, all published usable fuel in either wing system is available for crossfeed to ei-ther engine.

Approved fuel grades, operating limitationsand fueling considerations are covered in theLIMITATIONS section of this chapter.

through the normally open firewall shutoff valve. Just downstream of this valve is thelow-pressure, engine-driven boost pump. Fromthis pump, fuel is subsequently routed to thefirewall fuel filter and pressure switch, througha fuel heater which utilizes heat from engineoil, to the engine fuel pump, on to the fuel con-

trol unit (FCU), and then through the fuel flowtransmitter (prior to BB-1401, the fuel flowtransmitter is upstream of the fuel heater).Fuel is then directed through the dual fuelmanifold to the fuel sprayer nozzles and intothe annular combustion chamber. Fuel is alsotaken from just downstream of the firewallfuel filter to supply the auxiliary tank trans-fer system with motive fuel flow.

MAJOR COMPONENTLOCATIONS ANDFUNCTIONS

MAIN AND AUXILIARY FUEL

SYSTEMSEach fuel system is divided into a main and anauxiliary fuel system, with a total usable fuelcapacity of 544 gallons. See Figure 5-1 for Models 200 and B200, SN BB-666 and sub-sequent. See Figure 5-2 for Model 200 prior to SN BB-666.

The total usable fuel capacity of the main fuelsystem is 386 gallons (Figure 5-3).

The filler cap for the main fuel system is lo-cated on top of the leading edge of the wing,

ENGINE FUEL MANIFOLDLEGEND



5 -

3

F OR T RAI NI N G P UR

P O S E S

ONL Y

S

UP E R

KI N G

A I R

2 0 0

/ B 2 0 0

P I L O T T R A I NI N

G MA N U A L

F l i gh t S af e t y


L

F AUXILIARY

WING LEADING EDGE

WING LEADING EDGE

INTEGRAL (WET WING) BOXSECTION

BOXSECTION

DRAINVALVE

VENTFLOAT

VALVE

AIR INLET

DRAIN

RECESSED VENT

HEATED RAM VENT

FLAME ARRESTER

*BB 1401 & SUBSEQUENT LOCATED DOWNSTREAM OF FUEL CONTROL UNIT**BB 1193 & SUBSEQUENT DRAIN RELOCATED TO OUTBOARD SIDE OF NACELLE

DRAIN**

TRANSFER JET PUMP

STRAINER, DRAIN,AND FUEL SWITCH

F

FUEL CONTROL UNIT

ENGINE FUEL MANIFOLD

PRESSURE TANK

FIREWALL FUEL FILTER

ENGINE-DRIVEN BOOST PUMP (LP)

DRAIN VALVE (FIREWALL)

FIREWALL SHUTOFF VALVE

STANDBY BOOST PUMP (30 PSI)

NACELLE TANK

VENT FLOAT VALVE

CROSSFEED VALVE (NC)

P3 BLEED AIR LINE

ENGINE FUEL PUMP (HP)

FUEL HEATER

AIR FILTER

*FUEL FLOW TRANSMITTER AND INDICATORLEFT FUEL PRESSURE ANNUNCIATOR PRESSURE SWITCH

FUEL CONTROL PURGE VALVE

GRAVITY FLOW CHECK VALVE

STRAINER AND DEFUELING DRAIN VALVE

TRANSFER CONTROL MOTIVE FLOW VALVE (NC)

PRESSURE SWITCH FOR LEFT NO FUELTRANSFER LIGHT ON FUEL PANEL (6 PSI)

FUEL

FUEL AT STRAINER OR FILTER

FUEL UNDER LOW PUMP PRESSURE

HIGH-PRESSURE FUEL

FUEL CROSSFEED

FUEL RETUREN

GRAVITY FEED

FUEL VENT

FILLER

PROBES

SUCTION RELIEF VALVE

CHECK VALVE

FUEL FLOW TRANSMITTER

FUEL PRESSURE ANNUNCIATOR

F

L

Figure 5-1. Fuel System Schematic for the Super King Air 200 and B200 (BB-666 and Subsequent)

5 - 4

FROM FUEL NOZZLE MANIFOLD

FUEL DRAIN COLLECTOR TANK

FLOAT SWITCHFUEL DRAIN RETURN PUMPENGINE FUEL PUMP

FUEL CONTROL UNIT

LEGEND

FUELFUEL AT STRAINER OR FILTER



F OR T

RAI NI N G P URP O S E S

ONL Y

S UP E R

KI N G

A I R

2 0 0 / B 2 0 0

P I L O T T R A I NI N G MA N

U A L



F

F

L

AUXILIARY

WING LEADING EDGE WING LEADING EDGE

INTEGRAL (WET WING)BOX SECTION

DRAIN VALVE

VENT FLOAT VALVE

AIR INLET

DRAIN

BOX SECTION

RECESSED VENT

HEATED RAM VENT

FLAME ARRESTER

FUEL HEATER

PRESSURE SWITCH (10 PSI)


FLAME ARRESTER

ENGINE-DRIVEN

BOOST PUMP (LP, 30 PSI)


DRAIN VALVE FIREWALL

STANDBY BOOST PUMP (30 PSI)

NACELLE TANK

VENT FLOAT VALVE

CROSSFEED VALVE (NC)

(HP)


STRAINER AND DEFUELING DRAIN VALVE

FUEL CONTROL

PURGE VALVE

TRANSFER CONTROL

MOTIVE FLOW VALVE (NC)

GRAVITY FLOW CHECK VALVEPRESSURE SWITCH

(6 PSI)

DRAIN

TRANSFER JET PUMP

AIR INLET (PRIOR TO SI 1021)

STRAINER, DRAIN,

AND FUEL SWITCH

FUEL UNDER LOW PUMP PRESSURE

HIGH-PRESSURE FUEL

FUEL CROSSFEED

FUEL RETUREN

GRAVITY FEED

FUEL VENT

FILLER

PROBESSUCTION RELIEF VALVE

CHECK VALVE


FUEL PRESSURE ANNUNCIATOR

F

L

Figure 5-2. Fuel System Schematic for the Super King Air 200 (Prior to BB-666)



NACELLETANK

(57GALLONS)



Each auxiliary fuel system is equipped withits own filler port and antisiphon valve.

While the auxiliary fuel system is being used,fuel is transferred from the auxiliary tank tothe nacelle tank by a jet transfer pump, whichis mounted adjacent to the outlet strainer anddrain in the auxiliary fuel cell.

A swing check valve in the gravity feed lineprevents reverse flow into the outboard tankswhen the auxiliary transfer system is in use.When auxiliary fuel is exhausted, normal grav-ity flow from the outboard tanks to the nacelle

tanks begins.

AUXILIARY FUEL TRANSFERSYSTEM

When auxiliary fuel is available, this systemautomatically transfers fuel from the auxiliarytank to the nacelle tank. No pilot action is in-volved. The jet transfer pump in the auxiliary

tank operates on the venturi principle using thefuel and boost pump for motive flow. The en-gine-driven or electric low-pressure boostpump routes fuel through the normally-closedmotive flow valve the jet pump and into the

for engine starting. At the end of this time, themotive flow valve opens automatically andfuel transfer begins. The pilot should monitor the NO TRANSFER lights on the fuel panelto ensure that they are extinguished 30 to 50seconds after engine start. The pilot should alsomonitor the auxiliary fuel level during the be-ginning of the flight to ensure that the trans-fer of fuel is taking place.

Fuel pressure supplied by either the engine-driven boost pump or the electric standby boostpump (normally 25 to 30 psi) will open a fuelpressure-sensing switch and extinguish the

red FUEL PRESS warning light (Figure 5-4).A minimum pressure of 10 ± 1 psi is requiredto extinguish the light. This same FUEL PRESSswitch will also send a signal to the auxiliaryfuel transfer printed circuit board indicatingthat motive flow is available for fuel transfer.If there is fuel in the auxiliary tank, this cir-cuit board will open the motive flow valvewithin 30 to 50 seconds. With the motive flow

valve now open, fuel is permitted to flowthrough the auxiliary transfer line. If the fuelpressure in this auxiliary transfer line is atleast 4 to 6 psi, a normally-closed pressureswitch will open and extinguish the amber NO

WING LEADING EDGE(13 GALLONS)

INTEGRAL WET WING(35 GALLONS)

WING LEADING EDGE(40 GALLONS)

BOX SECTION(25 GALLONS)

BOX SECTION(23 GALLONS)

AUXILIARY(79 GALLONS)

Figure 5-3. Fuel Tank/Cell Capacities (Super King Air 200 and B200)





FER light because there is no more fuel left totransfer.

If the motive flow valve or its associated cir-cuitry should fail, it will go to the normally-

closed position. Loss of motive flow pressurewith fuel remaining in the auxiliary tank willilluminate the amber NO TRANSFER lighton the applicable side of the fuel control panel.The motive flow valve may be manually en-ergized to the open position by placing theAUXILIARY TRANSFER switch, normally inthe AUTO position, to the OVERRIDE posi-tion (Figure 5-5). This procedure will bypass

the automatic feature in the auxiliary transfer system and send DC power directly to the mo-tive flow valve.

On BB 32 and subsequent airplanes and on ear

On SN BB-2 through BB-31, selecting theOVERRIDE position of the switch takes power from the NO TRANSFER light, causing it toextinguish. Even though the light is extin-guished, the valve may or may not open. The

auxiliary fuel level must be monitored to en-sure that it is decreasing.

The amber NO TRANSFER lights installed onairplanes prior to SN BB-516 illuminate andstay bright. On SN BB-516 and subsequent,they are dimmed through the airplane’s auto-matic dimming system.


The fuel system incorporates two in-line motor-driven firewall shutoff valves, one on each side.E h i t ll d b di ( d d)

Figure 5-4. Fuel Pressure Warning Lights





AUXTRANSFER

CB

AUX TRANSFERSWITCH

OVER RIDE NOT EMPTY

FLOAT SWITCHMOTIVE FLOW

PRESSURE SWITCH



The firewall shutoff valves, like the standbyboost pumps, are powered by the No. 3 (left) andNo. 4 (right) dual-fed buses. The firewall shut-off valves are also powered from the hot batterybus. Therefore, they can be operated regardlessof battery-switch position. When these valvesare closed, fuel is cut off from the engine.

ENGINE-DRIVEN BOOST PUMP

The low-pressure, engine-driven boost pumpis mounted on a drive pad on the aft accessorysection of the engine. The boost pump deliv-

ers low-pressure fuel to the engine high-pres-sure fuel pump, thus preventing cavitation.The boost pump is protected against contam-ination by a strainer, and has an operating ca-pacity of 1,250 pph at a pressure of 25 to 30

In case of a low-pressure engine-driven boostpump failure, the L or R red FUEL PRESS lightilluminates on the warning annunciator panel(Figure 5-4). The light illuminates when pres-sure decreases below 10 ± 1 psi. Activation of the standby boost pump on the side of the fail-ure will increase the pressure and extinguishthe light.

Engine operation with the fuel pres-

sure light on is limited to 10 hours be-fore overhaul or replacement of thehigh-pressure main engine fuel pump.

CAUTION

L FUEL PRESS LOL FUEL PRESS LO

AUXTRANSFER

PCB

FROMCROSSFEED

SWITCH

MOTIVE FLOWVALVE

BOOST PUMPPRESSURESWITCH

FROM ENGINE

DRIVEN BOOSTPUMP

PRESSUREWARNING

JET TRANSFERPUMP

FROMAUX TANK

TONACELLE TANK

CB

AUTOEMPTY

TO

ENGINE

N.C.

Figure 5-6. Auxiliary Fuel Transfer System

STANDBY BOOST PUMP

An electrically-driven, low-pressure standby

boost pump located in the bottom of each na-celle tank performs three functions:

LOW FUEL PRESSURESWITCH

Mounted on top of the firewall fuel filter is afuel pressure-sensing switch. In the event of





1. It is a backup pump for use in the event of an engine-driven fuel boost pump failure.

2. It is used with aviation gas above 20,000feet.

3. It is used during crossfeed operation.

If a standby pump becomes inoperative, cross-feed can be accomplished only from the sideof the operative standby pump.

Electrical power for standby pump operation iscontrolled by lever-lock switches on the fuelcontrol panel (Figure 5-5) and DC power is

supplied from the dual fed buses (prior to BB-1098, except 1096 the standby boost pumps arealso powered by the hot battery bus). Theswi tches are labeled STANDBY PUMPON–OFF. With the master switch on, power issupplied from the No. 3 (left) or No. 4 (right)bus feeders through the STANDBY PUMP cir-cuit breakers on the fuel control panel to thepumps.

Prior to BB-1098, except 1096, battery power from the hot battery bus is also available for standby boost pump operation. Fuses locatedin the right wing center section adjacent to thebattery box protect this circuit. These circuitsuse diodes to prevent failure of one circuitfrom disabling the other circuit. During shut-down, both STANDBY PUMP switches and

the CROSSFEED FLOW switch must be po-sitioned to OFF to prevent battery discharge.


p gan engine-driven boost pump failure or anyother failure resulting in low pressure in thefuel line, the respective fuel pressure switchwill close, causing the red FUEL PRESS lighton the warning annunciator panel to illuminate(Figure 5-4). This light illuminates any time

pressure decreases below 10 ± 1 psi. The lightwill normally be extinguished by switching onthe standby boost pump on that side.

This switch also sends a signal to the auxil-iary fuel transfer printed circuit board advis-ing the system if fuel pressure is or is notavailable for auxiliary tank transfer.

FUEL FLOW TRANSMITTERAND GAGES

The fuel flow gages on the instrument panelindicate fuel flow in pounds per hour (Figure5-7). The following list indicates how thesegages are powered:

• Prior to BB-225 by 26-volt AC power

• BB-225 through 1483, including 1485,by DC power from the No. 1 and No. 2dual-fed buses.

• BB-1484, 1486 and subsequent by DCpower from the No. 1 and No. 2 dual-fedbuses or from the isolation bus.

With BB-1401 and subsequent, the transmit-ters were moved downstream of the fuel con-t ro l un i t t o on ly ind ica te fue l used fo r combustion. Prior to BB-1401, the fuel flowtransmitters were installed in the fuel line up-

FUEL MANIFOLDCLEARING

FUEL PURGE SYSTEM





FUEL HEATERFuel is heated prior to entering the fuel controlunit by an oil-to-fuel heat exchanger. An engineoil line is in proximity with the fuel line and,through conduction, a heat transfer occurs. Thepurpose of heating the fuel is to remove any iceformation which may have occurred or pre-clude any ice from forming, and which may re-

sult in fuel blockage at the fuel control unit (seeLIMITATIONS at the end of this chapter). Thefuel heater is thermostatically controlled tomaintain a fuel temperature of 70° to 90°F under normal conditions. If the fuel temperature risesabove 90°, the fuel will automatically bypass thefuel heater. If the fuel is extremely cold, and theoil temperature is too low, the unit may not becapable of preventing icing in the FCU. The oil

vs. fuel temperature graph in the LIMITATIONSsection will specify under what conditions icingmay occur. The fuel heater is automatic and re-quires no pilot action.

HIGH-PRESSURE ENGINEFUEL PUMP

The high-pressure engine fuel pump is enginedriven and is mounted on the accessory drivein conjunction with the fuel control unit. Thisgear-type pump supplies the fuel pressureneeded for a proper spray pattern in the com-

FUEL PURGE SYSTEM

(BB-666 and Subsequent)The fuel purge system (Figure 5-8) uses P3bleed air to purge the fuel manifolds of fuelwhen the condition lever is placed in the fuelcutoff position and the fuel pressure in the fuelmanifold decreases.

Fuel enters the fuel manifolds in the normalmanner via the flow divider. Incorporated inthe flow divider is the dump valve which func-tions to prevent fuel from the fuel controlfrom entering the purge line while the engineis in operation. P3 air is extracted from the en-gine compressor and sent to the airframe ser-vices (pressurization/pneumatics) just aft of the fireseal. At the point where the airframeservices distribution is separated, a small lineis tapped off and P3 air is sent via a filter andcheck valve to the purge tank. The output endof the purge tank also has a check valve, work-ing in conjunction with the dump valve, whichprevents the return of fuel or air from the fuelmanifolds to the purge tank.

In normal operation, the P3 air generated by theengine is held within the purge tank by theinput check valve and fuel pressure whichholds the dump valve shuttle closed. When theengine is shut down, fuel pressure on the dumpvalve shuttle decreases. The shuttle valve openswhen P3 pressure is greater than fuel manifoldpressure. This allows P

3air to enter the fuel

manifolds, forcing the remaining fuel in themanifolds into the burner can. Since combus-tion has not ceased, this small amount of fuelfrom the manifolds is now burned, which may

Figure 5-7. Fuel Flow Gages



FUEL

PUMP

ENGINEFUEL

CONTROLUNIT

FUELHEAT

FUELFLOW *



FUEL DRAIN COLLECTORSYSTEM

(BB-2 through BB-665)Fuel from the engine flow divider drains intoa collector tank mounted below the aft engineaccessory section. An internal float switchactuates an electric pump, which delivers thefuel to the fuel purge line just aft of the fuelpurge shutoff valve. A check valve in the lineprevents the backflow of fuel during engine

purging. A vent line plumbed from the top of the collector tank is routed through an in-lineflame arrester and downward to a drain mani-fold on the underside of the nacelle (Figure 5-2) The system receives power from the No 1

essary to supply fuel to the operative enginefrom the fuel system on the opposite side(Figure 5-9). The simplified crossfeed con-

t ro l posi t ions are labeled CROSSFEEDFLOW and OFF (Figure 5-5). The STANDBYPUMP switches must be positioned to OFFfor crossfeeding.

The auxiliary transfer switch must be

positioned to the AUTO position onthe side being crossfed. If auxiliaryfuel supply is required from the in-operative engine side, the firewallvalve must be opened provided en-

CAUTION

BOOST PUMP PRESSURE

HIGH-PRESSURE FUEL

ENGINE BLEED AIR

LEGEND

FILTER

CHECKVALVE

PURGE TANK

CHECKVALVE

PRIOR TO BB-1401, FUEL FLOWTRANSMITTER IS UPSTREAM OF

FUEL HEATER

TOPNEUMATICS

TOFLOW

PACKAGE

PURGELINE

DUMPVALVE

POPPETVALVE

FROMP3 AIR

FIRESEAL

*

Figure 5-8. Fuel Purge System



MOTIVE FLOWVALVE

FIREWALLSHUTOFF VALVE

LOW PRESSUREENGINE-DRIVEN

FUEL PUMP



1. The green FUEL CROSSFEED annun-ciator will illuminate (Figure 5-10).

2. The CROSSFEED valve will open.

3. The standby boost pump on the deliveryside will be turned on.

4. The motive flow valve on the receivingside will close, stopping auxiliary tankfuel transfer.

If there is fuel in the receiving side’s auxiliarytank when crossfeed is selected:

5. The NO TRANSFER light on the re-ceiving side will illuminate. (Note thatthis will not occur if there is no auxil

Illumination of the green FUEL CROSSFEEDlight on the caution/advisory panel indicatescrossfeed has been selected, not that the cross-

feed valve has moved. The Before Engi neStarting checklist contains a crossfeed test toensure operation of this valve. During this test,the pilot should ensure that both red FUELPRESSure lights extinguish once the CROSS-FEED switch is moved LEFT or RIGHT, in-dicating the valve has opened.

FUEL GAGING SYSTEMA capacitance-type fuel gaging system mon-itors fuel quantity in either the main or aux-iliary fuel system for each side. Two fuel

MOTIVE FLOWVALVE

CROSSFEEDVALVE

STANDBY BOOST PUMP

FIREWALLSHUTOFF

VALVE

Figure 5-9. Fuel Crossfeed System





Quantity is read directly in pounds. An error of 3% maximum may be encountered in the sys-tem. The readings are compensated for density

changes caused by temperature variations.

A FUEL QUANTITY selector switch on thefuel control panel placarded MAIN and AUX-ILIARY allows monitoring of the main or aux-iliary system fuel quantity. This switch isspring loaded to the main system and must beheld in the auxiliary position for reading.

The fully-independent indicating system oneach side of the airplane incorporates eightprobes: one in the inboard box fuel cell, onein the nacelle fuel cell, two in the integralwet-wing cell, two in the inboard leading edgecell, and two in the auxiliary tank.

Power is supplied through the capacitanceprobes to the quantity indicator.

FUELING

Fuels and fueling considerations are covered

is covered in the NORMAL PROCEDURESsection of the Fl ight Manual .

ANTISIPHON VALVE

An antisiphon valve installed at each filler port prevents loss of fuel in the event of im-proper securing or loss of the filler cap inflight.

VENT SYSTEM

The two wing fuel systems are vented throughrecessed ram vents coupled to protrudingheated ram vents on the underside of thewing adjacent to the nacelle. One vent oneach side is recessed and aerodynamicallyprevents ice from forming. The other vent isprotruding and is heated to prevent icing.Refer to Chapter 10, ICE AND RAIN PRO-

TECTION, for additional information.

An air inlet at the wingtip vents the integral(wet cell) tank, the auxiliary tank and for BB-479 d b h ll k P i

Figure 5-10. Fuel Crossfeed Advisory Light

FUEL DRAINS

There are five sump drains and a firewall fil-

ter drain in each wing. Drain locations areshown in Table 5-1.

each engine divided by the engine fuel con-sumption rate equals the number of hours to becharged against time between overhauls (TBO).

The pilot must be familiar with the consump-tion rate of his airplane and record the num-





LIMITATIONS

APPROVED FUEL GRADESAND OPERATING LIMITATIONS

Commercial Grades Jet A, Jet A-1, and Jet B,and Military Grades JP-4 and JP-5 are recom-mended fuels for use in the Super King Air 200and B200. They may be mixed in any ratio.

Aviation gasoline Grades 80 Red (formerly80/87), 91/98, 100LL Blue (same as 100LGreen in some countries), 100 Green (formerly

100/130), and 115/145 Purple are emergencyfuels. Emergency fuels may be mixed withrecommended fuels in any ratio. However,when aviation gasoline is used, operation islimited to 150 hours between engine over-hauls. The number of gallons taken aboard for

tion rate of his airplane and record the number of gallons taken aboard for each engine.

It is recommended that the pilot refer also tothe Limitations chart in the POH concerningstandby boost pumps and crossfeed opera-tions when aviation gasoline is used.

Takeoff is prohibited if either fuel quantitygage indicates less than 265 pounds of fuel or is in the yellow arc.

Crossfeed is utilized for single-engine oper-ation only.

Operation of either engine with its corre-

sponding fuel pressure warning annunciator (LFUEL PRESS or R FUEL PRESS) illuminatedis limited to 10 hours between overhaul or re-placement of the high-pressure main enginefuel pump.

Table 5-1. DRAIN LOCATIONS

DRAINS LOCATION

Leading edge tank Outboard of nacelle underside of wing

Integral tank Underside of wing forward of aileron

Firewall fuel filter Underside of cowling forward of firewall

Sump strainer Bottom center of nacelle forward of the wheel well

Gravity feed line Outboard side of nacelle

Aft of wheel well (Prior to BB 1193)

NOTE

Windmi l l i ng t ime need no t be

charged against this time limit.

The maximum allowable fuel imbalance is

or MIL-I-85470 must be mixed with the fuelat refueling to ensure safe operation.

CAUTION





The maximum allowable fuel imbalance is1,000 pounds. Check the Airc raft Flight Manual Supplements for maximum imbalanceduring autopilot operation.

APPROVED FUEL ADDITIVE

Ant i - i c ing add i t ive conforming toSpecification MIL-I-27686 or MIL-I-85470are the only approved fuel additives.

Engine oil is used to heat the fuel on enteringthe fuel control. Since no temperature mea-surement is available for the fuel at this point,it must be assumed to be the same as the OAT.

Figure 5-11 is supplied for use as a guide inpreflight planning, based on known or forecastoperating conditions, to allow the operator tobecome aware of operating temperatures atwhich icing of the fuel control could occur. If oil temperature versus OAT indicates that iceformation could occur during takeoff or inflight, anti-icing additive per MIL-I-27686

Anti-icing additive must be properlyblended with the fuel to avoid dete-rioration of the fuel cells. The addi-tive concentration by volume shall bea minimum of 0.10% and a maxi-mum of 0.15 %.

Anti-icing additive per MIL-I-27686is blended in JP-4 fuel per MEL-T-5624 at the refinery, and no further treatment is necessary. Some fuelsuppliers blend anti-icing additivein their storage tanks. Prior to refu-eling, check with the fuel supplier todetermine whether or not the fuelhas been blended. To assure proper concentration by volume of fuel onboard, only enough additive for theunblended fuel should be added.

CAUTION

FUELING CONSIDERATIONSDo not put any fuel into the auxiliary tanks un-

less the main tanks are full.

The airplane must be statically grounded to thei i it d th i i it t





servicing unit, and the servicing unit mustalso be grounded.

The fuel filler nozzle must not be allowed torest in the tank filler neck as the filler neckmight be damaged.

It is recommended that a period of three hoursbe allowed to elapse after refueling so thatwater and other fuel contaminants have timeto settle. A small amount of fuel should thenbe drained from each drain point and checkedfor contamination. This practice is advanta-geous because fuel filters must be cleanedevery 100 hours. In addition, fuel filters must

be cleaned whenever fuel is suspected of beingcontaminated.

ZERO-FUEL WEIGHT

The maximum zero-fuel weight of the Super King Air 200 is 10,400 pounds. The maxi-mum zero-fuel weight of the B200 is 11,000

pounds.

1. Fuel is heated prior to entering the fuelcontrol unit by:

5. The fuel panel check tests electrical con-tinuity to which items?



QUESTIONS



A. Bleed ai r f rom the eng ine’scompressor

B. Engine oil, through an oil-to-fuel heatexchanger

C. The friction heating caused by theboost pump

D. An air-to-fuel heat exchanger prior to the fuel control unit

2. Which of the following is not affectedwhen the crossfeed switch is moved to theright or left?

A. The override function for auxiliary

fuel transfer B. The crossfeed valve

C. T h e s ta n d by p u mp o n t he s i d esupplying the fuel

D. The motive flow valve on the sidebeing fed

3. Which of the following is electrical ly

powered?A. Engine-driven boost pump

B. Standby boost pump

C. Engine fuel pump

D. Fuel manifold pump

4. Which of the following is a function of

the electric standby boost pump?A. It functions as a backup pump for use

in the event of primary boost pumpfailure

A. Firewall valves only

B. Firewall valves, standby boost pumps,and crossfeed valve

C. Standby boost pumps only

D. Firewall valves and standby boost

pumps prior to BB-1096

6. When is crossfeed use authorized?

A. For single-engine operation

B. For climbs above 20,000 feet whenaviation gas is used

C. W h e n on e s t an d b y p u m p i sinoperative

D. When fuel pressure decreases below10 ± 1psi

7. Which of the following limitations appliesto operation with aviation gas?

A. A maximum altitude of 20,000 feetwi th bo th s t andby boos t pumpsoperative and 150 hours betweenoverhauls

B. A maximum altitude of 31,000 feetwith standby boost pump inopera-t i v e a n d 1 5 0 h o u r s b e t w e e noverhauls

C. A maximum altitude of 20,000 feetwith one standby pump inoperativeand 150 hours between overhauls

D. A maximum of 150 hours betweenoverhauls only



8. Operation of the engine with the FUELPRESS light illuminated is limited towhich of the following?

A. Te n h o u rs o f en g i n e o pe r a t i o nbetween main engine fuel pumpoverhauls or before replacement.



B. Ten hours of operation above 20,000feet.

C. Unlimited operation below 20,000feet.

D. Respective engine shutdown.





The information normally contained in this chapter is notapplicable to this particular aircraft.

CHAPTER 7

POWERPLANTCONTENTS





Page

INTRODUCTION ................................................................................................................... 7-1

OVERVIEW ............................................................................................................................ 7-1ENGINE .................................................................................................................................. 7-2

General ............................................................................................................................. 7-2

Major Sections ................................................................................................................. 7-2

Operating Principles......................................................................................................... 7-6

ENGINE LUBRICATION SYSTEM ...................................................................................... 7-8

General ............................................................................................................................. 7-8

Oil Tank............................................................................................................................ 7-8

Pumps............................................................................................................................... 7-8

Oil Cooler......................................................................................................................... 7-8Indication ......................................................................................................................... 7-8

Fuel Heater....................................................................................................................... 7-9

Operation.......................................................................................................................... 7-9

ENGINE FUEL SYSTEM....................................................................................................... 7-9

General ............................................................................................................................. 7-9

Indication ......................................................................................................................... 7-9

Starting Ignition ............................................................................................................. 7-15

Autoignition................................................................................................................... 7-15

Indication ....................................................................................................................... 7-15

Operation 7-16





Operation ....................................................................................................................... 7-16

PROPELLER ........................................................................................................................ 7-16

General........................................................................................................................... 7-16

Feathering ...................................................................................................................... 7-18

Unfeathering and Reversing .......................................................................................... 7-18

Basic Principles.............................................................................................................. 7-18

Control ........................................................................................................................... 7-18

Overspeed Control ......................................................................................................... 7-19

Fuel Topping (Power Turbine) Governor ...................................................................... 7-20

Reverse Operation.......................................................................................................... 7-21

Beta Mode Control......................................................................................................... 7-21

Propeller Operating Principles....................................................................................... 7-21

Powerplant Power Control............................................................................................. 7-23

Engine Instrumentation.................................................................................................. 7-26

Synchroscope ................................................................................................................. 7-28

Synchrophasing.............................................................................................................. 7-29

PROPELLER FEATHERING............................................................................................... 7-31

Autofeathering ............................................................................................................... 7-31

Operating Principles ...................................................................................................... 7-32

LIMITATIONS (POWERPLANT)........................................................................................ 7-34

General........................................................................................................................... 7-34

Powerplant ..................................................................................................................... 7-35

Engine Operating Limits 7-35





Engine Operating Limits................................................................................................ 7 35

Approved Fuels.............................................................................................................. 7-35

Propeller......................................................................................................................... 7-37

Powerplant Instrument Markings................................................................................... 7-37

Starter Limits ................................................................................................................. 7-37

QUESTIONS......................................................................................................................... 7-39

ILLUSTRATIONS

Figure Title Page

7-1 Super King Air 200 .................................................................................................. 7-2

7-2 PT6A Engine 7-3





7 2 PT6A Engine ............................................................................................................ 7 3

7-3 Engine Cutaway ....................................................................................................... 7-4

7-4 Compressor Bleed Valves......................................................................................... 7-5

7-5 Engine Gas Flow and Stations.................................................................................. 7-7

7-6 Oil Pressure/Temperature Gages.............................................................................. 7-8

7-7 Chip Detection Lights .............................................................................................. 7-9

7-8 Oil System Schematic ............................................................................................ 7-10

7-9 Fuel Low-Pressure Lights ...................................................................................... 7-11

7-10 Fuel Flow Gages..................................................................................................... 7-11

7-11 Fuel Schematic ....................................................................................................... 7-12

7-12 Simplified Fuel Control Schematic........................................................................ 7-14

7-13 Engine Start and Ignition Switches ........................................................................ 7-15

7-14 Engine Autoignition Switches................................................................................ 7-15

7-15 Ignition System Schematic..................................................................................... 7-16

7-16 Propellers................................................................................................................ 7-17

7-17 PROP GOV TEST Switch ..................................................................................... 7-19

7-18 Propeller Governor Test Schematic ....................................................................... 7-20

7-19 Propeller Onspeed Schematic ................................................................................ 7-22

7-20 Propeller Overspeed Schematic ............................................................................. 7-22

7-24 Propeller Control Lever.......................................................................................... 7-26

7-25 Friction Control Knobs .......................................................................................... 7-26

7-26 ITT Gages............................................................................................................... 7-27

7-27 Torque Gages ......................................................................................................... 7-27





q g

7-28 Propeller RPM Gages............................................................................................. 7-28

7-29 Engine RPM Gages................................................................................................ 7-28

7-30 Propeller Synchroscope and Switches (Type II) .................................................... 7-28

7-31 Type II System Schematic...................................................................................... 7-29

7-32 Type I System Schematic....................................................................................... 7-30

7-33 Propeller Synchroscope and Switch (Type I)......................................................... 7-31

7-34 Sync Light .............................................................................................................. 7-31

7-35 AUTOFEATHER Switch ....................................................................................... 7-31

7-36 Autofeather Lights ................................................................................................. 7-32

7-37 Autofeather System Schematic (Both Power Levers at Approximately

90% N1; Right Engine has Failed) ......................................................................... 7-33

7-38 Autofeather Test Schematic (Left Power Lever Below 200 ft-lb;

Right Power Lever Above 400 ft-lb) ..................................................................... 7-34

TABLES

Table Title Page

7-1 Engine Operating Limits

(PT6A-42 Engine BB-1439, 1444 and Subsequent).............................................. 7-35





7-2 Engine Operating Limits (PT6A-42 Engine Prior to BB-1439,

1444 and Subsequent) ............................................................................................ 7-35

7-3 Engine Operating Limits (PT6A-41 Engine) ......................................................... 7-36

7-4 Powerplant Instrument Markings........................................................................... 7-38

CHAPTER 7

POWERPLANT





INTRODUCTIONThis chapter deals with the powerplant of the Super King Air 200. All values, such as for pressures, temperatures, rpm, and power are used for illustrative meanings only. Actualvalues must be determined from the appropriate sections of the approved flight manual.

Information in this chapter must not be construed as being equal to or superseding anyinformation issued by or on behalf of the various manufacturers or the Federal Aviation

Administration.

OVERVIEW

#1 DC

GEN





ENGINEGENERAL

The engines used on the Super King Air 200are designated PT6A-41, while the B200 usesPT6A-42.

The PT6A (Figure 7-2) is a free-turbine, re-

verse-flow, lightweight turboprop engine, ca-pable of developing 850-shaft horsepower (903 equivalent shaft horsepower [ESHP]).

PT6 i d l b b 1960

MAJOR SECTIONS

For the purpose of this chapter, the engine(Figure 7-3) is divided into seven major sections.

1. Air Intake Section

2. Compressor Section

3. Combustion Section

4. Turbine Section

5. Exhaust Section

6. Reduction Gear Section

Figure 7-1. Super King Air 200



STARTER-GENERATOR



FUEL PUMP/FCU

TACHOMETER-GENERATOR(NG)

OIL SCAVENGE

PUMPS ANDFUEL BOOST PUMP

OPTIONALACCESSORY

DRIVE

PROPELLEROVERSPEEDGOVERNOR

PROPELLERGOVERNOR

TACHOMETER-GENERATOR

(NF)

AFT

7 - 4



F OR T R

AI NI N G P URP O S E S

O

NL Y

S UP E R

K

I N G

A I R

2 0 0 / B 2 0 0

P I L O T T R A I NI N G MA N U

A L

F l i g

h t S af e t y


Figure 7-3. Engine Cutaway



Air Intake SectionThe compressor air intake consists of a circular screen-covered, aluminum casting. Air is di-rected to the air intake by the nacelle air scoopon the lower side of the nacelle. The functionof the air intake section is to direct airflow tothe gas generator compressor.

by overboarding, or bleeding axial stage air toreduce backpressure on the centrifugal stage(Figure 7-4). This pressure relief helps prevent

compressor stall of the centrifugal stage.

The compressor bleed valves, one on eachside of the compressor, are pneumatic pis-tons, which reference the pressure differential



Compressor SectionThis section consists of a four-stage compres-

sor assembly, made up of three axial stages andone centrifugal stage. The function of the com-pressor is to compress and supply air for com-bustion, combustion cooling, pressurizationand pneumatic services, compressor bleed valveoperation, and bearing sealing and cooling.

Compressor Bleed Valves

At low N1 rpm, the compressor axial stagesproduce more compressed air than the cen-trifugal stage can use. Compressor bleed valvescompensate for this excess airflow at low rpm

, pbetween the axial and centrifugal stages.Looking forward, the low-pressure valve is lo-cated at the 9 o’clock position and the highpressure at 3 o’clock. The function of these

valves is to prevent compressor stalls andsurges in the low N1 rpm range.

At low N1 rpm, both valves are in the open po-sition. At takeoff and cruise N1 rpm, above ap-proximately 90%, both bleed valves will beclosed. If both compressor bleed valves wereto stick closed below approximately 90% N1,a compressor stall would result.

If one or both valves were to stick open, theITT would increase and torque decrease whileN1 rpm remained constant.

CONTROL PRESSURE

PISTON DAMPER(SPRING LOAD)

SLEEVE

AMBIENT PRESSURE

FINALORIFICE

PRIMARYORIFICE

DELIVERYAIR PASSAGE

P3

Combustion SectionThe PT6 engine utilizes an annular combustionchamber. Two, high-energy igniter plugs areinstalled in the combustion chamber, as wellas 14 equally-spaced simplex fuel nozzles.

Turbine Section

The function of the accessory section is todrive the engine and airplane accessories,which include:

• Fuel control unit (FCU) and high-pres-sure fuel pump

• Lubricating pump/scavenge pumps

• N tachgenerator





Turbine SectionThe PT6A uses three reaction turbines: a free,two-stage axial propeller (power) turbine anda single-stage compressor turbine. The two-

stage power turbine extracts energy from thecombustion gases to drive the propeller and itsaccessories through the planetary reductiongears. This combination is defined as NP. Thesingle-stage compressor turbine extracts en-ergy from the combustion gases to drive thegas generator compressor and the accessorygear section. This combination is defined asN1.

Exhaust SectionThis section is located immediately aft of thereduction gear section and it consis ts of an an-nular exit plenum, a heat-resistant cone, andtwo exhaust outlets at the 9 o’clock and 3o’clock positions.

Reduction Gear SectionThe reduction gear section at the front of theengine is a two-stage, planetary type. The pri-mary function of the reduction gear section isto reduce the high rpm of the free turbine tothe value required for propeller operation.

The reduction gear section is also used for

torquemeter operation and includes drive sec-tions for the propeller governor (with fueltopping governor sensing), the propeller over-speed governor, and a propeller tachgenera-

• N1 tachgenerator

• DC starter-generator

• Refrigerant compressor (right engineonly)

• Low-pressure fuel boost pump

Other drive pads are provided for optional op-erator equipment (Figure 7-2).

OPERATING PRINCIPLES

When the engine is rotating (Figure 7-5), air is

inducted through the nacelle air scoop to the en-gine air intake. Airflow is turned 180° in a for-ward direct ion and is then progressivelyincreased in pressure by a three-stage axial-flow and single-stage centrifugal-flow com-pressor. It is then directed forward throughdiffuser ducts towards the forward side of thecombustion chamber. The airflow is again turned180° and enters the combustion chamber, where

metered fuel is added to the air by 14 fuel spraynozzles. Two high-energy igniter plugs ignitethe gas mixture. The expanding gases moverearward through the combustion chamber andturn 180° forward to enter the turbine section.The compressor turbine extracts sufficient en-ergy from the expanding gases to drive the four-stage compressor and the accessory gear section.The remaining two stages of the free power tur-

bine extract the maximum amount of the re-maining energy from the combustion gases todrive the propeller and the propeller accessoriesthrough the reduction gearbox. The two-stage

bi i f bi d i l



POWER SECTION



COMPRESSOR SECTION

M O D U L E

1

M O D U L E

2

GAS GENERATOR SECTION

ENGINE LUBRICATIONSYSTEM

GENERAL

The engine lubrication system is a completelyself-contained and fully automatic system. It

INDICATION

Engine Oil PressureEngine oil pressure is sensed by a transmitter in the pressure pump outlet line and suppliedto a combination, pressure-temperature gage(Figure 7-6) on the engine instrument panel.The oil press re s stem req ires DC po er





y yprovides for cooling and lubrication of theengine bearings and the reduction and acces-sory drive gears, and for operation of the pro-peller control system, the torquemeter system,

the torque limiter, and the fuel heater system.

The engine oil system is a dry-sump systemconsisting of pressure, scavenge, and cen-trifugal air breather systems.

OIL TANK

The oil tank forms an integral part of the en-gine, located between the aft end of the com-pressor air inlet and the forward end of theaccessory gearbox.

A filler and dipstick are located at the 11o’clock position on the accessory case. The oiltank is vented to a centrifugal breather to pro-vide for air-oil separation.

PUMPS

The oil pumps consist of one pressure ele-ment and four scavenge elements. The pres-sure pump supplies lubrication pressure to thebearings and the accessory system drive gears.In addition, the pressure pump supplies oil tothe propeller control system, the torquemeter

system, reduction gears and the torque limiter.

OIL COOLER

The oil pressure system requires DC power.

Engine Oil TemperatureOil temperature is sensed by a resistance bulband transmitted to the same combination pres-sure/temperature gage (Figure 7-6) on the en-gine instrument panel. The power supply for the gage is from the DC power system.

Chip DetectionFor BB-1439, 1444 and subsequent, the cau-tion annunciator panel contains two amber lights marked L CHIP DETECT and R CHIPDETECT (Figure 7-7). Prior to BB-1444, ex-cept 1439, these are red lights on the warning

annunciator panel. They are operated by amagnetic chip detector located at the bottomof each reduction gearbox.

When either light illuminates it indicates that

Figure 7-6. Oil Pressure/TemperatureGages





FUEL HEATER

Oil scavenged from the accessory gearcase isdirected through an oil-to-fuel heater prior toits return to the oil tank.

OPERATIONWhen the engine is running, the oil pressurepump (Figure 7-8) draws oil from the tank, de-velops a higher pressure with the oil, and di-rects pressure oil through various filters tothe engine bearings, the accessory and re-duction drive gears, the propeller governor,and the engine torquemeter system. Oil pres-

sure is regulated and limited by a relief valve.Oil pressure and temperature are sensed andtransmitted to the cockpit gages. All oil isscavenged to the accessory gearcase except thereduction gearcase oil, which goes directly tothe oil cooler. A screened scavenge pump re-turns the gearcase oil to the tank through theoil-fuel heater; another scavenge pump scav-enges oil from the reduction gearcase and re-

turns this oil to the tank through the oil cooler.

ENGINE FUEL SYSTEM

divider, and two fuel manifolds each withseven simplex fuel nozzles.

INDICATION

Fuel PressureThe warning annunciator panel red lightsmarked L FUEL PRESS and R FUEL PRESS(Figure 7-9) are operated by pressure switchesthat sense outlet pressure at the engine-drivenboost (LP) pump. The lights will come on toindicate abnormally low (10 ± 1 psi) fuel pres-sure to the (HP) engine pump.

Fuel FlowFuel flow information is sensed by a transmit-ter in the engine fuel supply line and suppliedto the fuel flow gages (Figure 7-10) on the cen-ter instrument panel. For BB-225 airplanes andsubsequent, they are DC powered. Prior to BB-225, these fuel flow gages are AC powered.

FUEL SYSTEM OPERATION

The fuel control system for PT6A engines is

BB-1439, 1444 AND AFTER PRIOR TO BB-1444, EXCEPT 1439

Figure 7-7. Chip Detection Lights

7 - 1 0

S

FROM COOLER

TO COOLER

OIL TANK BREATHEROIL DIPSTICK

DIVERTERVALVE

PROPELLER GOVERNORAND BETA CONTROL

TORQUEMETER OILCONTROL VALVE



F OR T RAI NI N G P URP O S E S

O

NL Y

S UP E R

KI N G

A I R

2 0 0 / B 2 0 0

P I L O T T R A I NI N G MA N U

A L

F l i g

h t S af e t y


TO OIL PRESSUREINDICATOR

TO OIL TEMPERATUREINDICATOR

TO OIL PRESSUREANNUNCIATOR

FUELHEATER

SCAVENGEPUMP

ACCESSORYGEARBOX DRAIN

BYPASSVALVE

OIL

TANK

OIL FILTER ANDCHECK VALVE

BYPASS VALVEOVERPRESSURE

RELIEF VALVE

TANKDRAINTORQUEMETER

& TORQUE LIMITER

TORQUEMETERPRESSURE(INDICATOR)CHIP

DETECTOR

OIL SUPPLYTO PROPELLER

PRESSURE OIL

PROPELLER SUPPLY OIL

SCAVENGE OILBREATHER AIR

TORQUEMETER PRESSURE

LEGEND

PRESSUREREGULATINGVALVE

* OPTION

*

Figure 7-8. Oil System Schematic





pump, oil-to-fuel heat exchanger, high-pres-sure fuel pump, fuel control unit, fuel cutoff valve, fuel flow transmitter, flow divider, anddual fuel manifold with 14 simplex nozzles.

The low-pressure boost pump is engine-drivenand operates when the gas generator shaft(N1) is turning, to provide sufficient fuel headpressure to the high-pressure pump to main-tain proper cooling and lubrication. The oil-to-fuel heat exchanger uses warm engine oilto maintain a desired fuel temperature at thefuel pump inlet to prevent icing at the pump

filter. This is done with automatic temperaturesensors and requires no action by the pilot.

Fuel enters the engine fuel system through

let filter. Flow rates and pressures will varywith gas generator (N1) rpm. Its primary pur-pose is to provide sufficient pressure at the fuelnozzles for a good spray pattern at all modesof engine operation. The high-pressure pumpsupplies fuel at approximately 800 psi to thefuel side of the FCU.

Two valves included in the FCU ensure con-sistent and cool engine starts. When the igni-tion or start system is energized, the purgevalve is electrically opened to clear the FCUof vapors and bubbles. The excess fuel flowsback to the nacelle fuel tank. The spill valve,referenced to atmospheric pressure, adjuststhe fuel flow for cooler high-altitude starts.

Between the FCU fuel valve and the enginecombustion chamber, and part of the FCU, aminimum pressurizing valve cuts off fuel flowduring starts until fuel pressure builds suffi-ciently to maintain a proper spray pattern inthe combustion chamber. About 70 psi is re-quired to open the minimum-pressurizingvalve. The engine-driven high-pressure fuelpump maintains this required pressure. If thepump should fail, the valve would close andthe engine would flame out.

Figure 7-9. Fuel Low-Pressure Lights

Figure 7-10. Fuel Flow Gages

7 - 1 2

S

ENGINE

DRIVENFUEL PUMP

(HIGHPRESSURE)

FUEL FLOWTRANSMITTER

OIL-TO-FUELHEAT

EXCHANGER

FUELCONTROLUNIT

COCKPITGAGE

*

TOFUELTANK

PURGE LINE




O

NL Y

S UP E R

KI N G

A I R

2 0 0 / B 2 0 0

P

I L O T T R A I NI N G MA N U

A L

F l i g

h t S af e t y


FUEL PRESS

FIREWALLFUEL FILTER

ANDMANUAL

SHUT-OFFVALVE

ENGINEDRIVENBOOSTPUMP

TRANSDUCER

FLOWDIVIDER

P3 PURGE**TANK

P3 AIR

P3 AIR

NP

FUELTOPPING

GOVERNOR

N1

POWERAND

CONDITIONLEVERS

* PRIOR TO BB-1401, FUEL FLOW TRANSMITTERLOCATED UPSTREAM OF OIL-TO-FUEL HEAT EXCHANGER

** PRIOR TO BB-666, FUEL DRAIN COLLECTOR TANK

Figure 7-11. Fuel Schematic



fuel spray nozzles in the combust ion chamber.As the engine accelerates through approxi-mately 40% N1, fuel pressure is sufficient toopen the transfer valve to the secondary fuelnozzles. At this time all 14 nozzles are deliv-ering atomized fuel to the combustion cham-ber. This progressive sequence of primary andsecondary fuel nozzle operation providescooler starts. On engine startups, there is a def-

ing valve prevents fuel flow to the engine untilthe fuel pressure has increased enough to en-sure proper atomization of the fuel at the noz-zles. Once the minimum pressure valve hasopened, fuel will flow to the flow divider andthe fuel nozzles.

Aside from opening and closing the fuel cutoff valve, the condition lever adjusts N1 speed from



g p ,inite surge in N1 speed when the secondary fuelnozzles cut in.

In order to improve cold weather starting , SB-3214 changed seven primary and seven sec-ondary nozz les to 10 p r imary and four secondary. At a later date, SB-3250 changedthe nozzles back to seven and seven, but witha different arrangement and an improvedburner can.

During engine shutdown on BB-666 and sub-

sequent, any fuel left in the manifold is forcedout through the nozzles and into the combus-tion chamber by purge tank pressure. As thefuel is burned, a momentary surge in N1 rpmshould be observed. The entire operation is au-tomatic and requires no input from the crew.On BB-2 through BB-665, an EPA collector tank is used instead of the purge tank system.

FUEL CONTROL UNIT (FCU)

GeneralThe fuel control unit (Figure 7-12), which isnormally referred to as the FCU, has multiplefunctions, but its main purpose is to meter theproper fuel amount to the nozzles in all modesof engine operation. It is calibrated for start-

ing flow rates, acceleration, and maximumpower. The FCU compares gas generator speed(N1) with the power lever setting and regulatesfuel to the engine fuel nozzles. The FCU also

, j 1 pLOW IDLE to HIGH IDLE. The power lever,by adjusting the governor position in the FCU,adjusts the fuel-metering valve to allow more

or less fuel to the spray nozzles. In summary,the power lever controls fuel to the engine byadjusting the governor position, which in turnrepositions the fuel-metering valve in the FCU.

FCU OperationThe pneumatic section of the FCU determinesthe flow rate of fuel to the engine for all op-

erations. The power levers control enginepower from idle through takeoff power by op-eration of the gas generator (N1) governor inthe FCU. Increasing N1 rpm results in in-creased engine power.

For explanation purposes, consider the N1 gov-ernor bellows as a diaphragm. P3 air is intro-duced into the bellows in a manner that sets up

a differential pressure on each side of the di-aphragm. Therefore, any change in P3 pres-sure will move the diaphragm. When pressureis increased, the fuel-metering valve attachedto the bellows will move in an opening di rec-tion to increase fuel flow and increase N1 rpm.

As P3 pressure decreases, fuel flow also de-creases which reduces the N1 rpm. The N1

governor increases or decreases P3 pressure inthe bellows by varying the opening of relief orifices in the bellows.

7 - 1 4

F S U




O

NL Y

UP E R

KI N G

A I R

2 0 0 / B 2 0 0

P

I L O T T R A I NI N G MA N U

A L

F l i g

h t S af e t y


N1

N2

AIR FUEL

Figure 7-12. Simplified Fuel Control Schematic



the governor. When N1 rpm reaches the desiredspeed, the governor adjusts the P3 orifice to re-duce pneumatic pressure to match the fuel pres-sure required to maintain the desired N

1rpm.

The fuel topping (power turbine) governor protects against power turbine overspeed. If an overspeed occurs, and the propeller goes be-yond 106% of the requested propeller rpm,h f l i i d

switch is located on the left switch panel. It hasthree marked positions: ON–OFF–STARTERONLY. The ON position (UP) is lever locked andit provides for engine cranking and ignit ion op-eration. The STARTER ONLY position is a mo-mentary (spring loaded to center hold down)position and it only provides for engine motor-ing. In this position, the igniters do not function.



the fuel topping governor vents air to reducefuel flow. Reducing fuel flow decreases N1speed and accordingly power turbine speed.With propellers in reverse, the fuel-toppinggovernor restricts fuel flow to approximately95% of the requested propeller rpm.

ENGINE IGNITIONSYSTEM

GENERALThe engine ignition system is a high-energy,capacitance type consisting of a dual-circuitigniter box and two igniter plugs in the com-bustion chamber. The ignition system is di-vided into starting ignition and autoignition.

STARTING IGNITION

A three-position lever lock switch for each en-gine (Figure 7-13) controls this system. The

AUTOIGNITION

The autoignition system is controlled by a

two-position switch for each engine markedARM and OFF (Figure 7-14). Turning on anAUTO IGNITION switch arms the igniter circuit to an engine torque switch that isnormally open when the engine is develop-ing more than 400 foot-pounds of torque. Thesystem must be armed prior to takeoff andfor all phases of flight, and it should beturned off only after landing. If engine torque

drops to 400 foot-pounds or less when theautoignition is armed, the ignition systemwill energize to prevent engine flameout if the loss of power was caused by a momen-tary fuel or air interruption.

INDICATION

Green annunciator lights marked L and R IG-NITION ON tells the pilot that the igniters arereceiving power.

OPERATION

Starting Ignition

When DC power is available, turning on theignition and engine start switch (Figure 7-15)will apply DC power to the ignition ON light,FCU purge valve, and to the ignition exciter.The exciter, which operates at three cyclesper second will apply high energy power to

400 foot-pounds, the torque switch will closeand apply DC power to the ignition ON light,the FCU purge valve, and to the ignition ex-citer. Ignition will be continuous until power increases above 400 foot-pounds.

PROPELLER





per second, will apply high-energy power tothe igniter plugs in the combustion chamber.

AutoignitionWhen the AUTO IGNITION switch (Figure7-14) is at the ARM position, the ignitionsystem is inactive as long as engine torque isabove 400 foot-pounds. If torque decreases to

GENERAL

The PT6A engine drives a three- or four-blade,

oil-operated propeller (Figure 7-16). A threeblade Hartzell propeller is used on BB-2through B-1192 and has a blade angle rangeof +90° to –9°. A three blade McCauley pro-peller is used on BB-1193 through 1438, and1440 through 1443, and has a blade angle

ONOFFSTARTER ONLYIGNITION

ANDENGINE STARTER

ARMOFF

AUTOIGNITION

CLOSE400 FT-LBS

TORQUE SW

IGN ON

IGN EXCITER

IGNITER PLUGS





3-BLADE PROPELLER

range of +86.8° to –10°. Airplanes BB-1439,1444 through 1508 have a four -bladedMcCauley propeller with a blade angle range of +87.5° to –10°. Airplanes BB-1509 and subse-quent have a four-bladed Hartzell propeller with a blade angle range of +87.9° to –11°. Thepropeller control system provides for constant-speed operation, full feathering, reversing, andBeta mode control. Feathering is induced bycounterweights and springs

BASIC PRINCIPLES

Constant-speed propellers operated in threeconditions under the control of a propeller governor. These conditions are:

• Onspeed

• Overspeed

• Underspeed





counterweights and springs.

If an engine flames out in flight or if the pilotselects the condition lever to CUTOFF, thepropeller will not feather because of the wind-milling effect and governor action. Featheringin flight should be manually selected by usingthe propeller control lever.

A conventional oil-operated propeller gover-nor achieves normal propeller operation inthe constant speed range. A preset oil-operated

overspeed governor is provided in case of fail-ure of the normal propeller governor. In ad-dition to the normal and overspeed propeller governors, a fuel topping function, integralwith the primary governor, provides protec-tion against propeller overspeed, as well aslimiting rpm in the reverse ranges.

FEATHERING

Feathering is a function of counterweights at-tached to each blade root and spring forces inthe propeller cylinder.

UNFEATHERING ANDREVERSING

Unfeathering and reversing functions are doneby hydraulic (engine oil) pressure developedby a high-pressure oil pump, which is an in-tegral part of the propeller primary governor.

p

Onspeed

Onspeed is defined as the condition of oper-ation in which the selected rpm and actualrpm are the same.

Overspeed

Overspeed is the condition of operation inwhich the actual rpm is greater than the se-lected rpm.

Underspeed

Underspeed is the condition of operation inwhich the actual rpm is less than the se-lected rpm.

CONTROL

Speed (rpm) control is a function of the pro-peller governor. This unit is engine-drivenand operates on the principle of balancing twoopposing forces, both of which are variables.These forces are speeder spring force and fly-weight force.

Speeder Spring ForceSpeeder spring force is a function of, and var-ied by, the position of the propeller controll

If the speeder spring force is greater than fly-weight force, the propeller would be operat-ing in an underspeed condition.

If the flyweight force is greater than speeder spring force, the propeller would be operatingin an overspeed condition.

When the speeder spring and flyweight forcesare equal the propeller is onspeed

lever (2,120 being the highest setting, propeller levers full forward).

Test SystemThe overspeed governor incorporates a testsystem controlled by a two-position switch(Figure 7-17) for both propellers. The switchis marked PROP GOV TEST. The switch is lo-cated on the pilot’s left subpanel (BB 2 through





are equal, the propeller is onspeed.

Unbalance of speeder spring and flyweightforces is used to position a pilot valve to ac-complish the following:

• Direct governor boosted high oil pres-sure to the propeller servo piston to re-duce the blade angle.

• Shut off governor-boosted high oil pres-sure to the propeller servo piston andconnect the piston chamber to the oil

sump, allowing the counterweights andpropeller spring force to increase theblade angle, to include feather if de-sired. When the speeder spring and fly-weight forces are equal, the pilot valveis positioned appropriately to maintaina constant blade angle.

OVERSPEED CONTROLThe normal rpm control range of the primarygovernor is from 1,600 rpm to 2,000 rpm; thelatter is 100% rpm.

If the primary governor fails to limit rpm to2,000, a second (overspeed) governor, driven bythe reduction gearbox, operates in parallel with

the primary governor. This is called the over-speed governor. The overspeed governor has apreset speeder spring tension which limits pro-peller rpm to the preset limit of 2,120 rpm (prior

cated on the pilot s left subpanel (BB-2 through162 had two switches).

A solenoid valve is associated with each over-speed governor. The valve is energized whenthe PROP GOV TEST switch is moved to theTEST position. When energized, the valve ap-plies governor pump pressure to change thefixed value of the overspeed governor as listedabove, to a range of from 1,830-1,910 rpm.

Operating PrinciplesWith the engine running and the propeller control lever full forward, moving the gov-ernor test switch to TEST will open a solenoid

valve and admit primary governor pump pres-sure to a hydraulic reset valve on the over-speed governor. Movement of the reset valvewill raise the pilot valve, simulat ing an over-

Figure 7-17. PROP GOV TEST Switch



PROP GOVTEST SWITCH SPEED SETTING SCREW

SPEEDERSPRING

FLYWEIGHT

FLYWEIGHT

PROPELLER LEVER

SPEEDERSPRING TEST

SOLENOID



FUEL TOPPING (POWERTURBINE) GOVERNOR

If a mechanical failure causes the propeller tolock or stick, it will not respond to oil pres-sure changes. The primary and overspeed gov-ernors, although still operating normally, willbe unable to control propeller rpm with oilpressure. The fuel topping governor (FTG), anintegral part of the primary governor, acts toreduce fuel flow, which in turn reduces pro-

ll With l k d ll (fi d

The fuel-topping governor is designed to ventair pressure from the FCU, which results in afuel flow reduction. The propeller rpm atwhich the FTG activates is determined by pro-peller control lever position. With the pro-peller locked, the FTG will reduce fuel flowwhen the overspeed reaches approximately106% of the selected propeller rpm.

The FTG utilizes the same flyweights andpilot valve mechanism of the primary gover-nor If the primary governor fails the fuel top

TRANSFERGLAND

RESTRICTOR

FINE PITCH

ENGINE OILSUPPLY

OIL DUMP TOREDUCTION BOX

OIL DUMP TOREDUCTION BOX

PROPELLEROVERSPEEDGOVERNOR

CONSTANTSPEEDPROPELLERGOVERNOR

RELIEFVALVE

GOVERNORPUMP

Figure 7-18. Propeller Governor Test Schematic

REVERSE OPERATION

When full reverse is selected, the power leverssend three commands:

1. Spool the compressor to 83% ± 5% N1with a fuel flow increase.

2. Decrease the propeller blade angle to –9°or –10°.

3 Reset the FTG to 95% of the rpm se

PROPELLER OPERATINGPRINCIPLES

OnspeedWhen the upward force of the governor fly-weights (Figure 7-19) is equal to the downwardforce of the speeder spring, the governor pilotvalve is positioned to shut off the governor pump pressure from the propeller piston and





3. Reset the FTG to 95% of the rpm se-lected by the propeller lever.

The maximum allowable propeller speed in re-verse is 1,900 rpm; however, this is not anoverspeed limitation for the propeller or power turbine. The 1,900-rpm limit, which is con-trolled by the FTG, assures that the propeller does not attain 2,000 rpm, which brings thepropeller on speed and begins to interfere withthe reverse operation.

BETA MODE CONTROLBeta control defines a range of operation inwhich the pilot can reduce the residual idlethrust of the propeller by reducing blade angle.This reduction in blade angle and, therefore,propeller thrust, is accomplished by liftingthe power levers aft into the ground fine rangeon BB-1439, 1444 and subsequent. For prior

aircraft this is accomplished by lifting thepower levers aft to a position just above thered and white lines (reverse range) on thethrottle quadrant.

The propeller used in the King Air Series in-cludes a Beta valve, which forms an integralpart of the propeller governor. The pilot canmechanically position this valve, within a lim-

ited (ground) range, described above, to effectpropeller blade angle changes. Propeller servopiston movement is fed back to the valve bya mechanical follow-up system to null the

pump pressure from the propeller piston andisolate the propeller cylinder from the gearcasedrain. This, in effect, hydraulically locks the

blades at a specific angle. This condition doesnot prevail for very long as changes in altitude,temperature, airspeed, and inherent leakage atthe prop transfer sleeve require blade anglechanges. In effect, in any constant-speed con-dition, the governor is hunting through a verynarrow range to maintain the selected rpm.

OverspeedWhen an overspeed occurs, the governor fly-weight force (Figure 7-20) exceeds the speeder spring force. This occurs when the propeller hasaccelerated above the selected rpm. The in-creased flyweight force will raise the governor pilot valve and reduce oil pressure at the propeller piston, allowing the counterweights and springto increase blade angle and decelerate the pro-peller until an onspeed condition occurs.

Underspeed

When an underspeed condition occurs, the pro-peller decelerates below the selected rpm andthe speeder spring force overcomes the force of the flyweights (Figure 7-21). As a result, the

pilot valve moves down and allows the gover-nor pump to apply oil pressure to the propeller servo piston, resulting in a decrease in bladeangle. This allows the propeller to accelerate



HYDRAULICOVERSPEEDGOVERNOR

REVERSELEVER

OIL

PRIMARY PROP GOVERNOR1,600 - 2,000 RPM

PILOTVALVE

2,120 RPMNORMAL(2,080 PRIORTO BB-1444,EXCEPT 1439)

GOVERNORPUMP

PROPLEVER



GOVERNORPUMP

OIL

PRIMARY PROP GOVERNOR

1,600 - 2,000 RPMOVERSPEED

REVERSELEVER


TOCASE

TOCASE

BETAVALVE

AUTOFEATHER SOLENOID (NC)

PROPLEVER

PILOTVALVE


TRANSFERGLAND

LOW PITCH(HIGH OIL PRESSURE)


TOCASETO

CASE

BETAVALVE

Figure 7-19. Propeller Onspeed Schematic



GOVERNORPUMP

OIL


UNDERSPEED

REVERSELEVER


PROPLEVER

PILOTVALVE




POWERPLANT POWERCONTROL

The powerplant (engine-propeller combina-

tion) is controlled by the interaction of threelevers (Figure 7-22): a condition lever, a power lever, and a propeller control lever.

Condition LeverThe condition lever (Figure 7-22) is me-chanically connected to the FCU to operatea fuel cutoff valve that shuts off metered

fuel to the fuel manifold.

The condition levers located on the power lever quadrant (last two levers on the right

1439; or 52% in the PT6A-41) gas genera-tor, or N1 rpm. HIGH IDLE will establish afuel flow that will sustain 70% N1 rpm. Thereis a progressive increase in fuel flow as thecondition lever is moved from LOW IDLE toHIGH IDLE, and any rpm may be selected be-tween LOW IDLE and HIGH IDLE.

Power LeversPower levers (Figure 7-22) are located on thepower lever quadrant (first two levers on the leftside) on the center pedestal and they are me-

chanically interconnected through a cam box tothe FCU, the Beta valve and follow-up mecha-nism, and the fuel topping (NP) governor. Thepower lever quadrant permits movement of the

TOCASE

TO

CASE

BETAVALVE

TRANSFERGLAND

LOW PITCH(HIGH OIL PRESSURE)


Figure 7-21. Propeller Underspeed Schematic



POWERLEVERS




POWERLEVERS

CONDITIONLEVERS

CONDITIONLEVERS

PROPELLERLEVERS

PROPELLERLEVERS

1444, except 1439) or reverse range. The pilotmust lift the power levers up and over this de-tent to select ground fine/Beta or reverse.

The function of the power levers in the forwardthrust (Alpha) range is to establish a gas gen-erator rpm through the gas generator gover-nor (N1) and a fuel flow that will produce andmaintain the selected N1 rpm. In the groundfine (Beta) range, the power levers are usedto reduce the propeller blade angle thus re

Ground Fine (Beta) andReverse Control

The geometry of the power lever linkage (Figure7-23) through the cam box is such that power lever increments from idle to full forward thrusthave no effect on the position of the Beta valve.When the power lever is moved from idle intothe reverse range, which requires the power levers to be lifted over a second gate in BB-1439 1444 d b i i i h





to reduce the propeller blade angle, thus re-ducing residual prop thrust. In the reverserange, the power lever functions to:

1. Select a blade angle proportionate to theaft travel of the lever.

2. Select a fuel flow that will sustain theselected reverse power.

3. Reset the fuel topping governor (NP)from its normal 106% to a range of ap-proximately 95%.

1439, 1444 and subsequent, it positions theBeta valve to direct governor pressure to thepropeller piston, decreasing blade angle throughzero into a negative range (Figure 7-23). Thetravel of the propeller servo piston is fed backto the Beta valve to null its position and, in ef-fect, provide many negative blade angles all theway to full reverse. The opposite will occur when the power lever is moved from full reverseto any forward position up to idle, therefore pro-viding the pilot with manual blade angle con-

trol for ground handling.

GOVERNORPUMP

OIL


POWER/REVERSELEVER


TOCASEDRAINTO

CASE DRAIN

BETAVALVE NC

PROPLEVER

POWER

LEVER

PILOTVALVE

2,080 RPMNORMAL

REV IDLE LO HI

FX LO HI

APPROXIMATELY1870 RPM IN TES

MODE

Propeller Control LeverThe propeller control lever (Figure 7-24) op-erates in the throttle quadrant (the two cen-

ter levers) on the center pedestal and it ismechanically connec ted to the primary pro-peller governor. In the forward thrust, or constant-speed range, the propeller controllever selects rpm from low rpm to high rpm(1,600 to 2,000 rpm) by changing the set-ting of the primary propeller gove rnor. The





ting of the primary propeller gove rnor. Thepropeller control lever is also used to feather the propeller by moving the lever aft into the

feather detent position. This action posi-tions the primary propeller governor ’s pilotvalve to dump oil from the propeller se rvopiston chamber and allows the propeller counterweights and springs to move the pro-peller blades to the full feather position. Adetent (requiring more force to overcome)at the low rpm position, prevents inadver-tent movement of the propeller lever into the

feather range.

ENGINE INSTRUMENTATION

Engine Temperature (ITT)GagesEngine operating temperature at station T

5is

sensed by eight thermocouple probes locatedbetween the gas generator turbine and the firststage power turbine. The probes are connectedin parallel to provide the best average reading.

Interstage turbine temperature (ITT) mea-surement is calibrated to provide a very ac-curate reading. This is done by a temperaturetrimmer located on top of the engine. Thistemperature trimmer is connected in parallelwith the ITT harness, and it is factory preset.

The temperature sensed by the thermocouplesis sent to gages (Figure 7-26) on the center in-strument panel calibrated in degrees Celsiusand designated ITT. On BB-1484, 1486 andsubsequent, the gages use DC power. Prior toBB-1486, excluding BB-1484, the gages areself-energizing and do not require DC power.

Engine Power (Torque)Figure 7-24 Propeller Control Lever

Figure 7-25. Friction Control Knobs

(Figure 7-27) on the engine instrument panelthat is calibrated in foot-pounds of torquetimes 100. The torquemeter chamber receivesa supply of oil at a relatively constant pressure

from the engine lubricating system. On BB-1484, 1486 and subsequent, the gages use DCpower. Prior to BB-1486, except 1484, it ispowered by the 26-volt AC bus.



B200 — BB-1484 1486 AND AFTER



but it can move a limited amount in axial di-

rection because of helical splines. Therefore,the first-stage ring gear is a reaction member that reacts to an increase or decrease of appliedtorque by moving aft as engine torque is in

Torque LimiterEngine torque is automatically limited to apreset value by a torque limiter that is suppliedwi th a to rque p ressure s igna l f rom thetorquemeter.

At a predetermined torque pressure of 2,368to 2,447 foot-pounds, the torque limiter willbleed off and change the pneumatic servo pres-

B200 BB 1484, 1486 AND AFTER

B200 — PRIOR TO BB-1486, EXCEPT 1484

200

Figure 7-26. ITT Gages


PRIOR TO BB-1486, EXCEPT 1484

Figure 7-27. Torque Gages

Propeller RPMPropeller rpm output is sent to a gage (Figure7-28) on the engine instrument panel cali-

brated directly in propeller revolutions per minute. On BB-1484, 1486 and subsequent,DC power is required. Prior to BB-1486, ex-cept BB-1484, these gages do not requireaircraft DC electrical pow er, as they are op-erated by tachgenerators.






Engine RPM (N1)Engine or gas generator (N1) rpm is also sentto a gage (Figure 7-29) on the engine in-strument panel. The gage is calibrated in

percentage of design 100% rpm. On BB-1484, 1486 and subsequent, DC power is re-quired. It does not require aircraft electrical

i BB 1484 i l di BB 1485

SYNCHROSCOPE

A synchroscope (Figure 7-30) with black andwhite cross patterns is located on the lower rightcorner of the pilot’s instrument panel to aid inmanual propeller synchronization. The disc willrotate in the direction of the higher rpm engine.The disc will stop rotating when the enginesare synchronized. Input signals to the synchro-scope are from the propeller tachgenerators.



Figure 7-28. Propeller RPM Gages


Figure 7-29. Engine RPM Gages



SYNCHROPHASING

General

Two synchrophasing systems are available. Theyare identified as Type I and Type II systems.

Type II System (BB-935 andSubsequent)

The Type II synchrophaser system is an

ControlThe system is controlled by a two-positionswitch (Figure 7-30) located on the lower right

side of the pilot’s instrument panel.

Operation Type II SystemTurning the control switch on will supply DCpower to the electronic control box. Input sig-nals representing propeller rpm are receivedfrom magnetic pickups on each propeller The



The Type II synchrophaser system is anelectronic system, certified for takeoff and

landing (Figure 7-31). It functions to matchthe rpm of both propellers and establish ablade phase relationship between the rightand left propellers to reduce cabin noise toa minimum.

The system can not reduce rpm of either pro-peller below the datum selected by the pro-peller control lever. Therefore, there is no

indicating light associated with the Type IIsystem.

from magnetic pickups on each propeller. Thecomputed input signals are corrected to a com-

mand signal and sent to an rpm trimming coillocated on the propeller governor of the s lowengine and its (propeller) rpm is adjusted tothat of the other propelle r.

NOTE

If the synchrophaser is on and failsto synchronize the propellers, turn itoff, then manually synchronize thepropellers and turn it back on.

PROP SYNCOFF

SYNCCONTROLLER

RPM SENSORRPM SENSOR

PROPELLER GOVERNOR (PRIMARY)

PROPELLER SPINNER

Type I System BB-2 throughBB-934

The Type I system uses the master-slave con-cept (Figure 7-32). The left propeller is themaster propeller and the right propeller is theslave. The system functions to adjust the rpmof the right propeller to that of the left, withina limited rpm range and at the same time it pro-vides a specific blade phase relationship be-tween the left and right propellers The overall

of the pilot’s instrument panel. Being a mas-ter slave system, it should be off during groundoperation, takeoff, and landing, because if themaster engine fails, the rpm of the slave en-

gine will decrease a limited amount. The pro-pellers should be manually synchronizedbefore turning the system on.

An amber light (Figure 7-34) on the caution/ advisory panel will come on if the synchro-nizer system is on and the landing gear is se-





tween the left and right propellers. The overalleffect of the synchrophaser system is to reduce

noise level in the cabin to a low va lue.

ControlSystem control is achieved by a two-positionswitch (Figure 7-33) on the lower right side

lected down.

Operation Type I SystemWhen the synchrophaser switch is on (Figure7-33), DC power is available to the control box.Input signals are received by the control box

SYNCHROSCOPE

SYNCCONTROLLER

RPM MONOPOLE RPMMONOPOLE

SLAVE

MSYNCACTUATOR

LDG GRUP

DOWNSYNC ON

MASTER

PROPELLER OVERSPEED GOVERNOR



forces tend to induce high blade angles or toward feather.

Feathering is normally accomplished with the

propeller control lever (Figure 7-22). Movingthis lever aft to the FEATHER position will me-chanically raise the governor pilot valve anddump oil from the propeller cylinder. The coun-terweights and springs will then rapidly feather the propeller.

Al if th i i h t d th d



from monopoles on each propeller overspeedgovernor. These signals represent propeller rpm. The pulse rate difference of the signalsis corrected to a command signal, which istransmitted to an actuator on the right engineprimary governor housing. The actuator, in

turn, trims the right propeller governor tomatch its rpm to the left (master) propeller.This adjustment does not affect the positionof the propeller control lever. When turned off,the stepping motor or actuator will run to a neu-tral position.

PROPELLERFEATHERING

h ll d C l ll i l

Also, if the engine is shut down on the groundusing the condition lever, the oil pressure de-

creases and the centrifugal force of the coun-terweights plus the springs will eventuallyfeather the propeller. However, this is not a rec-ommended procedure. The prop should befeathered with the prop control lever.

AUTOFEATHERING

An autofeather system is available in theevent of engine failure. This system willrapidly feather the affected propeller byopening a solenoid valve on the overspeedgovernor and will dump propeller controloil. The counterweights and springs willrapidly feather the propell er.

ControlAutofeather is controlled by a single switch(Figure 7-35) for both propellers. The switchis marked ARM, OFF, and TEST.

ArmingTurning the switch to the ARM position ap-plies power to a microswitch in each power lever quadrant. The switches will close when

the power levers are advanced to a position thatshould produce approximately 90% N1 rpm.

Figure 7-34. Sync Light

Figure 7-33. Propeller Synchroscopeand Switch (Type I)

When this occurs, electrical power is finallytransmitted to torque switches.

Once engine torque is over 400 foot-pounds, the

opposite engine’s autofeather annunciator willilluminate.

IndicationTwo green lights (Figure 7-36) on the cau-tion/advisory panel marked L and R AUTO

OPERATING PRINCIPLES

Assume that the autofeather system is armedfor takeoff. As the power levers are advanced,

the microswitches will close at a position in thequadrant representing 90% N1 rpm. Electricalpower will now be applied to engine torque-sensitive switches (two for each engine). Oneswitch on each engine is set to open at ap-proximately 200 foot-pounds of torque andthe second switch on each engine opens at 400foot pounds of torque When passing through





/ y pFEATHER will illuminat e if the autofeather

system is armed, the power levers are ad-vanced to approximately 90% N1 rpm or greater, and the engines are developing power in excess of 400 foot-pounds of torque.

TestingThe TEST position of the autofeather systemi s u se d t o b y p a s s t h e p o w e r l e v e r m i -

croswitches and induce arming at a muchlower power setting to test the integrity of the torque switches, the arming relays, thedump solenoid valve, and the arming lightswithout high power settings. The autofeather system is designed for use only during crit-ical power periods such as takeoff, approach,and landing, and it should be turned off under all other operating conditions.

foot-pounds of torque. When passing through90% N1, a green AUTOFEATHER light for

each engine should be on, indicating a fullyarmed condition for both engines.

AUTOFEATHERING

If an engine fails (Figure 7-37) (for example,during takeoff), a torque switch will closewhen torque decays to 400 foot-pounds and the

AUTOFEATHER light of the operating en-gine will extinguish, indicating that its auto-feather circuit is disarmed. Then as torque onthe failing engine decays to 200 foot-pounds,a second torque switch closes. The armingrelay will be energized, and the dump valvelocated on the overspeed governor will opento dump propeller servo oil and produce rapidfeathering. In addition, the autofeather light

for the failed engine will extinguish.



NC

DUMPVALVE

TORQUESWITCH

TORQUESWITCH

200

400

LEFTPOWERLEVER

SWITCH

ARMINGRELAY



Autofeather TestThe TEST position (Figure 7-38) of the AUTO-FEATHER switch bypasses the power lever

90% N1 switches. With both engines set to ap-proximately 500 foot-pounds of torque, mov-ing the switch to the TEST position and reducingpower slowly on one engine, the opposite en-

NOTE

If the condition levers are not set atLOW IDLE, it may not be possibleto reduce torque below 200 foot-pounds, which would result in thepropeller not cycling during test.

L AUTOFEATHERL AUTOFEATHER

R AUTOFEATHERR AUTOFEATHER

NC

DUMPVALVE

TORQUESWITCH

TORQUESWITCH

400

200

TEST

C/B

AUTOFEATHER

ARM

RIGHTPOWERLEVER

SWITCH* CLOSED AT HIGH N1

OFFAUTOFEATHER

LIGHTS

ARMINGRELAY

Figure 7-37. Autofeather System Schematic (Both Power Levers at Approximately90% N1; Right Engine has Failed)



NC

DUMPVALVE

TORQUESWITCH

TORQUESWITCH

LEFTPOWERLEVER

SWITCH

ARMINGRELAY400

200



UNFEATHERING

With the prop levers set full forward, pro-peller unfeathering occurs automatically with

oil pressure as the engine is started and theblade angle will decrease to the datum set bythe Beta/reverse mechanism (approximately18°) A h f h i i

LIMITATIONS(POWERPLANT)

GENERAL

The limitations contained in Section II of the

L AUTOFEATHERL AUTOFEATHER

R AUTOFEATHERR AUTOFEATHER

NC

DUMPVALVE

TORQUESWITCH

TORQUESWITCH

400

200

TEST

C/B

AUTOFEATHER

ARM

RIGHTPOWERLEVER

SWITCH* CLOSED AT HIGH N1

OFF AUTOFEATHERLIGHTS

Figure 7-38. Autofeather Test Schematic (Left Power Lever Below 200 ft-lb;Right Power Lever Above 400 ft-lb)



POWERPLANT

Manufacturer: Pratt & Whitney Aircraft of Canada LTD, Engine Model No. PT6A-41/42.

ENGINE OPERATING LIMITS

The following limitations in Tables 7-1, 7-2and 7-3 shall be observed. Each column pre-

sents limitations. The limits presented do notnecessarily occur simultaneously. Refer to Pratt & Whitney Maintenance Manual for spe-cific actions required if limits are exceeded.

APPROVED FUELS

See Chapter 5, FUEL SYSTEM.

Table 7-1. ENGINE OPERATING LIMITS (PT6A-42 ENGINE BB-1439, 1444



OPERATING TORQUE MAXIMUM GAS GENERATOR PROP OIL OIL

CONDITION SHP FT-LB OBSERVED RPM N1 RPM PRESS. TEMP

(1) ITT °C RPM % NP PSI (2) °C (3) (4)

STARTING --- --- 1,000 (5) --- --- --- --- -40 (min)

LOW IDLE --- --- 750 (6) 22,875 61 (min) (13) 60 (min) -40 to 99

HIGH IDLE --- --- --- --- (7) --- --- -40 to 99

TAKEOFF ANDMAX CONT 850 2,230 800 38,100 101.5 2,000 100 to 135 0 to 99

MAX CRUISE 850 2,230 (8) 800 38,100 101.5 2,000 100 to 135 0 to 99

CRUISE CLIMB ANDREC (NORMAL) CRUISE 850 2,230 (8) 770 38,100 101.5 2,000 100 to 135 0 to 99

MAX REVERSE (9) 850 --- 750 --- 88 1,900 100 to 135 0 to 99

TRANSIENT --- 2,750 (5) 850 38,500 (10) 102.6 (10) 2,200 (5) --- 0 to 104 (11)

(AND SUBSEQUENT)



(1) ITT °C RPM % NP PSI (2) °C (3) (4)

STARTING --- --- 1,000 (5) --- --- --- --- -40 (min)

LOW IDLE --- --- 750 (6) 21,000 56 (min) --- 60 (min) -40 to 99HIGH IDLE --- --- --- --- (7) --- --- -40 to 99

TAKEOFF AND

Table 7-2. ENGINE OPERATING LIMITS (PT6A-42 ENGINE PRIOR TO BB-1439,1444 AND SUBSEQUENT)





(1) ITT °C RPM % NP PSI (2) °C (3) (4)

STARTING --- --- 1,000 (5) --- --- --- --- -40 (min)

LOW IDLE --- --- 660 (6) 19.500 52 (min) --- 60 (min) -40 to 99

HIGH IDLE --- --- --- --- (7) --- --- -40 to 99

TAKEOFF (12) 850 2,230 750 38,100 101.5 2,000 105 to 135 10 to 99

Table 7-3. ENGINE OPERATING LIMITS (PT6A-41 ENGINE)



MAX CONT AND

MAX CRUISE 850 2,230 (8) 750 38,100 101.5 2,000 105 to 135 10 to 99CRUISE CLIMB ANDREC CRUISE 850 2,230 (8) 725 38,100 101.5 2,000 105 to 135 0 to 99

MAX REVERSE (9) --- --- 750 --- 88 1,900 105 to 135 0 to 99

TRANSIENT --- 2,750 (5) 850 38,500 (10) 102.6 (10) 2,200 (5) --- 0 to 104 (11)

FOOTNOTES:

1. Torque limit app lies within range of 1 ,600-2 ,000 prope ller rpm (N2). Below 1,600 propeller rpm torque is limited to 1,100 ft-lbs.

2 . When gas genera tor speeds are above 27,000 rpm (72% N

1

) and oil temperatures are between 60°C and 71°C, normal oil pres-

sures are: 105 to 135 psi below 21,000 feet; 85 to 135 psi at 21,000 feet and above.

During extremely cold starts, oil pressure may r each 200 psi. Oil pressure between 60 and 85 psi is undesirable; it should be tolerated only forthe completion of the flight, and then only at a reduced power setting not exceeding 1,100 ft-lbs torque. Oil pressure below 60 psi is unsafe; itrequires that either the engine be shut down, or that a landing be made at the nearest suitable airport, using the minimum power required tosustain flight. Fluctuations of ± 10 psi are acceptable.

3. A minimum oil temperature of 55°C is recommended for fuel heater operation at takeoff 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 . T he se va lu es are t ime l im ited to fi ve se co nd s.

6. High ITT at ground id le may be corrected by reducing accessory load or increasing N1 rpm.7. At approximately 70% N1.

8 . Cru ise to rque values vary with al t itude and tempera ture .

9 . T his o pe ra tion i s t ime l im ited to on e m in ute.

10. These values are time l imited to 10 seconds.

11. Values above 99°C are t ime limited to 10 minutes .

12. These values are time l imited to five minutes .

13. 1,100 rpm for McCauley Propeller, 1,180 rpm for Hartzell Propeller.

PROPELLER

Manufacturer:

• Prior to BB-1193 and BL-37Hartzell Propeller, Inc.Diameter 98.5 inches

• BB-1193 through 1438,BB-1440 through 1443,BL-37 through 138McCauley PropellerDiameter 98.0 inches

Sustained propeller overspeeds faster than2,000 rpm indicate failure of the primarygovernor. Flight may be continued at pro-peller overspeeds up to 2,120 rpm (2,080

rpm prior to BB-1444, except 1439) pro-vided torque is limited to 1,800 foot-pounds.Sustained propeller overspeeds faster than2,120 rpm (or 2,080 as indicated above) in-dicate failure of both the primary governor and the secondary governor, and such over-speeds are unapproved.





• BB-1439, 1444 through 1508,McCauley PropellerDiameter 94.0 inches

• BB-1509 and subsequentHartzell PropellerDiameter 93.0 inchesRotational Speed Limits

Rotational Speed Limits:• 2,200 rpm (Transient)—Not exceedingfive seconds

1,900 rpm—Reverse

2,000 rpm—All other conditions

Propeller Rotational Overspeed

LimitsThe maximum propeller overspeed limit is2,200 rpm and is time-limited to five seconds.

POWERPLANT INSTRUMENTMARKINGS

The powerplant instrument markings are givenin Table 7-4.

STARTER LIMITS

Use of the starter is limited to:

40 seconds ...................................... ON60 seconds .................................... OFF

Then, if necessary:

40 seconds ...................................... ON60 seconds .................................... OFF

Then, if necessary:

40 seconds ...................................... ON30 minutes .................................... OFF



Table 7-4. POWERPLANT INSTRUMENT MARKINGS

RED LINE YELLOW ARC GREEN ARC RED LINEINSTRUMENT MINIMUM CAUTION NORMAL MAXIMUM

LIMIT RANGE OPERATING LIMIT

INTERSTAGE TURBINETEMPERATURE (ITT) * --- --- 400°C to 800°C 800°C

TORQUEMETER --- --- 0 to 2,230 ft-lb 2,230 ft-lb

PROPELLER

BB-1484, 1486 AND SUBSEQUENT



PROPELLERTACHOMETER (N2) --- --- *** 2,000 rpm

GAS GENERATORTACHOMETER (N1) --- --- 61 to 101.5% 101.5%

OIL TEMPERATURE --- --- 0°C to 99°C 99°C

OIL PRESSURE ** 60 psi 60 to 100 psi 85 psi to 135 psi 135 psi

BB-2 THROUGH 1485, EXCEPT 1484

RED LINE YELLOW ARC GREEN ARC RED LINE

INSTRUMENT MINIMUM CAUTION NORMAL MAXIMUM


INTERSTAGE TURBINETEMPERATURE (ITT) * --- --- 400°C to 800°C 800°C

TORQUEMETER --- --- 400 ft-lb to 2,230 ft-lb 2,230 ft-lb

PROPELLERTACHOMETER (N2) --- --- 1,600 rpm to 2,000 rpm 2,000 rpm

GAS GENERATORTACHOMETER (N1) --- --- --- 101.5%

OIL TEMPERATURE --- --- 10°C to 99°C 99°C

OIL PRESSURE ** 60 psi --- 100 psi to 135 psi 200 psi

PT6A-41 ENGINE

RED LINE YELLOW ARC GREEN ARC RED LINE

INSTRUMENT MINIMUM CAUTION NORMAL MAXIMUM


INTERSTAGE TURBINE

1. The PT6A engine power section con-sists of:

A. One compression stage and four tur-bine stages.

B. A two-stage reaction turbine.

C. A two-stage turbine and a centrifu-gal compressor.

D. Twin-spool, two-stage turbines.

4. During ground operation at LO IDLE,you note that ITT is exceeding 750°C.Which of the following actions wouldyou consider best to reduce ITT?

A. Move the propeller control lever tothe low rpm position

B. Reduce accessory load or increaseN1 rpm



QUESTIONS



p , g

2. The function of the reduction gear systemis to provide gear reduction:

A. For the propeller

B. Between the compressor and thepower turbine

C. For the airplane’s accessory drivesection

D. Between the compressor and the com-pressor turbine

3. If a chip detector light illuminates, youmust do one of the following:

A. Continue normal flight operationsand have the filter checked after landing.

B. Reduce torque to 500 foot-poundsfor the remainder of the flight.

C. Check engine instruments and, if normal, no action is required.

D. Shut the engine down and land as soonas practical.

C. Move the power lever into the

ground fine (Beta)/reverse rangeD. Shut down and have the propeller LO

IDLE stops checked

5. When using maximum reverse power atHI IDLE and full increase rpm, you wouldexpect a maximum propeller rpm of:

A. 1,900

B. 2,200

C. 1,830

D. 2,000

6. D ur in g a g ro un d st ar t of th e ri gh tengine, the IGNITION ON light shouldilluminate:

A. At 10% N1 rpm.B. When the condition lever is moved

to LO IDLE.

C. At a stabilized 12% N1.D. When the start switch is moved to

the IGNITION and ENGINE STARTposition.

7. When the AUTO-IGNITION switch is inthe ARM position, ignition is:

A. Continuous.

B. Inactive but armed if torque isgreater than 400 foot-pounds.

C. Controlled by the stall warningsystem.

D. Continuous when torque is greater than 400 foot-pounds.

8. After lift-off, if an autofeather is initiated,

10. Which o f the fol lowing i s the mos taccurate definition of Engine TorqueReadout?

A. Power developed by the gas genera-tor

B. Thrust supplied by the propeller

C. Ratio of compressor inlet to exhaustoutlet

D. Power delivered to the propeller





8. After lift off, if an autofeather is initiated,

the immediate requirement is to:A. Continue to fly the airplane and

allow the propeller to feather andstop.

B. Move the power lever to idle.

C. Move the condition lever to cutoff.

D. Reduce electrical loads.

9. As you start taxiing, the OAT is +38°C.You notice that the right engine N1 rpmis decreasing and ITT is rising rather fast.Which of the following is the best actionto take?

A. Turn off the refrigerant compressor and adjust the condition lever to ahigher N1

B. Move the condition lever to LOIDLE

C. Advance the power lever to increaseN1 rpm

D. Move the propeller lever to FULLDECREASE RPM

CHAPTER 8

FIRE PROTECTION

CONTENTS

Page

INTRODUCTION ................................................................................................................... 8-1

FIRE DETECTION ................................................................................................................. 8-1





General ............................................................................................................................. 8-1

Indicators.......................................................................................................................... 8-4

FIRE EXTINGUISHING ........................................................................................................ 8-4

General ............................................................................................................................. 8-4

Controls, Indicators, and Operation ................................................................................. 8-4Limitations ....................................................................................................................... 8-4

TESTING OF THE SYSTEMS............................................................................................... 8-6

PORTABLE FIRE EXTINGUISHERS................................................................................... 8-7

QUESTIONS........................................................................................................................... 8-8

ILLUSTRATIONS

Figure Title Page

8-1 Fire Detection System—BB-1439, 1444 and After ................................................. 8-2

8-2 Fire Detection System—BB-2 through 1443 Except 1439...................................... 8-3

8-3 Fire-Extinguishing System....................................................................................... 8-5

8-4 Gage Location .......................................................................................................... 8-6

8 5 Portable Fire E ting isher 8 7





8-5 Portable Fire Extinguisher........................................................................................ 8-7

TABLES

Table Title Page

8-1 Temperature vs. Pressure Data ................................................................................. 8-6

FIRE

CHAPTER 8FIRE PROTECTION





INTRODUCTION

The two engines each have independently operating fire-detection systems. A temper-ature-sensing cable or three flame detectors per engine (operating through an amplifier)turn on the appropriate warning light. Separate fire-extinguishing systems are availableas an option. Crew activation is required to release the extinguishing chemical agent intothe nacelle with the fire.

FIRE DETECTIONGENERAL

FIRE PULL

WARN



L LEXT DET

R R FIRE SENSORELEMENT

PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

R ENG FIRE

PUSHTOEXT

OKD

LENG FIRE

PUSHTO EXT

OKD

DETAIL A



A

B

OFFTEST SWITCH

ENG FIRE SYS

DETAIL B(WITH FIRE

EXTINGUISHER)

L

DETR

OFF

TEST SWITCH

ENG FIRE SYS

DETAIL B(WITHOUT FIRE

EXTINGUISHER)

B

C

A

28 VDC

TEST SWITCH

SENSOR RESPONDER

SIMPLIFIED CIRCUIT RESPONDER ALARM

SWITCH (N.O.)

SENSOR ELEMENT

ISOLATOR

SENSOR ELEMENTSENSORRESPONDER

FIRE SENSORELEMENT

C

PRINTEDCIRCUITCARDS



PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

RENG FIRE

PUSHTOEXT

OKD

LENG FIREPUSH TOEXT

OKD

PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

RENG FIRE

PUSHTO EXT

OKD

LENG FIRE

PUSH TOEXT

OKD

APPROACH

PLATE

BRT

PRESS

TO

TEST

AIRPLANES WITH 12-STATION WARNING ANNUNCIATOR PANEL

AIRPLANES WITH 20 STATION WARNING ANNUNCIATOR PANEL



L 3EXT

DET

R2

OFF

TEST SWITCH

B

A

FIRE DETECTORS

CONTROLAMPLIFIERS

FIRE DETECTORS

AIRPLANES WITH 20-STATION WARNING ANNUNCIATOR PANEL

1

DETAIL A

tor system operates at a high preset thresholdlevel, but occasionally the system may be se toff by sunlight if it enters nacelle openings atthe appropriate angle to reach the detectors.

Power is supplied from the No. 1 dual-fed busthrough a circuit breaker on the right sidepanel.

INDICATORS

When the temperature-sensing cable is acti-vated or the light threshold is reached (prior to BB-1444, except 1439) indicating a possi-

CONTROLS, INDICATORS, ANDOPERATION

A three-lens control indicator is located on the

glareshield when the optional extinguisher system is incorporated (Figure 8-3). The threelenses are:

• Red—L (or R) ENG FIRE PUSH TOEXT

• Amber—D

• Green—OK





to BB 1444, except 1439) indicating a possi

ble fire, the appropriate light on the warningannunciator panel comes on. Assuming theintegrity of the wiring or sensor cable has notbeen compromised and the fire goes out, thelight will extinguish. Both systems can againdetect the outbreak of fire.

With the fire-extinguishing system installed,fire warning is indicated by the L or R ENG

FIRE PUSH TO EXT switchlights located onthe glareshield at each end of the warning an-nunciator. Fire warning is also simultaneouslyindicated by the red warning annunciators.

FIRE EXTINGUISHING

GENERALFire in either engine compartment is smoth-ered by engulfing the nacelle compartmentwith bromotrifluoromethane (CBrF3) pres-surized with dry nitrogen. There are threespray bars per engine compartment (Figure8-3), each one supplied by one common fireextinguisher supply cylinder per engine. One

squib per bottle incorporates a pyrotechniccartridge which releases the entire contents.The squib is fired by depressing the switch-

The red L (or R) ENG FIRE PUSH TO EXTlens indicates a detected fire. The three-lenscontrol indicator is pushed to activate the ap-propriate extinguisher.

The amber D lens indicates that the extin-guisher has been discharged, and the supplycylinder is empty.

The green OK lens confirms circuit continu-ity during the test function.

When a red warning light indicates a fire andit is confirmed by the pilot, the appropriate (Lor R) engine should be shut down and the fire-extinguishing switchlight depressed. This firesthe appropriate squib, releasing the contents

through the tubing. When the bottle is dis-charged, the amber D light illuminates.

The pressure gages, one located on each fire-extinguishing supply cylinder, reflect the con-tents of the bottle. They can be read only whileon the ground because they are located in thewheel wells. See Figure 8-4 and Table 8-1 for temperatures vs. pressure data and for the

gage location.

LIMITATIONS



PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

PRESS TO RESET

MASTER

WARNING

PRESS TO RESET

MASTER

CAUTION

RENG FIRE

PUSH TOEXT

OKD

LENG FIRE

PUSHTOEXT

OKD

DETAIL A

C



L LEXT DET

R R

OFF

TEST SWITCH

EXPLOSIVESQUIB

FIRE EXTINGUISHERSUPPLY CYLINDER

C

R MONITOR

MODULE

L MONITORMODULE

A

B

SUPER KING AIR 200/B200 PILOT TRAINING MANUALFlightSafety

international



Each engine has its own self-contained extin-guishing system, which can be used only oncebetween recharging. This system cannot be usedto extinguish a fire in the opposite engine.

TESTING OF THE

For BB-1439, 1444 and subsequent, when theswitch is placed in the DET L or DET R po-sition, the illumination of the correspondingENG FIRE light assures the integrity of the

cable and continuity of the electrical wiring.

Prior to BB-1444, except 1439, the four-posi-

Figure 8-4. Gage Location

TEMPERATURES -40°/-40° -29°/-20° -18°/0° -6°/20° 4°/40° 16°/60° 27°/80° 38°/100° 48°/120°

°C/°F

PSI MINIMUM 190 220 250 290 340 390 455 525 605

to to to to to to to to to

PSI MAXIMUM 240 275 315 365 420 480 550 635 730

Table 8-1. TEMPERATURE VS. PRESSURE DATA

NOTE: PRESSURES ARE EXTRACTED FROM THE BEST AVAILABLE INFORMATION AND SHOULD ONLY BE USED AS A GUIDE.

During testing, the pilot’s and copilot’s redMASTER WARNING light flashes, and, if the optional extinguisher system is installed,the red lenses placarded L ENG FIRE–PUSH

TO EXT and R ENGINE FIRE–PUSH TOEXT illuminate. Failure of the fire detectionannunciators in any of the test positions indi-cates a malfunction in that system. When thelight fails to come on during testing, a no-gosituation exists. Should there be no responsein any position, check the circuit breaker.

For testing the extinguishing systems, the cir-

SUPER KING AIR 200/B200 PILOT TRAINING MANUALFlightSafety

international



cuitry of the squibs is checked for continuityby rotating the TEST SWITCH FIRE DETand FIRE EXT through the two (LEFT andRIGHT) EXT positions (Figure 8-3). Theamber D light and the green OK light shouldilluminate, indicating that the bottle charge de-tector circuitry and squib-firing circuitry areoperational and that the squib is in place(Figure 8-3).

PORTABLE FIREEXTINGUISHERS

There are two portable fire extinguishers in-side the airplane. One is in the cabin, the other is in the cockpit. One is normally installed on

the floor on the left side of the airplane for-ward of the airstair entrance door, just aft of the rearmost seat; the other is underneath thecopilot’s seat (Figure 8-5).

Figure 8-5. Portable Fire Extinguisher

1. H ow ma ny ti me s c a n t he fi re -e xt in -

guishing system be fired between supplycylinder recharges, per engine?

A. One

B. Two

C. Three

D. Four

3. The fire detection system is tested by the

flight crew using the TEST SWITCH.The switch:

A. Supplies an electrical signal similar to the one that the detectors send to thewarning annunciating system.

B. Heats up an infrared source by eachdetector.

C. Merely checks the annunciator system



QUESTIONS



2. The amber D light, when illuminated (other than for test purposes), indicates:

A. The supply cylinder is full.

B. The supply cylinder is empty.

C. T h e s up p l y cy l i nd e r is b e i n gdischarged.

D. The supply cylinder is available for discharge.

operation.D. Directs a small amount of bleed air to

heat the detectors.

4. I n t he te s ti ng mo de , if th e T E STSWITCH is in either LEFT or RIGHTEXT position, the green OK light failsto i l luminate, but the amber D does

illuminate, what does this mean?A. The bottles are empty.

B. The lights are definitely burned out.

C. The generators are not powering thesupply bus.

D. The squib-firing circuitry may notwork.

CHAPTER 9

PNEUMATICS

CONTENTS

Page

INTRODUCTION ................................................................................................................... 9-1

GENERAL............................................................................................................................... 9-1





System Description and Location .................................................................................... 9-1

Bleed-Air Warning System.............................................................................................. 9-3

Bleed-Air Control ............................................................................................................ 9-5

Door Seal System............................................................................................................. 9-7

Flight Hourmeter.............................................................................................................. 9-7

LIMITATIONS ........................................................................................................................ 9-7

QUESTIONS........................................................................................................................... 9-8

ILLUSTRATIONS

Figure Title Page

9-1 Pneumatic and Vacuum Systems Diagram............................................................... 9-2

9-2 Bleed-Air Ejector ..................................................................................................... 9-3

9-3 Suction Gage and Pressure Gage.............................................................................. 9-3

9-4 Bleed-Air Warning System Diagram....................................................................... 9-4

9-5 L & R BL AIR FAIL Warning Lights ...................................................................... 9-5





9-6 BLEED AIR VALVE Switches ................................................................................ 9-5

9-7 Pneumatic Plastic Tubing ......................................................................................... 9-5

9-8 Bleed-Air Control Diagram...................................................................................... 9-6

9-9 Cabin Door Air Seal................................................................................................. 9-7

9-10 Hourmeter................................................................................................................. 9-7

L R

COBLEED AIR

CHAPTER 9PNEUMATICS





INTRODUCTION

The Super King Air utilizes an engine bleed-air pneumatic system to provide bleed air for the door system (door seal line), the ice protection systems (surface deice), thebleed-air warning system, the rudder boost, the hourmeter, and the brake deice system.Also, pneumatic air that is exhausted overboard via a venturi creates a negative pres-sure that is used by the vacuum system.

GENERALSYSTEM DESCRIPTION AND

VALV E

5 1 5

2 0

AIR



LEGENDHIGH PRESSURE BLEED AIR

REGULATED BLEED AIR

VACUUM

TODEICEBOOTS

EXHAUSTOVERBOARDEJECTOR

LANDING GEARRESERVOIR

(HYDRAULIC

PNEUMATIC PRESSUREGAGE

(IN COCKPIT)

PRESSURESWITCHRIGHT

SQUATSWITCH

DEICE

DISTRIBUTOR

VALVE

HOURS

FLIGHT

1/10



VACUUMREGULATOR

RIGHT BLEED-AIR WARNING SYSTEM

RIGHTP AIR

LEFT BLEED-AIR WARNING SYSTEM

LEFTP AIR

AIRSTAIRDOOR SEAL

LINE

(HYDRAULICGEAR ONLY)

LEFTSQUAT

SWITCH

CLOSED ON

GROUND(NO)

CHECK VALVECHECK VALVE

PNEUMATICAIR VALVE

(NO)

PNEUMATICAIR VALVE

(NO)

18 PSIPRESSURE

REGULATOR

PRESSURATIONCONTROLLER,OUTFLOW AND

SAFETY VALVES

GYROINSTRUMENTS

GYROSUCTION

(IN COCKPIT)

60 PSID

P SWITCH

RUDDERBOOSTSYSTEM

VALVEL SERVO

R SERVO

RUDDER BOOST RUDDER BOOST

LEFTNC

RIGHTNC

15 PSIREGULATOR

4 PSIREGULATOR



troller, and surface deice originates througha venturi (bleed air ejector) which is exhaustedoverboard (Figure 9-2). One engine can sup-ply sufficient bleed air for all associated sys-

tems. In addition, the brake deice systemreceives bleed air that is tapped off down-stream of each instrument air valve (Figure 9-1). Refer to Chapter 10, ICE AND RAINPROTECTION, for more information on thebrake deice system. Refer to Chapter 15,FLIGHT CONTROLS, for information on therudder boost system.

Engine bleed air is ducted from each engine toit ti L R fl t l it t d

and the flight hourmeter. An ejector changespressure to a vacuum to operate gyro instru-ments, pressurization controller, and outflowand safety valves. The flow control unit reg-

ulates the mixture of engine bleed air for pres-surization with ambient air. Pressurization air is routed through the wings and, finally, intothe cabin where it is used for heating, cooling,and pressurization.

A suction gage (Figure 9-3), which is cali-brated in inches of mercury and is located onthe copilot’s right subpanel, indicates gyro

suction To the right of the suction gage is ati (Fi 9 3) hi h



Engine bleed air is ducted from each engine toits respective L or R flow control unit mountedon the firewall. A pressure supply line tees off the engine bleed-air line forward of the fire-wall and flow control unit. This supply line con-tains pneumatic pressure to operate the surfacedeicer, rudder boost, door seal, brake deice(hot brakes) system hydraulic reservoir (BB-1193 and after, including BB-1158 and 1167),

suction. To the right of the suction gage is apneumatic pressure gage (Figure 9-3) whichindicates air pressure available to the deice dis-tributor valve, vacuum system, bleed air warn-ing, rudder boost, hourmeter, and door seal.The pneumatic pressure gage is calibrated inpounds per square inch (psi).

BLEED-AIR WARNING SYSTEMThe bleed-air warning system is installed toalert the pilot when a pressurization line or pneumatic line ruptures, exhausting hot enginebleed air into the airframe.

Whenever the temperature from this rupturereaches approximately 204°F (Figure 9-4),the plastic tubing melts, which results in theillumination of either the L BL AIR FAIL or the R BL AIR FAIL warning lights (Figure 9-5). A severe bleed-air leak could result in a de-

Figure 9-2. Bleed-Air Ejector



ENGINE P3

BLEED-AIRCONNECTOR

ENGINE P3

BLEED-AIRCONNECTOR

AMBIENTAIR

AMBIENTAIR

PLUGS

BLEED-AIRWARNINGSWITCHES

ENVIRONMENTALMIXING PLENUM

FLOW CONTROLVALVE



UNREGULATED PNEUMATIC BLEED AIR

REGULATED PNEUMATIC BLEED AIR

BLEED AIR WARNING LINE

LEGEND

ENVIRONMENTAL BLEED AIR

VACUUM

FIREWALLFIREWALL

PLUGS PLUG

MANIFOLD(18 PSI REGULATOR)

RH N.O. PNEUMATICAIR VALVE

AIR-TO-AIRHEAT EXHCHANGER

BLEED-AIRBYPASS VALVE

BLEED-AIRBYPASS VALVE

AIR INLET AIR INLET

REAR SPAR

LH N.O. PNEUMATICAIR VALVE

EJECTOR

LH ENGINE

BLEED AIR INLET

AIR SOURCE—LH

BLEED AIR WARNING



crease in engine torque and an increase in ITT.Therefore, whenever the applicable BLEEDAIR VALVE switch (Figure 9-6) is placed intoINST and ENVIR OFF position, the pilot

should monitor the engine instruments for an

Figure 9-5. L & R BL AIR FAIL WarningLights

Figure 9-7. Pneumatic Plastic Tubing



should monitor the engine instruments for anincrease in torque and a decrease in ITT. Thisindicates that the leak has been isolated, if itwas a severe leak.

However, regardless of engine instruments,any time the bleed-air warning light illumi-nates, the respective bleed-air valve must bepos i t ioned in the INSTrument and

ENVIRonmental OFF position.

The plastic tubing (Figure 9-7) lies alongside

below the copilot’s feet. When this switch(one of two switches) closes, the applicable BLAIR FAIL light illuminates.

NOTE

The bleed-air warning annunciator will not extinguish after closing thebleed-air valves. When the bleed-air control switch is in the OPEN posi-tion, it requires DC power to open theflow control unit shutoff valve. Whenthe switch is in the INST & ENVIROFF position, it requires DC power to close the pneumatic instrumentair valve. Both positions receive their power from the bleed-air control CB.

BLEED-AIR CONTROL

Bleed air entering the cabin, used for pressur-ization and environmental functions, is con-trolled by the two BLEED AIR VALVESswitches which are marked OPEN, ENVIR OFF,and INST & ENVIR OFF. When the switch is

in the OPEN position, both the environmentalflow control unit and the pneumatic instrumentair valve open. When the switch is in the ENVIR

Figure 9-6. BLEED AIR VALVE Switches

9 - 6

F OR T RAI N

I N

S UP E R

KI N G

A I

NC

NC

NO

NO

R BL AIR OFFR BL AIR OFF

L BL AIR OFFL BL AIR OFF

BLEED AIR VALVESOPEN

INSTR & ENVIR OFF

ENVIROFF


INSTR & ENVIR OFF

ENVIROFF

LH FLOWCONTROLSHUTOFF

LHPNEUMATICBLEED AIRSHUTOFF

RH FLOWCONTROLSHUTOFF

RHPNEUMATIC

BLEED AIRSHUTOFF

PNEUMATIC ANDENVIRO OFF

PNEUMATIC ANDENVIRO OFF

ENVIROOFF

ENVIROOFF

OPEN

OPEN



G P URP O S E S

ONL Y

IR

2 0 0 / B 2 0 0

P I L O T T R A I NI N G MA N U A L


i n

t er n a t i on al

NC

NO

NO

NO

NO

6-8 SEC

NO

CABINPRESSDUMP

TEST

PRESS

SHUTOFF

RAM AIRDOORSOLENOID

CABINPRESETSOLENOID

CABINPRESSURESAFETYVALVE

DOORSEALSOLENOID

TIMEDELAYPCB

TEST

PRESS

DUMP

LH GEARSAFETYSWITCH

RH AMBIENTAIR SHUTOFFVALVE

LH AMBIENTAIR SHUTOFFVALVE

RH GEAR SAFETY SWITCH

CABIN AIR TEMP

UP

UP

DN

DN5A

DUALFED

BUS NO. 1

DUALFED

BUS NO. 2

Figure 9-8. Bleed-Air Control Diagram

DOOR SEAL SYSTEM

The entrance door to the cabin utilizes air from the pneumatic system to inflate the door seal (Figure 9-9) after the airplane lifts off.Bleed air is tapped off the manifold down-stream of the 18-psi pressure regulator. After the tap, the regulated air passes through a 4-psi regulator and to the normally-open valvethat is controlled by the left landing gear safetyswitch.

LIMITATIONS



R

1000

FLIGHT

HOURS 1/10

500

0

1000

SUPPLY PREMADE IN U

OXYG

Figure 9-10. Hourmeter



FLIGHT HOURMETER

The FLIGHT hourmeter (Figure 9-10) pro-

vides a readout of the airplane’s flight time.The meter is located on the copilot’s rightsubpanel. In order for it to operate, pneumaticbleed air must be supplied, and DC power must be available through the flap control cir-cuit breaker. In addition, weight must be re-moved from the right landing gear strut toaffect the squat switch.

LIMITATIONSThe pneumatic system limitations are as follows:

• Pneumatic gage indicates, within a greenarc, the normal operating range of 12 to20 psi, and the maximum operating limit(red line) of 20 psi.

• Vacuum (suction) gage indicates, withina narrow green arc, the normal suctionfrom 15,000 to 30,000 feet MSL of 3.0to 4.3 in. Hg, or from 15,000 to 35,000feet MSL of 2.8 to 4.3 in. Hg. A widegreen arc indicates the normal vacuumrange from sea level to 15,000 feet MSLof 4.3 to 5.9 in. Hg.

Figure 9-9. Cabin Door Air Seal

1. To what systems does the pneumatic sys-

tem supply bleed air?A. Electrical and hydraulics

B. Air data computer

C. Vacuum, hourmeter, hot brakes, door seal, surface deice, rudder boost, andhydraulic reservoir (if installed)

D. W in d shi e l d , r a d ia n t h e a t , f l i gh tcontrols

5. What are the engine instruments moni-

tored for after a bleed-air warning lighthas illuminated when a BLEED AIRVALVE switch is place d in INST &ENVIR OFF?

A. Increase in torque, decrease in ITT

B. Increase in ITT, decrease in torque

C. Steady N1 rpm, decrease in ITT

D. Increase in N1 rpm, decrease in ITT



QUESTIONS



2. Where does the negative pressure for thevacuum system originate?

A. 18 psi regulator

B. Pneumatic bleed-air venturi

C. Refrigerant compressor

D. Safety/dump valve

3. Approximately what temperature wi llcause the plastic pneumatic tubing of thebleed-air warning system to fail?

A. 306°F

B. 406°F

C. 260°F

D. 204°F

4. A b le ed -a ir l ea k co ul d re su lt i n adecrease in “__________” and an in-crease in “__________”

A. Engine torque, N1B. Engine rpm, ITT

C. Engine temperature, N1

D. Engine torque, ITT

6. What lights illuminate when there is afailure in the pneumatic system?

A. Either L BL AIR FAIL or R BL AIRFAIL only

B. Either L BL AIR FAIL or R BL AIRFAIL or both

C. BLEED AIR FAIL

D. R FAIL or L FAIL

7. What is the maximum operating pressurelimit of the pneumatic system?

A. 12-20 psi

B. 18 psi

C. 6 psi

D. 20 psi

8. From sea level to 15,000 feet MSL, whatis the normal vacuum range of the vacuumsystem?

A. 3.0-4.3 in. Hg

B. 3.0-4.3 psi

C. 4.3-5.9 in. Hg

D. 4.3-5.9 psi

CHAPTER 10

ICE AND RAIN PROTECTION

CONTENTS

Page

INTRODUCTION................................................................................................................. 10-1

GENERAL ............................................................................................................................ 10-1

ICE PROTECTION—PNEUMATIC SOURCE ................................................................... 10-2





Wing and Horizontal Stabilizer Deice System .............................................................. 10-2

Controls, Indicators and Operation................................................................................ 10-5

Brake Deice System....................................................................................................... 10-6

Control and Indicator ..................................................................................................... 10-6

Operation ....................................................................................................................... 10-6

ICE PROTECTION—ELECTRICAL SOURCE.................................................................. 10-8

Windshield Heat............................................................................................................. 10-8

Controls.......................................................................................................................... 10-8

Operation ....................................................................................................................... 10-8

Propeller Heat ................................................................................................................ 10-8

Controls, Indicators and Operation.............................................................................. 10-10

Pitot Heat ............................................................................................................................ 10-11

Controls and Operation................................................................................................ 10-11

Stall Warning Vane Heat.............................................................................................. 10-12

WINDSHIELD WIPERS .................................................................................................... 10-17

Controls and Operation................................................................................................ 10-17

WING ICE LIGHTS ........................................................................................................... 10-18Location and Control................................................................................................... 10-18

LIMITATIONS.................................................................................................................... 10-18

QUESTIONS....................................................................................................................... 10-22





ILLUSTRATIONS

Figure Title Page

10-1 Weather-Protected Airplane Surfaces .................................................................... 10-210-2 Ice and Rain Protection Controls and Indicators

(BB-1439, 1444 and Subsequent) .......................................................................... 10-3

10-3 Ice and Rain Protection Controls and Indicators(Prior to BB-1444, Except BB-1439)..................................................................... 10-4

10-4 Wing and Horizontal Stabilizer Deice Boots System Control ............................... 10-5

10-5 Brake Deice System ............................................................................................... 10-7





10-6 Windshield Anti-Ice System .................................................................................. 10-9

10-7 Propeller Boots Heat-Control and Indicator ........................................................ 10-10

10-8 Pitot Probes and Heat Controls............................................................................ 10-12

10-9 Stall Warning Vane and Heat Controls ................................................................ 10-12

10-10 Heated Fuel Vent and Control.............................................................................. 10-13

10-11 Powerplant Intake Ice Protection (BB-1439, 1444 and Subsequent) .................. 10-14

10-12 Powerplant Intake Ice Protection (Prior to BB-1444, Except BB-1439)............. 10-15

10-13 Engine Intake Inertial Vane Positions and Bypass Door ..................................... 10-17

10-14 Windshield Wiper Control ................................................................................... 10-17

10-15 Wing Ice Inspection Light and Control................................................................ 10-18

CHAPTER 10ICE AND RAIN PROTECTION





INTRODUCTION

Ice, rain, and fogging can adversely affect a flight. Several systems have been includedon the Super King Air to protect those surfaces susceptible to the effects of weather.

Three sources of energy are used to prevent or to break up ice formations on the airplane’ssurfaces: engine bleed-air (pneumatics), electrical power, and engine exhaust.

GENERAL

Surfaces kept ice-free by engine bleed-air (pneumatics) are:

Surfaces kept ice- and/or water-free by elec-trical energy are:

Surfaces kept ice-free by engine exhaustgases are:

• The air inlets for both engines

Figure 10-1 illustrates the location of the sur-faces so protected.

Heated pitot tubes, stall warning vane, wind-shield panes, fuel vents, and the engine inletlips prevent ice from forming and are com-ponents of the anti-ice systems.

The inflatable boots on the wings and hori-zontal stabilizer and the electrically-heatedpropeller deicers remove accumulated ice andare considered to be the deice system

ICE PROTECTION—PNEUMATIC SOURCE

WING AND HORIZONTALSTABILIZER DEICE SYSTEM

The leading edges of the wings and horizon-tal stabilizer are protected against an accu-mulation of ice buildup. Inflatable bootsattached to these surfaces are inflated whennecessary by pneumatic pressure to breakaway the ice accumulation and are deflated bypneumatic-derived vacuum. The vacuum isalways supplied while the boots are not in useand are held tightly against the skin.





are considered to be the deice system.

Also, to prevent ice from accumulating on theengine compressor intake screen, an inertialvane separating system is installed.

The ice and rain controls and indicators are lo-cated on the main instrument panel (Figures10-2 and 10-3).

and are held tightly against the skin.

Never take off or land with the bootsinflated. Do not operate deice bootswhen OAT is below –40°C (–40°F).

CAUTION





MAXGEAR EXTENSION

MAXGEAR RETRACT

MAXGEAR EXTENDED

MAXAPPROACH FLAP

MAXFULLDOWNFLAP

MAXMANEUVERING

181 KNOTS

163 KNOTS

181 KNOTS

200 KNOTS

157 KNOTS

181 KNOTS

AIRSPEEDS (IAS)

OPERATION LIMITATIONSTHISAIRPLANEMUST BEOPERATEDASANORMALCATEGORYAIRPLANEINCOMPLIANCEWITH

THEOPERATINGLIMITATIONSSTATEDINTHEFORMOF PLACARDS,MARKINGSANDMANUALS

NO ACROBATICMANEUVERSINCLUDING SPINSARE APPROVED

THISAIRPLANEAPPROVED FORVFR, IFR, &DAY&NIGHT OPERATIONANDINICING CONDITIONS

CAUTION

STALLWARNINGISINOPERATIVEWHENMASTERSWITCHISOFF

STANDBYCOMPASSISERRATICWHENWINDSHIELDANTI-ICEAND/ORAIR CONDITIONING ISON

DO NOT OPERATEONDRYGLASS

WINDSHIELD WIPERSOFF

PARK SLOW

FAST

OFF

MASTER

PANELLIGHTS

ON

OVERHEADFLOOD

LIGHTS

OFFBRT

INSTRUMENTINDIRECT

LIGHTS

OFFBRT

AVIONICSPANELLIGHTS

OFFBRT

ENGINEINSTRUMENT

LIGHTS

OFFBRT

PILOTFLIGHTLIGHTS

OFFBRT

OVERHEAD

SUBPANEL&CONSOLE

LIGHTS

OFFBRT

SIDEPANELLIGHTS

OFFBRT

COPILOT GYROINSTRUMENT

LIGHTS

OFFBRT

COPILOTFLIGHTLIGHTS

OFFBRT

4 0 6 08020

0 3010 20

0

PUSH

FORVOLTS

100

DCVOLTS

%LOAD

Beechcraft

4 0 6 08020

0 3010 20

0

PUSH

FOR VOLTS

100

DCVOLTS

%LOAD

Beechcraft

100

40 608020

0 3010 20

0

PUSH

FOR VOLTSDC VOLTS

% LOAD

Beechcraft

40 608020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

DO NOT OPERATEON DRY GLASS

WINDSHIELD WIPERSOFF

PARK SLOW

FAST

08020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

400 410420380

390

100 1301 10 1 20PUSH

FOR VOLTS

ACVOLTS

FREQ



PARKING BRAKEOFF

ENGINEANTI-ICE

ON

MAIN

OFF

ACTUATORSTANDBY

LEFT RIGHT

COLLINS

LANDING TAXI ICE NAV RECOG

OFF

LIGHTS

L E FT R I GH T

I N

P U L L

0

10 20

30PROP AMPS

3 3N3

F O R

0

3 0

6 0 9 0

1 2 0 150 1 8 0

2 1 0 2 4

0 2 7 0

3 0 0

3 3 0

COMPASSCORRECTIONCALIBRATE WITH

RADIOON

S T E E

STALLWARN

BRAKE

ICE PROTECTIONPROPWSHLD ANTI-ICE

NORMAL

DEICECYCLESINGLE PITOT

P ILOT COP ILOT LEFT RIGHT

OFF

OFF

AUTO INNER FUEL VENT

HI

OFF OUTER

ICE VANE

AUTOFEATHER

EXTEND

RETRACT

OFF

ARM

LEFT RIGHT



Each wing has an inboard and an outboardboot. The horizontal section of the tail hasonly one boot from the left and right segmentsof the horizontal stabilizer. The vertical sta-bilizer is not, nor does it have to be, deiced

(Figure 10-1).

CONTROLS, INDICATORS ANDOPERATION

The three-position switch in the ice protectiongroup labeled DEICE CYCLE SINGLE–OFF– MANUAL controls the operation of the boots.

This switch is spring-loaded to the center OFF

be selected to the SINGLE cycle (up) positionand released (Figure 10-4). Pressure-regu-lated bleed air from the engines’ compressorssupply air through a distributor valve to inflatethe wing boots. After an inflation period of six

seconds, an electronic timer switches the dis-tributor to deflate the wing boots with vacuum,and a four-second inflation begins in the hor-izontal stabilizer boots. After these boots havebeen inflated and deflated, the cycle is com-plete, and all boots are again held down tightlyagainst the wings and horizontal stabilizer byvacuum. The spring-loaded switch must beselected up again for another cycle to occur.

Each engine supplies a common bleed-air



p gposition. When approximately one-half to oneinch of ice has accumulated, the switch should

manifold. To ensure the operation of the sys-tem if one engine is inoperative, a check valve

ENGINE P3

BLEED AIRSOURCE

ENGINE P3

BLEED AIR

SOURCE

BLEED AIR FLOWCONTROL UNIT

BLEED AIR FLOWCONTROL UNIT

DEICEBOOT

DEICEBOOT

DEICEBOOT

DEICEBOOT

BRAKE DEICEVALVE BRAKE DEICE

VALVE

PNEUMATICSHUTOFF

VALVE

PNEUMATICSHUTOFFVALVE

PNEUMATICCONTROL

ASSEMBLY

VACUUM REGULATOR

LEGENDSTALLBRAKE


NORMAL

SURFACEDEICE

P ILO T C OP ILO T LEFT RIGHT

OFF

FUEL VENT

HIOFF

A UT O M AN U AL

OFF

is incorporated in the bleed-air line from eachengine to prevent the loss of pressure throughthe compressor of the inoperative engine.

If the boots fail to function sequentially, theymay be operated manually by selecting thedown position of the same DEICE CYCLEswitch. Depressing and holding it in the MAN-UAL (down) position inflates all the boots si-multaneously. When the switch is released, itreturns to the (spring-loaded) OFF position,and each boot is deflated and held by vacuum.

A single circuit breaker located on the copi-lot’s side panel, receiving power from the No.1 dual-fed bus, supplies the electrical opera-

main landing gear. If installed, this high-pres-sure and high-temperature air is routed througha solenoid control valve in each main whee lwell, through a flexible hose on the main gear strut, and to the distribution manifold aroundthe brake assembly (Figure 10-5).

The brake deice system can be used on theground or in flight to prevent or melt away anyice accumulation.

CONTROL AND INDICATOR

The BRAKE DEICE switch in the anti-icegroup on the pilot’s right subpanel (Figures10-2 and 10-3) activates the valves, allow-





tion of both boot systems.

The boots operate most effectively when ap-proximately one-half to one inch of ice hasformed. Very thin ice will crack and couldcling to the boots and/or move aft into un-protected areas.

When operated manually, the boots shouldnot be left inflated longer than necessary toeliminate the ice, as a new layer of ice maybegin to form on the expanded boots and be-come unremovable.

If one engine is inoperative, the loss of its pneu-matic pressure does not affect boot operation.

Refer to LIMITATIONS in this chapter for additional information.

Electrical power to the boot system is requiredto inflate the boots in either single-cycle or manual operation, but with a loss of this power,the vacuum will hold them tightly against theleading edge.

BRAKE DEICE SYSTEM

10 2 and 10 3) activates the valves, allowing the pneumatic air to enter the brake man-ifolds. When this switch is activated, bothsolenoid valves are opened, and the greenBRAKE DEICE ON light on the caution ad-visory annunciator panel illuminates to ad-vise that both solenoids are being activated

to the open position (Figure 10-5). The lightdoes not, however, ensure that the valveshave actually opened. Conversely, if theBRAKE DEICE switch is turned off, the lightshould extinguish. However, it is possiblethat the valves are stuck in the open position.Confirmation that the valves are openingand closing can be made by observing a slightincrease or decrease in ITT when BRAKE

DEICE is cycled. The circuit breaker for thebrake deice system is located on the copilot’sside panel in the weather group labeledBRAKE DEICE.

OPERATION

With the landing gear extended, the brakedeice system may be operated on a continu-

ous basis, provided that the limitations listedin that section are observed.

F OR

T RAI N

S UP E R

KI N G

18 PSI

PRESSURE

REGULATOR

18 PSI

PNEUMATIC

PRESSURE

PNEU

RIGHT

P3

AIR

PNEU

LEFT

P3

AIR

N CN C



1 0 - 7

NI N G P URP O S E S ONL Y

G

A I R

2 0 0 / B 2 0

0 P I L O T T R A I NI N G MA N U A L



LEFT

BRAKE

DEICE

MANIFOLD

RIGHT

BRAKE

DEICE

MANIFOLD

VDC

BRAKE

DEICE C/B

N.C.N.C.

BRAKE DEICE

N.C. VALVES

GEAR

UPLOCK

10

MIN

BRAKE DEICE

TIMER PCB

BRAKEDEICE

DUAL FED

BUS NO.1

Figure 10-5. Brake Deice System

A minimum of 85% power on each engine isnecessary to maintain proper boot inflation if the hot brake system is on.

A 10-minute timer is activated when the gear

is retracted, which allows sufficient time for the brakes to dry.

The system should not be used continuouslyabove 15°C ambient temperature. Both in-strument (pneumatic) valves must be open for use of the system.

The brake deice system is the single biggestuser of engine bleed air. During an enginefailure, the rudder boost system may be in-

ti h th b k d i t i i

over a detent before it can be moved to the HI(down) position, preventing inadvertent se-lection of the HI position when moving theswitch from NORMAL to OFF.

The two control units receive power throughtwo 5-amp control circuit breakers located ona panel on the forward pressure bulkhead, notaccessible by the crew in flight. The windowheaters are each supplied by 50-amp circuitbreakers located in the power distributionpanel under the floor forward of the main spar.

OPERATIONEither or both windshields may be heated a t





operative when the brake deice system is in usebecause there isn’t enough differential pres-sure to activate the system.

ICE PROTECTION—

ELECTRICAL SOURCE

WINDSHIELD HEAT

Both windshields are heated by resistance wireembedded in the glass. A thermal sensor withinthe lamination monitors the glass temperatureand feeds a control signal into a controller unit.

The controller regulates the current flow to theembedded wire. Normally, a constant temper-ature of +95°F to +105°F is maintained (Figure10-6). However, at cold temperatures and highairspeeds, the system may not be able to main-tain an ice-free windshield.

The windshields can be operated at two heatlevels. Normal heating supplies heat to the

broadest area. High heating supplies a higher intensity of heat to a smaller but more essen-tial viewing area.

yany time, as overheating is prevented by ther-mal sensors. Each window is fed from the leftor right generator bus through a circuit breaker located in the power distribution panel under the floor forward of the main spar. The panelswitch closes a relay, which supplies current

to the windshields, subject to the control of thetemperature controller and thermal sensors.

Windshield heat may be used at any time, butit causes erratic operation of the magnetic com-pass, and could result in distorted visual cues.

PROPELLER HEAT

An electrically-heated boot on each blade, de-ices the propellers (BB-2 through 815, 817-824, 991; BL 1-29, these boots were dividedinto an inner and outer segment). The boot,firmly cemented in place, receives currentfrom a slip ring and brush assembly on the pro-peller shaft. The slip ring transmits current tothe deice boot. The centrifugal force of thespinning propeller and airblast breaks the ice

particles loose from the heated blades.



HEATINGWIRES

OVERTEMP

SENSOR OVERTEMPSENSOR95° TO 105°



STALLWARN

BRAKEDEICE


NORMAL

SURFACEDEICE

SINGLEPITOT

PILOT COPILOT LEFT RIGHT

OFF

OFF

AUTO MANUAL FUEL VENT

HIOFF

CB

CB

CB

CB

TEMPERATURECONTROLLER

TEMPERATURECONTROLLER

On models BB-816, 825-990, 992 and subse-quent; BL 30 and subsequent, the following se-quence is followed:

• For 90 seconds—entire right propeller

• For 90 seconds—entire left propeller

On models prior to BB-2 through 815, 817-824, 991; BL 1-29, the most common timer isused and this sequence is followed:

• For 30 seconds—right outer elements

• For 30 seconds—right inner elements

• For 30 seconds—left outer elements

• For 30 seconds—left inner elements

For both versions, manual bypass of the timer is possible. Refer to LIMITATIONS in thischapter for additional information on pro-peller deicing.

Figure 10-7 shows the control and circuitbreakers for the two configurations.


The propeller deice boots are controlled by acircuit-breaker type switch and a two-position

PROP toggle switch. When the possibility of ice buildup exists, the PROP AUTO switch la-beled AUTO–OFF should be set to the AUTO





For 30 seconds left inner elements

Once the system is turned on for automatic op-eration, it cycles continuously.

position, initiating the timer sequencing of the boots. An ammeter labeled PROP AMPSon the copilot’s left subpanel indicates the

STALLWARN

BRAKEDEICE


NORMAL

DEICECYCLESINGLE PITOT

PILO T CO PILOT LEFT RIGHT

OFF

OFF


HIOFF

COLLINS

BB-2 THROUGH 815, 817-824, 991

0

10 20

30PROP AMPS

STALLWARN

BRAKEDEICE


NORMAL

PITOT

OFF

MANUAL

PILO T CO PILOT

LEFT RIGHT

LEFT RIGHT

OFF

OFF


HIOFF OUTER

DEICECYCLESINGLE

0

10 20

30PROP AMPS

current flow to the propeller elements (Figure10-7).

Normal current flow within the green arc is 18to 24 amperes for all 4-bladed airplane versions

(14 to 18 for 3-bladed versions). The amme-ter may flicker as the timer sequences to thenext combination of boots, but this flicker isvery difficult to see.

The ammeter should be monitored to makecertain that current flow is approximately thesame for all timer positions. Variations couldindicate that uneven heating is occurring, re-

sult ing in possible propeller vibrat ions.However, loss of one heating element (whenthe prop ammeter indicates a less than green

panel in the ice group. The manual system’scircuit breakers are located on the fuel controlcircuit-breaker panel, located on the pilot’s leftside panel in the PROP DEICE group. Thecontrol circuit breaker is for the INNER/

OUTER switch, depending on the model. ThePROP LEFT and PROP RIGHT circuit break-ers control power to the prop elements in themanual mode.

Although this system is called aprop deice system, pilot manage-ment of the system should be as ananti-ice system

CAUTION





the prop ammeter indicates a less than greenarc value) does not mean that the entire sys-tem must be turned off. (Refer to the appro-priate section of the Fl ight Manual .)

A manual backup of the automatic sequenc-ing is installed in case the timer fails to oper-

ate properly. The PROP MANUAL–OFFswitch (or on earlier aircraft as listed above,the PROP INNER–OUTER switch), providescurrent to the boots (Figure 10-7). On air-planes with a single boot element per pro-peller, with the PROP AUTO switch in theOFF position, holding the PROP MANUALswitch in the MANUAL position for approx-imately 90 seconds deices both props at thesame time, applying heat to all the boots. Onairplanes with a two-segment boot per pro-peller, the spring-loaded switch must be heldto the OUTER position until the ice has beendislodged from both propellers’ outer boots.Then it must be held to the INNER position todeice both propellers’ inner boots.

The PROP AMPS ammeter does not register current flow in the MANUAL mode of oper-ation. The increased load, however, can be

anti ice system.

PITOT HEAT

A heating element in each pitot probe pre-vents ice and moisture buildup. There is no

thermal protection for the heating system ex-cept its own circuit-breaker switch.

CONTROLS AND OPERATION

Each pitot heater has its own circuit-breaker switch that can be left in the ON position dur-ing flight (Figure 10-8).

The two circuit-breaker switches are fed off separate dual-fed buses. The left is on the No.1 and the right is on the No. 2 dual-fed bus.

It is recommended that the pitot heat not be op-erated on the ground except for testing or for short intervals to remove ice or snow fromthe mast. However, it should be turned on for takeoff when icing conditions are suspected.



PITOT PROBES



STALL WARNING VANE HEAT

Heat is applied to both the mounting plate andthe vane. There is no thermal protection of theheating element except its own control cir-cuit-breaker switch.

Control and OperationA circuit-breaker switch labeled STALLWARN in the ICE group controls the heatingfunction. Due to the left landing gear squatswitch, the current flow to the heater is min-

imal while the airplane is on the ground. Inflight, full current is supplied (Figure 10-9).

FUEL VENT HEAT

Controls and OperationElectric heaters prevent ice formation in the fuelvent system. Each wing fuel system has its own

anti-ice system, operated by the two switchesin the ICE group labeled FUEL VENT (Figure

STALL WARNING VANE

HEAT CONTROLS

Figure 10-9. Stall Warning Vane andHeat Controls

HEAT CONTROLS

Figure 10-8. Pitot Probes and HeatControls

MISCELLANEOUSSYSTEMSPOWERPLANT

The engine air inlet lips are heated by engineexhaust gases to prevent the formation of ice(Figures 10-11 and 10-12). On airplanes BB-1266, BL-129 and subsequent, hot engine ex-haust flows from the left stack, through the lip,and exits out the right stack. Prior to BB-1266,hot engine exhaust is routed downward andinto each end of the inlet lip and eventually

ducted out through the bottom of the lip.

The system is automatic and does not requireil t ti



HEATED FUEL VENT



On earlier models prior to 1979, each P3pneumatic fuel control l ine is protectedagainst ice by an electrically-heated jacketwhich receives electric current if the enginecondition levers are moved out of fuel cut-off range. On later models and all B200 air-planes, a heated jacket and a filter is installedfor this purpose.

pilot action.

To prevent the engine compressor inlet screenfrom accumulating ice, an inertial vane sep-arating system is installed. When the ice vanesare lowered, they deflect the airstream slightly

downward, creating a venturi effect. At thesame time, an inertial vane bypass door under the cowling is also opened, allowing an exit.

As the ice particles or water droplets enter theengine inlet, the airstream with these particlesis accelerated because of the venturi effect.These frozen moisture particles, due to thegreater mass and, therefore, greater momentum,

accelerate past the screen area and vent overboardthrough the bypass door. However, the airstreammakes the sudden turn easier because the air isfree of the ice particles which are being de-flected rearward and overboard.

The inertial vane and the inertial vane bypassdoor are closed for normal flying conditions,thus directing the air into the powerplant in-

take and oil cooler.

HEAT CONTROLS

Figure 10-10. Heated Fuel Vent andControl



FLAPS

DOWN

TAKEOFF

AND

APPROACH

UP20

60

80

4

6

241

1

0

.5

.5

2CABIN CLIMBTHDS FT PER MIN

7

1

2

3

45

6

0PSI

CA B IN

A L T0 0 F T

40 5

10

1520

25

30

35

0



ENGINE ANTI-ICE

AUTOFEATHER

ON

MAIN

OFF

ACTUATOR

STANDBY

O GO

LEFT RIGHT

OFF

ARM

ENG AUTO

IGNITION

PILOT

AIR

PULL


NORMAL

DEICE


OFF

LIGHTS

PILOT COPIL OT LEFT RIGHT

LEFT RIGHT

OF

F


HI

DEFROST

AIR

PULL

RIGHTLEFT

OFF

150



FLAPS

DOWN

TAKEOFF

AND

APPROACH

UP20

60

80

4

6

241

1

0

.5

.5

2CABIN CLIMBTHDS FT PER MIN

7

1

2

3

45

6

0PSI

CA B IN

A L T0 0

F T

40 5

10

1520

25

30

35

0



ICE VANE

AUTOFEATHER

EXTEND

TEST

RETRACT

OFF

PROP GOV TEST

ARM

LEFT RIGHT

TEST

OFF

ARM

ENG AUTO IGNITION

PILOT

AIR

PULL

ONSTALL

WARN

BRAKE


NORMAL

DEICECYCLE

SINGLE PITOT


OFF

LIGHTS

PILOT COPILOT LEFT RIGHT

LEFT RIGHT

OFF

OFF


HI

DEFROST

AIR

PULL

ON

RIGHTLEFT

OFF

ICE VANECONTROLS

OUTER


To extend or retract the ice vanes, the ENG

ANTI-ICE toggle switches (prior to BB-1444,except 1439, ICE VANE toggle switches) aremoved to the appropriate position. They arelocated on the pilot’s left subpanel (Figures 10-11 and 10-12).

In the ice-protection mode, the extended po-sition of the vane and the bypass door is in-dicated by the green annunciator lights, L

ENG ANTI-ICE and R ENG ANTI-ICE (prior to BB-1444, 1439; L ICE VANE EXT and RICE VANE EXT) on the caution-advisorypanel. When retracted, the lights extinguish.

nunciator indication provided the appropri-ate circuit breaker has already been pulledvia the checklist. When the vane is success-fully positioned with the manual system, theamber annunciator light(s) will extinguish.

Prior to BB-1444, except 1439

Once the manual override system hasbeen engaged (i.e., any time the man-

ual ice vane T-handle has been pulledout), do not attempt to retract or ex-tend the ice vanes electrically, evenif the T-handle has been pushed back

CAUTION





panel. When retracted, the lights extinguish.

In addition, two amber lights labeled L and RENG ANTI-ICE (prior to BB-1444, except1439 L ICE VANE and R ICE VANE) are pro-vided on the caution-advisory panel. If either

engine’s inertial vane and inertial vane by-pass door have not attained the selected posi-tion (either open or closed) within 15 seconds,the appropriate light illuminates.

For BB-1439, 1444 and subsequent a backupsystem consists of dual actuators and con-trols. Illumination of the L and R ENG ANTI-ICE (amber) annunciators indicates that the

system did not operate to the desired posi-tion. Immediate illumination of the L or RENG ANTI-ICE (yellow) annunciator indi-cates loss of electrical power, whereas de-layed illumination indicates an inoperativeactuator. In either event, the STANDBY ac-tuator should be selected.

Prior to BB-1444, except 1439, a mechani-

cal backup system is provided for manuallylowering or raising the vanes. It is actuate dby pulling the T handles just below the pilot’s

if the T handle has been pushed backin, until the override linkage in theengine compartment has been prop-erly reset on the ground. However,the pilot can raise or lower the vanesrepeatedly, any time, with the man-

ua l s y s t e m e nga ge d . ( S e e t he Raytheon Maintenance Manual for resetting procedure.)

NOTE

Lowering the ice vanes will resultin a slight ITT rise and a significantloss of torque at normal cruise power settings.

The circuit breakers for the ice vanes are lo-cated on the copilot’s right side panel in theengine group and are labeled MAIN ENGANTI-ICE and STBY ENG ANTI-ICE (prior to BB-1444, except 1439 only one circuitbreaker exists labeled ICE VANE CONTROL).

The movable vane and the bypass door must bel d i h i h i i i



INERTIAL VANE RETRACTED INERTIAL VANE BYPASS DOOR EXTENDED



WINDSHIELD WIPERS

CONTROLS AND OPERATION

The dual wipers are driven by a mechanism op-erated by a single electric motor, all locatedforward of the instrument panel.

The windshield wiper switch is located on the

OPERATION LIMITATIONSTHISAIRPLANEMUST BEOPERATEDASANORMALCATEGORYAIRPLANEINCOMPLIANCEWITHTHEOPERATIN G LIMITATIONSSTATED INTHE FORM OF PLACARDS, MARKINGSANDMANUALS

NO ACROBATICMANEUVERSINCLUDING SPINSAREAPPROVEDTHISAIRPLANEAPPROVED FORVFR, IFR, &DAY&NIGHT OPERATIONANDINICING CONDITIONS

CAUTION

STALLWARNING ISINOPERATIVEWHENMASTERSWITCHISOFFSTANDBYCOMPASSISERRATICWHENWINDSHIELDANTI-ICEAND/ORAIRCONDITIONINGISON

DONOTOPERATEONDRYGLASS

WINDSHIELDWI PERSOFF

PARK SLOW

FAST

OFF

MASTERPANELLIGHTS

ON

OVERHEADFLOODLIGHTS

OFFBRT

INSTRUMENTINDIRECTLIGHTS

OFFBRT

AVIONICSPANELLIGHTS

OFFBRT

ENGINEINSTRUMENT

LIGHTS

OFFBRT

PILOTFLIGHTLIGHTS

OFFBRT

OVERHEADSUBPANEL&CONSOLE

LIGHTS

OFFBRT

SIDEPANELLIGHTS

OFFBRT

COPILOTGYROINSTRUMENT

LIGHTS

OFFBRT

COPILOTFLIGHTLIGHTS

OFFBRT

DO NOT OPERATE

ON DRY GLASS

WINDSHIELD WIPERS

OFF

PARK SLOW

FAST

Figure 10-13. Engine Intake Inertial Vane Positions and Bypass Door

INERTIAL VANE BYPASS DOOR EXTENDED

Windshield wipers may be damagedif used on a cracked outer panel.

The circuit breaker is on the copilot’s right CBpanel in the WEATHER group.

WING ICE LIGHTS

LOCATION AND CONTROL

The wing lights are located on the outboardside of each nacelle. The circuit-breaker switchis located on the pilot’s right subpanel in the

LIMITATIONS

Safe operation in icing conditions is dependentupon pilot knowledge regarding atmospheric

conditions conducive to ice formation, famil-iarity with the operation and limitations of theinstalled equipment, and the exercise of good judgment when planning a fl ight in to areaswhere possible icing conditions might exist.

When icing conditions are encountered, theperformance characteristics of the airplanewill deteriorate.

Increased aerodynamic drag increases fuelconsumption, thereby reducing the airplane’srange and making it more difficult to maintain

CAUTION





is located on the pilot s right subpanel in theLIGHTS group above the ICE group (Figure10-15).

range and making it more difficult to maintainspeed.

Decreased rate of climb must be anticipated,not only because of the decrease in wing andempennage efficiency, but also because of

the possible reduced efficiency of the pro-pellers and increase in gross weight. Also,the use of the inertial ice vanes may result inlost performance.

Abrupt maneuvering and steep turns at lowspeeds must be avoided because the airplanewill stall at higher than published speeds withice accumulation. On final approach for land-

ing, increased airspeed must be maintainedto compensate for this increased stall speed.After touchdown with heavy ice accumula-tion, landing distances may be as much astwice the normal distance due to the increasedlanding speed.

During descent, a minimum of 85% power oneach engine is necessary to maintain proper

boot inflation if the airplane is equipped withand using the hot brake system.

WING ICE INSPECTION LIGHT

If the landing gear is retracted, the systemmay not be operated longer than 10 minutes,which is one timer cycle. The annunciator light should be monitored. If it does not au-tomatically go out after approximately 10 min-

utes following gear retraction, the systemshould be manually turned off.

Both engine bleed-air sources must be in op-eration to use the brake deice system on bothsides.

A minimum speed of 140 KIAS is necessaryto prevent ice formation on the underside of

the wing, which cannot be adequately deiced.

Windshield heat may be used at any time, butit causes erratic operation of the magnetic com-

If the amber ENG ANTI-ICE (prior to BB-1444, except 1439, ICE VANE) annunciatorsilluminate upon extension (Figures 10-11 and10-12), the ice vanes may not have positionedproperly.

The STBY actuators (or prior to BB-1444,except 1439, manual control) should be usedto retract or to extend them. A reliable backupcheck on the position is to closely monitor engine torque. Normal torque may be re-gained with the power levers, observing theITT limits.

If in doubt, extend the vanes. Engine

CAUTION





it causes erratic operation of the magnetic compass, and could result in distorted visual cues.

Windshield wipers may be damagedif used on a cracked outer panel.Heating elements may be inopera-tive in area of crack.

During sustained icing conditions, 226 KIASis the maximum effective airspeed due to thelimitations of the windshield heating system.

In flight, the boots should be cycled onceevery time the ice accumulation is approxi-mately one-half to one inch thick.

Should either engine fail in flight, there is suf-ficient air for the entire deice operation (ex-cept for the hot brake operation). Should theautomatic cycling of the boots fail, the MAN-UAL position should be used for inflation.

While in flight, the engine ice vanes must be

, gicing can occur even though no sur-face icing is present. If freedom fromvisible moisture cannot be assured,engine ice protection should be ac-tivated. Visible moisture is moisture

in any form: clouds, ice crystals,snow, rain, sleet, hail, or any com-bination of these. Ice vanes should beretracted at +15°C and above to as-sure adequate engine oil cooling.Operation of strobe lights will some-times show ice crystals not normallyvisible.

Prior to BB-1444, except 1439, once the icevanes have been actuated manually, do not at-tempt to retract or extend them electricallyuntil they have been reset, as this may causedamage to the system.

During flight in icing conditions, fuel ventheat, pitot heat, prop deice, windshield heat,

and stall warning heat should all be on.

The wing ice lights should be used as re-

CAUTION



NOTES

For manual backup on airplanes BB-816, 825-990, 992 and subsequent;BL-30 and subsequent, the switch isheld to the ON position for approxi-mately 90 seconds. This backup sys-tem may be repeated as required andthe loadmeter should be monitoredfor a deflection of approximately 5%.

For manual backup on airplanesprior to BB-2 through 815, 817-824,991; BL-1 through 29, the PROP

INNER/OUTER swi tch i s pos i -tioned first to the OUTER, then tothe INNER position for 30 secondsin each position and the loadmeter





pmonitored for a deflection of ap-proximately 5%.

1. T he w in g an d ho ri zo nt al s ta bi li ze r leading edges are deiced by:

A. Pneumatically-inflated boots.

B. Pneumatically-heated boots.

C. Pneumatically-inflated and heatedboots.

D. Pneumatically-inflated/electrically-heated boots.

2. If wing and horizontal stabilizer bootswere inflated with only a thin coating of ice on them:

A. Th e sys t e m wo u l d wo r k m os t

5. If the airplane is f lying through icingconditions, what is the minimum speednecessary to keep the bottom of the wings’leading edges ice-free?

A. 100 knots

B. 120 knots

C. 140 knots

D. 160 knots

6. If there is a loss of electrical power to thetimer of wing and horizontal stabilizer boots while they are inflated:

A. The boots wil l nei ther inf late nor



QUESTIONS



. e sys t e wo u d wo os tefficiently.

B. The ice would only crack and maynot break loose.

C. The ice would only begin to melt and

then refreeze.D. The cracking ice might rupture the

boot.

3. When the deice boots are automaticallycycled, the timer sequence is as follows:

A. W in g s a n d h o ri z o n ta l s t a b il i z e r simultaneously, 10 seconds.

B. Inboard boots on wings, six secondsoutboard and horizontal stabilizer,four seconds.

C. Wings and tail, six seconds expanded,four seconds contracted.

D. Wing, six seconds; horizontal stabi-lizers, four seconds.

4. If the boots are held inflated too longthey:

. e boots w e t e ate odeflate.

B. The boots will stay inflated.

C. T h e b oo t s w il l c o ll a p se u n d er vacuum.

D. The boots deflate very slowly.

7. I f t he BR AK E D EI CE s wi tc h i n t heanti-ice group is selected to the ONposit ion, and the lower annunciator panel light BRAKE DEICE ON is illu-minated, the:

A. Brake mani fo lds a re mos t li ke ly

receiving hot bleed air.B. B r a ke m a n i fo l d s a r e d ef i n it e l y

receiving hot bleed air.

C. Brake manifolds are a t operat ingtemperature.

D. Brake manifolds are receiving anadequate supply of bleed air.

8. After the wheels have retracted into thewheel wells:

9. Brake deicing is the largest single load onthe bleed-air system. If the brake deicingis used with other pneumatic systems, suchas boot inflation, at what level must theengine N1 be maintained?

A. 85%B. 75%

C. 65%

D. 55%

10. The windshield temperature is regulatedand affected by:

A. Cockpit ambient temperature.B. Outside ambient temperature.

C. H e a t s en s o rs w h i ch s e n s e gl a s stemperature.

13. The engine compressor inlet screen isprotected from ice particles by:

A. An electrically-heated structure of in-take vanes.

B. An inertial vane system.

C. A p n e u m a ti c a l l y - he a t e d i n t a k emanifold.

D. Hot exhaust gases blown across theintake.

14. P r io r to BB-1444 , excep t 1439 , themechanical backup operat ion of the

eng ine ine r t i a l vanes in case o felectromechanical failure:

A. Can be performed only once per flight.

B. Can be performed twice per flight.





D. An accumulation of ice and snow.

11. The current requirements of the propeller boots must be monitored because:

A. A heavy current flow in one boot mayor may not be sufficient to trip thebreaker.

B. Overheating propellers can seriouslyweaken the structural integrity.

C. Defective boots can cause unevendeicing and serious vibration.

D. Heavy current flow may burn up the

brushes and slip rings.

12. During icing conditions in flight, the stallwarning:

A. Is reliable as long as the stall warningvane heat is on.

B. Is unreliable unless the wing boots andwarning vane heat boots are both on.

C. Is unreliable.

D Indication speeds is automatically

C. Must be used to the exclusion of theelectr ical system for the f l ight’sduration.

D. Must be used only if the electrome-chanical system fails.

15. The windshield wipers may be used under which of these circumstances?

A. On the ground or in flight on a wetwindshield

B. On the ground or in flight up to 200knots on a wet windshield

C. On the ground or during takeoff on awet or dry windshield

D. Under any circumstances

16. Each of two fuel vent systems is kept icefree by:

A. The hot oil heat exchanges around thevent probes.

B. Continuously active electric heatersaround the vent probes.

17. Engine air intake lips are:

A. Heated by electrothermal boots.

B. Heated by exhaust gases when theengine is operating.

C. Heated by extracting bleed air whenthe engine is operating.

D. Not heated because of new nacelledesign.

18. The following statements are applicableto flight in icing conditions with oneexception. Which is it?

A. Increased fuel consumption will occur B. Reduced propeller efficiency is likely

C. I n c r ea se d s ta l l spe e d s a re t o b eexpected

21. The upper temperature limitation for safelyusing the engine anti-icing vanes is:

A. +25°C

B. +20°C

C. +15°CD. +10°C

22. In case of bleed air failure from either source:

A. All of the pneumatic deice and anti-icing equipment may still be used.

B. Only the brake deice system may not

be used.C. Do not use either brake or boot deice

except in an emergency.

D. Do not use wing deice simultaneously





D. The engines may run a little cooler

19. Just prior to brake release with the OAT+5°C or less and visible moisture encoun-

tered, what action should the pilot take?A. The inertial separator ice vanes should

be extended immediately.

B. The inertial separator ice vanes shouldbe ex tended jus t a f t e r l i f to f f i sachieved.

C. The inertial separator ice vanes shouldbe extended only after 500 feet is

reached.D. The inertial separate ice vane should

be extended only after maximum en-gine takeoff power has been achieved.

20. The deice boots should not be operatedwhen the OAT is below:

A. –30°C

B. –40°CC –50°C

g ywith any other pneumatic system.

23. Prior to BB-1444, except 1439, the man-ual control of the ice vanes:

A. May be used interchangeably with theelectromechanical controls.

B. Must be used exclusively throughoutthe flight after manual control hasbeen used once.

C. Cannot be used unless there has beena failure of the electromechanicalsystem.

D. Is entirely independent of the elec-tromechanical system.

24. When the p rope l le r deice sys t em isoperated manually, the PROP ammeter reads:

A. 14-18 amperes

B. 10-15 amperes

C. 0 amperes

D. 8-10 amperes

25. If, during flight through icing conditions,the propel ler deic ing system drawsexcessive current (higher than green arc)but does not trip the circuit breaker:

A. Disable that breaker manually.B. Run the deice system only to get rid

of excessive vibration.

C. Normal heating may be continuedith i l i i





with an occasional increase in rpm asneeded.

D. Operate the deice system in manualmode.

CHAPTER 11

AIR-CONDITIONING SYSTEM

CONTENTS

Page

INTRODUCTION ................................................................................................................. 11-1

GENERAL............................................................................................................................. 11-1

System Description and Location .................................................................................. 11-1

Air-Conditioning System Controls ................................................................................ 11-8

LIMITATIONS .................................................................................................................... 11-11





QUESTIONS....................................................................................................................... 11-12

ILLUSTRATIONS

Figure Title Page

11-1 Super King Air Air-Conditioning System, BB-1180 and After

(With Aft Evaporator)............................................................................................. 11-211-2 Super King Air Air-Conditioning System Prior to BB-1180

(With Aft Evaporator)............................................................................................. 11-3

11-3 AIR CND N1 LOW Advisory Light ...................................................................... 11-4

11-4 Floor and Ceiling Outlets ....................................................................................... 11-4

11-5 Cockpit “Eyeball” Outlets ...................................................................................... 11-5

11-6 Receiver-Dryer Sight Gage .................................................................................... 11-5

11-7 Air Control Knobs .................................................................................................. 11-6





11-8 DUCT OVERTEMP Caution Light ....................................................................... 11-6

11-9 ELECTRIC HEAT Switch...................................................................................... 11-7

11-10 RADIANT HEAT Switch and Panel ...................................................................... 11-7

11-11 ENVIRONMENTAL Group Switches and Knobs ................................................. 11-8

11-12 CABIN TEMP MODE Selector Switch ................................................................ 11-8

11-13 Air-Conditioning System Control Diagram ........................................................... 11-9

11-14 Ram-Air Scoop ..................................................................................................... 11-11

CHAPTER 11AIR-CONDITIONING SYSTEM





INTRODUCTION

The Super King Air’s air-conditioning system (Figures 11-1 and 11-2) provides thecrew and passengers with cooling, heating and unpressurized ventilation. In addition tothe heating afforded by the air-conditioning system, electric heat (radiant heat prior toBB-1444, except 1439) is available as an option. The air-conditioning system may beoperated in the heating mode and the cool ing mode either under automatic mode con-trol or manual mode control.

GENERAL• A belt-driven, engine-mounted com-

pressor (right engine)

R f i l bi



PNEUMATICBLEED AIR

AMBIENT AIRMODULATING

VALVE

PNEUMATICTHERMOSTAT

CONDENSER BLOWER

OUTLET AIR

RECEIVER-DRYER

MIXING PLENUM

WINDHSHIELD DEFROSTER(ON GLARESHIELD)

FWD PRESSURE BULKHEAD

CREWHEAT DUCT

INSTRUMENT PANEL

PILOT'S VENT

AIR CONTROLWINDSHIELD DEFROSTER

CONTROL

ENVIRONMENTAL BLEED-AIR FLOW CONTROL UNITINCLUDING MODULATINGAND SHUTOFF VALVE

AMBIENT AIRMODULATINGVALVE

ENVIRONMENTAL

BLEED-AIRSHUTOFF VALVE

PNEUMATICTHERMOSTAT

REFRIGERANTCOMPRESSOR

CONDENSER

RAM-AIR SCOOP

FRESH AIR VALVE(CLOSED WHEN PRESSURIZED)

VENT BLOWERFWD EVAPORATORAIR FILTER

FWD EVAPORATOR

RETURN AIR VALVE

RETURN AIR FILTER

COPILOT'SVENT AIR CONTROL

CABIN AIRCONTROL

CEILING DUCT/ FLOOR DUCT DIVIDER

COPILOT'SCEILING OUTLET

DUCT OVERTEMPSENSOR

CABIN AIRCONTROL VALVE

FORWARDHEATER *



FLOOR DUCTTO AFT FLOOROUTLETS

TO CEILINGOUTLETS

FWD

SIDEVIEW

DOOR

DETAIL A

AIR-CONDITIONED AIR

CEILING

OUTLET

HOT ENGINE BLEED AIR


RECIRCULATED CABIN AIR(AIR CONDITIONED WHENEVAPORATOR IS ON)

AMBIENT AIR

PRESSURE VESSEL

LEGEND

FLOOR OUTLET

CEILING OUTLETS

AIR-TO-AIRHEAT

EXCHANGER

CABIN-HEATCONTROL VALVE

FLOOROUTLET

CEILINGOUTLET

AIR INLETSCOOP

FIREWALL

BLEED-AIRSHUTOFF

VALVE ENVIRONMENTAL BLEED-AIR SHUTOFF VALVE

AND SHUTOFF VALVE

FLOOROUTLET

DOOR (COOLED AIR TOFLOOR OUTLETS)

CEILING OUTLET

AFT EVAPORATORAIR FILTER

AFT EVAPORATOR

CEILING OUTLET

FLOOR OUTLET

CEILING OUTLET


AIR-TO-AIRHEAT EXCHANGER

FIREWALLCONTROL VALVE

REFRIGERANT LINESAIR INLET

FLAPPERVALVE

AFTHEATER *

PNEUMATIC BLEED-AIRSHUTOFF VALVE




PNEUMATICBLEED-AIR


PNEUMATICTHERMOSTAT

CONDENSER BLOWER

OUTLET AIR

RECEIVER-DRYER

MIXING PLENUM

WINDHSHIELD DEFROSTER(ON GLARESHIELD)

FWD PRESSURE BULKHEAD

CREWHEAT DUCT

INSTRUMENT PANEL

PILOT'S VENT

AIR CONTROLWINDSHIELD DEFROSTER

CONTROL



ENVIRONMENTALBLEED-AIRSHUTOFF VALVE

PNEUMATICTHERMOSTAT

REFRIGERANTCOMPRESSOR

CONDENSER

RAM-AIR SCOOP

FRESH AIR VALVE(CLOSED WHEN PRESSURIZED)

VENT BLOWER

FWD EVAPORATORAIR FILTER

FWD EVAPORATOR

RETURN AIR FILTER

RETURN AIR VALVE

COPILOT'SVENT AIR CONTROL

CABIN AIR CONTROL

CEILING DUCT/

FLOOR DUCT DIVIDERCOPILOT'SCEILING OUTLET

DUCT OVERTEMPSENSOR

CABIN AIRCONTROL VALVE



FLOOR DUCTTO AFT FLOOROUTLETS

TO CEILINGOUTLETS

FWD

SIDEVIEW

DOOR

DETAIL A

AIR-CONDITIONED AIR

CEILINGOUTLET

HOT ENGINE BLEED AIR


RECIRCULATED CABIN AIR(AIR CONDITIONED WHENEVAPORATOR IS ON)

AMBIENT AIR

PRESSURE VESSEL

LEGENDFLOOR OUTLET

CEILING OUTLETS

AIR-TO-AIRHEAT

EXCHANGER


FLOOROUTLET

CEILINGOUTLET

AIR INLETSCOOP

FIREWALL

SHUTOFFVALVE ENVIRONMENTAL BLEED-

AIR SHUTOFF VALVE

AND SHUTOFF VALVE

FLOOROUTLET

DOOR (COOLED AIR TO

FLOOR OUTLETS)

CEILING OUTLET

AFT EVAPORATORAIR FILTER

AFT EVAPORATOR

CEILING OUTLET

FLOOR OUTLET

FLAPPER VALVE


AIR-TO-AIRHEAT EXCHANGER

FIREWALLCONTROL VALVE

REFRIGERANT LINESAIR INLET

CEILINGOUTLET


PNEUMATIC BLEED-AIRSHUTOFF VALVE

• An evapora tor wi th an op tiona l a f tevaporator.

• A receiver-dryer.

• An expansion valve or two expansion

valves if aft evaporation is installed.• A bypass valve.

The plumbing from the compressor, which ismounted on the right engine, is routed throughthe right wing and then forward to the con-denser coil, receiver-dryer, expansion valve,bypass valve, and evaporator—all of which arelocated in the nose of the airplane.

The high- and low-pressure limit switches andthe N1 speed switch (engine speed) preventcompressor operation outside of establishedlimitation parameters. The N1 speed switchdi th l t h h th

The forward vent blower moves recirculatedcabin air through the forward evaporator, intothe mixing plenum, into the floor-outlet ducts,and ceiling eyeball outlets. Approximately75% of the recirculated air passes through the

floor outlets while approximately 25% of theair is routed through the ceiling outlets, by-passing the mixing plenum (Figure 11-4).

The forward vent blower, with the system inAUTOmatic normally runs at low speed.

If the cooling mode is operating, refrigerantcirculates through the forward evaporator,

cooling the output air. All the air entering theceiling-outlet duct is cooler than the air en-ter ing through the f loor outlets i f ei ther BLEED AIR VALVE switch is in the OPEN po-sition. This air discharges through “eyeball”

tl t l (Fi 11 5) i th k it





disengages the compressor clutch when the en-gine speed is below 62% N1 and air condi-tioning is requested. When the N1 speed switchopens, and if air conditioning is being re-

quested, the green AIR CND N1 LOW advi-sory annunciator (Figure 11-3) will illuminate.

outlet nozzles (Figure 11-5) in the cockpitand cabin. Each nozzle is movable so theairstream may be directed as desired. Also, thevolume of air can be adjusted from full open

to closed by twisting the nozzle. As the noz-zle is twisted, a damper opens or closes toregulate airflow.

Cool air also enters the floor-outlet duct, butin order to provide cabin pressurization, warmbleed air also enters this duct any time either BLEED AIR VALVE switch is in the OPEN po-sition. Therefore, pressurized air discharged

from the floor outlets is always warmer thanthe air discharged from the ceiling outlets, nomatter what temperature mode is used.

Figure 11-3. AIR CND N1 LOW AdvisoryLight



The crew can view these components by re-moving the upper-compartment access panel,located on top of the nose section left of cen-terline. This, however, is not a normal preflightaction. If there are bubbles seen through the

sight glass (Figure 11-6), the refrigerant sys-tem is low on the refrigerant gas being used.If, after adding more refrigerant gas, bubblesare still appearing, then the system needs tobe evacuated and recharged.

RECEIVER-DRYER AND

SIGHT GAGE



NOTE

On the Super King Air 200, prior to BB310 and all cargo door airplanes, alever on each floor outlet register (ex-cept the forward facing register in thebaggage compartment) can be movedvertically to regulate the airflow. OnBB 310, 343 and all subsequent pas-senger door models, this feature hasbeen deleted. A vane-axial blower in

the nose section draws ambient air through the condenser to cool the re-frigerant gas when the cooling modeis operating. On Serial Nos. BB-345and subsequent and BL-1 and subse-quent (and any earlier serials that havecomplied with Beechcraft ServiceInstructions No. 0968 by the instal-lation of Kit Number 101-5035-1 S or

101-5035-3 S), this blower shuts off when the gear is retracted.

An optional aft evaporator and blower isavailable for additional cooling. It is located

below the center aisle cabin floor behind therear spar. The additional unit increases theairplane’s cooling capacity from 18,000 Btu(with the forward evaporator only) to 32,000Btu. Refrigerant flows through the aft evap-orator any time it flows through the forw ardevaporator; however, the additional coolingis provided only when the aft blower is op-erating, recirculating cabin air through the

aft evaporator, and routing it to the aft fl oor and ceiling outlets.

Figure 11-5. Cockpit “Eyeball” Outlets

Figure 11-6. Receiver-Dryer Sight Gage

HeatingBleed air from the compressor of each engineis delivered into the cabin for heating, as wellas pressurization. When the left landing gear safety switch is in the ground position, theambient air valve in each flow control unit isclosed; therefore, only bleed air is delivered.When airborne, bleed air is mixed with out-side ambient air from the ambient air valve ineach flow control unit until a cold air tem-perature closes off the ambient flow. Thenonly bleed air is delivered.

In the cockpit, additional air can be providedby adjusting either the pilot’s damper, whichis controlled by the PILOT AIR knob (Figure11-7), or the copilot’s damper, which is con-trolled by the COPILOT AIR knob. Movementof these knobs affects cockpit temperature by

defroster. When the knob is pushed all the wayin, the valve opens, allowing the air in the ductto be directed into the cabin floor outlets.

The DEFROST AIR knob (Figure 11-7) con-

trols a valve on the pilot/copilot heat ductwhich admits air to two ducts that deliver thewarm air to the defroster, located below thewindshields and at the top of the glareshield.

The rest of the air in the bleed-air duct mixeswith recirculated cabin air and is routed aftthrough the floor-outlet duct, which handles75% of the total airflow. If the airflow be-

comes too low in the ducting, the amber DUCTOVERTEMP caution light (Figure 11-8) illu-minates, indicating that the duct temperaturehas reached approximately 300°F.





t t p t t p t yadjusting the air volume (Figure 11-7). TheCABIN/COCKPIT AIR knob (simply CABINAIR prior to BB-1444, except 1439) controls

air volume to the cabin (Figure 11-7) and is lo-cated on the copilot’s left subpanel below andinboard of the control column. This knob con-trols the cabin air control valve. When thisknob is pulled out of its stop, a minimumamount of air passes through the valve to thecabin, thus increasing the volume of air avail-able to the pilot and copilot outlets and the

ICEEMERGENCY

PILOTAIR

PULLON

STALLWARNBRAKEDEICE


NORMAL

SURFACE

DEICESINGLE PITOT

LANDING TAXI ICE NAV

OFF

LIGHTS

P IL OT C OP IL OT LEFT

LEFT RIGHT

OFF

OFF

A UTO MA NU AL

HI

DEFROSTAIR

PULLON

OFF

COPILOTAIR

PULL

ON

INCR

ENVIRONMENTAL

HIGH

AUTO

LO

DECR

INCR

CABIN TEMP

CABIN TEMP MODE

OFF

MANUALTEMP

VENTBLOWER

MAINHEAT

AUTO

MAINCOOL CABIN

AIR

PULL

DECRELECHEAT

AFTBLOWER

ON

INSTR & ENVIR OFF

ENVROFF

OPENRIGHT

BLEED AIR VALVES

OFF

LEXTR

LDET

R

Figure 11-8. DUCT OVERTEMP CautionLight

Electric Heat (BB-1439, 1444and Subsequent)A supplemental electric heating system isavailable for cabin comfort. It is operated by

a solenoid-held switch on the copilot’s leftsubpanel placarded ELEC HEAT–OFF (Figure11-9). This system can be used in conjunctionwith an external power unit for warming thecabin prior to starting the engines, and is usedin the manual heat or automatic temp controlmode only.

ing elements. When the electric heat system isselected to OFF, the ELEC HEAT ON annun-ciator must be extinguished to indicate power is removed from the heating elements before theblowers are switched to OFF.

NOTE

The electric heat system will drawapproximately 300 amps.

The system is available for ground operationonly. If the aircraft takes off with the electricheat ON, the squat switch will remove electricpower via the solenoid-operated electric heatswitch and the switch goes to the OFF posi-tion.

Radiant Heating (Prior to BB-1444





This system uses one forward heating elementlocated in a forward duct and one aft heating el-ement located in the aft evaporator plenum.Both the forward and the aft blower must be op-erating during electric heat operation. An ELECHEAT ON advisory annunciator is provided toindicate that the power relays are in the closedposition to apply electrical power to the heat-

Radiant Heating (Prior to BB-1444except BB-1439)An optional electric radiant heating system is

available for the Super King Air. This system isturned on or off by the RADIANT HEAT switch(Figure 11-10) located in the ENVIRONMENTALgroup on the copilot’s left subpanel. This systemuses overhead heating panels to warm the cabinprior to engine start, as well as to provide sup-plemental heat in flight (Figure 11-10). Duringground operations when using radiant heat, the useof an auxiliary power unit is highly encouraged.

Figure 11-9. ELECTRIC HEAT Switch

NOTE

The radiant heating system shouldbe used with the manual tempera-ture control mode only.

AIR-CONDITIONING SYSTEMCONTROLS

The ENVIRONMENTAL control sect ion(Figure 11-11) on the copilot’s left subpanelprovides automatic or manual control of theair-conditioning system. This section con-tains all the major controls of the environ-

mental function, which are:

• BLEED AIR VALVE switches.

• F o r wa r d V E N T B L O WE R c o n tr o lswitch.

Automatic Mode ControlWhen the CABIN TEMP MODE selector switch (Figure 11-12) is in the AUTO position,the air delivery system (Figure 11-13) oper-ates automatically to establish the temperaturerequested by the pilot. To reach the desiredtemperature setting, the automatic temperaturecontrol modulates the bypass valves and air conditioning compressor. For greater heating,bleed air is allowed to bypass the air-to-air heat exchangers in the wing center sections. For greater cooling, the bleed air is allowed to





• AFT (evaporator) BLOWER ON/OFFswitch (if installed).

• ELECTRIC HEAT switch (if installed,BB-1439, 1444 and subsequent).

• RADIANT HEAT switch (if installed,prior to BB-1444, except 1439).

• MANUAL TEMPerature switch.

• CABIN TEMPera ture l eve l con tro lswitch.

• CABIN TEMP MODE selector switch.Figure 11-12. CABIN TEMP MODE Selector

Switch



LEFT ENGINEBLEED AIR

RIGHT ENGINEBLEED AIR

MANCOOL

MANHEAT

AUTOOFF

CABIN TEMP MODE

MANUALTEMP

INCR

DECR

AUTO TEMPCONTROLLER

DUCTCABINSELECTOR

RH BYPASSVALVE MOTOR

LH BYPASSVALVE MOTOR

COOL

COOL

MANUAL

HEATOR COOL

HEAT

HEAT

AUTO

MANUALCOOL

TO CABIN

AIR TO AIRHEATEXCHANGER

TO CABIN

AIR TO AIRHEATEXCHANGER

TEMPSENSORS



pass through the air-to-air heat exchangers toreduce its temperature. In either case, the re-sultant bleed air is mixed with recirculatedcabin air (which can be additionally cooled if the air conditioning compressor is activatedin the cooling mode) in the forward mixingplenum.

The CABIN TEMP level control (Figure11-11) provides regulation of the temper-ature level in the AUTO mode. The pilot canadjust the temperature in the aircraft byturning the CABIN TEMP level control asrequired. A temperature-sensing unit behind

the first set of passenger oxygen masks (or BB-54 through BB-310, in the cockpit ceil-i d i t BB 54 i th l l ft

Manual Mode ControlWhen the CABIN TEMP MODE selector switch is in the MAN COOL or MAN HEATposition, regulation of the cabin temperatureis accomplished manually by momentarily

holding the MANUAL TEMP switch (Figure11-11) to either INCRease or DECRease po-sition as desired.

When released, this switch returns to the cen-ter (OFF) position. When held in either posi-tion, it results in modulation of the bypassvalves in the bleed-air lines. The pilot shouldallow one minute (30 seconds per valve) for both valves to move fully open or fully closed.Only one valve moves at a time to vary the

AIR CONDITIONER

LH BYPASSVALVE MOTOR

SWITCH

Figure 11-13. Air-Conditioning System Control Diagram

according to the position of the cabin-heatcontrol valves whether or not the refrigerantsystem is working.

NOTE

The air-conditioner compressor doesnot operate unless the bypass valvesare closed. To ensure that the valvesare closed, select MAN COOL thenhold the MANUAL TEMP switch inthe DECR position for one minute.

Airflow ControlFour additional manual controls on the sub-panels may be used to partially regulate cock-pit comfort when the cockpit partition door isclosed and the cabin comfort level is satis-factory (Figure 11-7). These controls are:

Vent Blower ControlThe VENT BLOWER switch located in theENVIRONMENTAL group, controls the for-ward vent blower (Figure 11-11). This switchhas three positions, HI–LO–AUTO. When itis in the AUTO position, the blower operatesat a low speed if the CABIN TEMP MODE se-lector switch is in any position other thanOFF. When the VENT BLOWER switch is inAUTO and the mode selector is in the OFF po-sition, the blower ceases to operate. Any timethe blower switch is in the LO position, theblower operates at low speed, even if the modeselector is in the OFF position. Likewise, if theblower switch is in the HI position, the blower operates at high speed even if the mode selector is in the OFF position.

If the optional aft evaporator unit is installedi th i l ft bl i l i t ll d





y ( g )

1. PILOT AIR CONTROL KNOB

2. DEFROST AIR CONTROL KNOB

3. CABIN AIR CONTROL KNOB

4. COPILOT AIR CONTROL KNOB

When these control knobs are fully pulled out,they provide maximum airflow to the cockpit;when fully pushed in, they provide minimum

airflow. During flights in warm air, such asshort, low-altitude flights in the summer, allthe cabin ceiling outlets should be fully openfor maximum cooling. During high-altitudeflights, cool-night flights, and flights in coldweather, the ceiling outlets should be closedfor maximum cabin heating.

Bleed-Air ControlThe BLEED AIR VALVE switches control the

in the airplane, an aft blower is also install edunder the floor next to the evaporator in therear of the cabin. The aft blower, which drawsin cabin air, blows it across the evaporator and to the aft floor and ceiling outlets; it op-erates at high speed only. The AFT BLOWERswitch (Figure 11-11), located in the ENVI-RONMENTAL group, controls the blower,which is independent of any other control.

NOTE

If the aft blower is turned on duringthe heating mode of operation, thedoor between the aft-blower duct andthe warm air (floor-outlet) duct opens.This stops the flow of bleed air to theaft floor registers and delivers recir-culated cabin air (which comes fromunder the floor and will be cooler than cabin air) to the aft floor regis-ters and ceiling outlets (DETAIL A,Figures 11-1 and 11-2). For BB-1439,1444 and subsequent both the vent

On Super King Air 200 airplanes, serials prior to BB-39, some airplanes were delivered witha two-speed aft blower which did not have a sep-arate AFT BLOWER switch, but was controlledby the forward VENT BLOWER switch and a

temperature sensor. In this installation, aftblower operation is entirely automatic and can-not be controlled by the pilot (Figure 11-11).

Unpressurized VentilationFresh air is available during unpressurized flightwith the CABIN PRESS switch in the DUMPposition. This ambient (ram) air is obtained via

the fresh air door and the ram-air scoop in theairplane nose section (Figure 11-14). This door is open only during unpressurized flight whenthe switch is in the DUMP position and there is0 psid. This allows the forward blower to drawram air into the cabin. This air is mixed with re-

NOTE

A flight conducted with the bleed-air switches placed in any position other than OPEN will also result in un-pressurized flight, but the fresh air

door will not be open.

LIMITATIONS

The following limits are imposed upon theair-conditioning system:

• Underpressure limit—2.5 psi (whichdisengages the air-conditioning clutch).

• Overpressure limit, forward evapora-tor—290 psi (which disengages the com-pressor clutch).

O li it ith ft t





circulated cabin air in the plenum chamber andthen directed to both the floor registers andceiling outlets.

On early Super King Air 200 models, the vol-ume of air from the registers is regulated by mov-ing a sliding handle (lever) at the side of eachinboard-facing register. On BB 310, 345 andafter on the Super King Air B200 the air vol-ume is regulated by the CABIN AIR controlknob (Figure 11-7).

• Overpressure limit with aft evaporator —340 psi (which disengages the com-pressor clutch).

NOTE

Prior to Serial No . 345, a 50°F OATswitch is located on top of the con-denser for auto only:

• OAT above 50°F, condenser motor runs.

• OAT below 50°F, condenser motor stops.

• After Serial No. 345, the nose gear limitswitch stops the condenser motor whenthe gear is up.

• Air-conditioning system rated at 18,000BTU with forward blower only.

• Air-conditioning system rated at 32,000BTU with aft blower and forward blower

operating.• AIR CND N1 LOW illuminates if right

1. When the engine speed falls below 62% N1the compressor clutch disengages and the

green advisory annunciator illuminates.A. AIR CND N1 LOW

B. DUCT OVERTEMP

C. AIR CND LOW

D. ENG SPD N1 LOW

2. How much of the recirculated air passes

through the ceiling outlet ducts?A. 75%

B. 50%

C. 25%

D. 60%

6. On the Super King Air B200, the air vol-ume passing through the floor registers is

controlled by:A. Sliding handle

B. CABIN AIR knob

C. Adjusting the engine N1 speed

D. Radiant heat switch

7. What is the source of fresh air during un-

pressurized flight with the PRESS switchin the DUMP position?

A. Ram air

B. Ram air, bleed-air heating system

C. Refrigerant air, ram air

D Refrigerant air bleed air heating



QUESTIONS



3. How is the airstream adjusted on the “eye-ball” outlets?

A. By twisting the nozzle

B. By pushing in the nozzle

C. By moving a sliding lever

D. By pos i t ioning VENT BLOWERswitch to LO

4. What is the airplane’s cooling capacity

with the aft evaporator, without the aftevaporator?

A. 1,800 BTU, 18,000 BTU

B. 3,200 BTU, 32,000 BTU

C. 1,800 BTU, 9,000 BTU

D. 32,000 BTU, 18,000 BTU

5. What control is adjusted if the bleed-air mixture is too warm for the crew?

D. Refr igerant air, bleed-air heat ingsystem

8. Prior to BB-1444, excep t 1439 whenshould the radiant heating system be used?

A. With the manual temperature controlmode

B. When the automatic temperature con-trol mode is used

C. Only when airborne

D. Whenever the aft blower is off

9. What adjustment is made if the cockpittemperature is too hot when the plane isbeing heated?

A. PILOT AIR, COPILOT AIR, DE-FROST AIR, and CABIN AIR knobsfully pushed in or as required

B. PILOT AIR, COPILOT AIR, and DE-FROST AIR knobs fully pulled out

C C k i h d “ b ll” l

10. When the CABIN TEMP MODE selector switch is in the MAN COOL position,how is the cabin temperature lowered?

A. Momentarily depressing the MAN-UAL TEMP switch to INCR

B. Momentarily depressing the MAN-UAL TEMP to DECR

C. Momentar i ly holding the CABINTEMP level control to DECR

D. Momentar i ly holding the CABINTEMP level control to INCR

11. How does the pilot ensure that the air-to-

air heat exchanger valves are closed?

A. Turn the CABIN TEMP selector allthe way clockwise

B. Momentarily place the CABIN TEMPMODE switch to MAN COOL

C S l MAN COOL h h ld h





C. Select MAN COOL, then hold theMANUAL TEMP switch in the DECRposition for one minute

D. Hold the MANUAL TEMP switch inthe INCR position for one minute

12. When does the door between the af t-blower duct and the warm duct open, thusstopping the flow of heated air to the aftfloor registers?

A. When the aft blower is turned on

B. If the aft blower is turned off duringthe heating mode of operation

C. When in manual temperature controlmode

D. During unpressurized flight

CHAPTER 12

PRESSURIZATION

CONTENTSPage

INTRODUCTION................................................................................................................. 12-1

GENERAL ............................................................................................................................ 12-1

System Description and Location.................................................................................. 12-3

Operation ....................................................................................................................... 12-5

Preflight Operation ........................................................................................................ 12-8

In-Flight Operation ........................................................................................................ 12-8





Descent and Landing Operation .................................................................................... 12-8

LIMITATIONS ...................................................................................................................... 12-9

QUESTIONS....................................................................................................................... 12-10

ILLUSTRATIONS

Figure Title Page

12-1 Pressurization Controls .......................................................................................... 12-2

12-2 Electronic Flow Control Unit(BB-1180 and Subsequent, BL-71 and Subsequent) ............................................. 12-3

12-3 Pneumatic Flow Control Unit (Prior to BB-1180, Prior to BL-71) ....................... 12-4

12-4 Outflow Valve ........................................................................................................ 12-6

12-5 Safety Valve ........................................................................................................... 12-6

12-6 Pressurization Controller........................................................................................ 12-7

12-7 Cabin Altimeter...................................................................................................... 12-7

12-8 CABIN CLIMB Indicator ...................................................................................... 12-7





12-9 CABIN PRESSURE Switch .................................................................................. 12-8

12-10 ALT WARN Annunciator....................................................................................... 12-8

TABLES

Table Title Page

12-1 Pressurization Controller Setting for Landing ....................................................... 12-9

CHAPTER 12PRESSURIZATION





INTRODUCTION

On Super King Air 200s, BB-2 through BB-194, the pressurization system is designedto provide a normal working pressure differential of 6.0 ± 0.1 psi, which provides cabinpressure altitudes of approximately 3,900 feet at an altitude of 20,000 feet, 9,900 feetat 31,000 feet, and 11,700 feet at 35,000 feet. The normal working pressure different ialfor Super King Air 200s, Serial Nos. BB-195 up to the B200, is 6.1 psi.

On Super King Air B200 airplanes, the pressurization system is designed to provide anormal working pressure differential of 6.5 ± 0.1 psi, which provides cabin pressure al-titudes of approximately 2,800 feet at 20,000 feet, 8,600 feet at 31,000 feet, and 10,400feet at 35,000 feet.



FILTER

FLOW CONTROL

PRESSURE

MOISTUREACCUMULATION

DRAIN

PLUG

STATIC

STATIC

OUT-FLOWVALVE

CABIN PRESETSOLENOID

NO LGSAFETYSWITCH



SAFETYVALVE

RATE ALTITUDE

RESTRICTOR

VACUUM SOURCEFROMPNEUMATICMANIFOLD

CONTROL SWITCHCABIN PRESSURE

DUMP SOLENOIDNC

LEGEND

CABIN AIR

VACUUM SOURCE

STATIC AIR



SYSTEM DESCRIPTION ANDLOCATION

Bleed air from each engine is used to pressurizethe pressure vessel (cabin and cockpit areas).A flow control unit in each of the engine na-celles controls the volume of the bleed air andcombines ambient air with it to provide a suit-able air density for pressurization. The BLEEDAIR VALVE switch in the ENVIRONMEN-TAL group on the copilot’s left subpanel con-trols the flow control unit (Figure 12-2). Whenthis switch is either in the ENVIR OFF or theINSTR & ENVIR OFF position, the flow con-

trol unit is closed. These switch positions willalso illuminate the green L or R BL AIR OFFadvisory annunciator light. When it is in theOPEN position, the mixture of engine bleedair and ambient air flows through the flowcontrol unit and through or around the air-to-air heat exchangers.

cabin (Figure 12-2). Each unit consists of anambient temperature sensor, an electronic con-troller, and an environmental air control valveassembly, interconnected by a wire harness.

The control valve assembly consists of:

• Mass flow transducer

• Ambient flow motor and modulatingvalve

• Check valve that prevents the bleed air from escaping through ambient air intake

• Bleed air flow transducer

• Bleed air flow motor and modulatingvalve (including bypass line)

• Air e jector

• Flow control solenoid valve

• Environmental shutoff valve



Electronic Flow Control Unit(BB-1180 and Subsequent, BL-71 and Subsequent)Electronic flow control units control the massflow of both ambient and bleed air into the

After engine start up when the flow control unitis energized, the bleed air modulating valve

closes. When it is fully closed, it actuates thebleed-air shaft switch, signaling the electroniccontroller to open the solenoid valve. Thisenables P3 bleed air to pressurize the envi-ronmental shutoff valve, causing it to open.

BLEED-AIR

FLOW TRANSDUCERPOWER

SQUAT

SWITCH

AMBIENT

AIR

INLET

AMBIENT

FLOW CONTROL

MOTOR

ELECTRONIC

CONTROLLER

BLEED-AIR

FLOW CONTROL

MOTOR

SOLENOID (N.C.)

ENVIRONMENTALSHUTOFF

VALVE (N C )

COCKPIT BLEED AIR

VALVE SWITCHFIRESEAL

AMBIENT

TEMPERATURE

SENSOR

The bleed-air shaft continues to open until thedesired bleed-air flow rate to the cabin isreached. (The flow rate is sensed by the bleed-air flow transducer and controlled by the elec-tronic controller per the input of the ambienttemperature sensor.)

As the airplane enters a cooler environment,ambient airflow is gradually reduced andbleed-air flow gradually increased to maintaina constant inflow and to provide sufficientheat for the cabin. At approximately 0°C am-bient temperature, ambient airflow is com-pletely closed off and the bleed-air valvebypass section is opened, as necessary, toallow more bleed air flow past the fixed flowpassage of the air ejector.

Flow Control Unit (Prior toBB-1180, Prior to BL-71)Each flow control unit (Figure 12-3) consistsof an ejector and an integral bleed-air modu-lating valve, firewall shutoff valve, ambientair modulating valve, and a check valve thatprevents the bleed air from escaping throughthe ambient air intake. The flow of bleed air through the flow control unit is controlled asa function of atmospheric pressure and tem-perature. Ambient airflow is controlled as afunction of temperature only. When the bleed-air valve switches on the copilot’s left subpanelare in the open position, a bleed-air shutoff electric solenoid valve on each flow controlunit opens to allow the bleed air into the unit.



PRESSUREREGULATOR

PNEUMOSTAT(PNEUMATIC

THERMOSTAT)



AMBIENTFLOW

N.O.AMBIENT AIRMODULATINGVALVE

N.O.SOLENOID VALVE

TOCABIN

AIR TOAIR HEATEXCHANGER

BYPASSVALVE AMBIENT

SENSEANEROID

BYPASSVALVE

N.C.SOLENOID

REGULATOR

EJECTORFLOW

CONTROLACTUATOR

CHECKVALVE

EJECTOR

TO LH L.G.SAFETYSWITCH

FILTER

N.C.FIREWALLSHUT-OFF

VALVE

T O O

P E N

TO CLOSETO OPEN



As the bleed air enters the flow control unit,it passes through a filter before going to thereference pressure regulator. The regulator will reduce the pressure to a constant value (18to 20 psi). This reference pressure is then di-rected to the various components within theflow control unit that regulate the output to thecabin.

One reference pressure line is routed to thefirewall shutoff valve located downstream of the ejector. A restrictor is placed in the line im-mediately before the shutoff valve to providea controlled opening rate. At the same time,the reference pressure is directed to the am-bient air modulating valve, located upstreamof the ejector, and to the ejector flow controlactuator.

A pneumatic thermostat with a variable ori-fice is connected to the modulating valve. Thepneumatic thermostat (pneumostat) is located

another variable orifice of the pneumatic ther-mostat and a variable orifice controlled by anisobaric aneroid. The pneumostat orifice isrestricted by decreasing ambient temperature,and the isobaric aneroid orifice is restricted bydecreasing ambient pressure. The restrictionof either orifice will cause a pressure buildupon the ejector flow control actuator, permit-ting more bleed air to enter the ejector.

OPERATION

The flow control units regulate the rate of air-flow to the pressure vessel. The bleed air por-

tion is variable from approximately 5 to 14pounds per minute depending upon ambienttemperature. On the ground, since ambientair is not available, cabin inflow is variable andlimited by ambient temperature. Inflight, am-bient air provides the balance of the cons tantairflow volume of 12 to 14 pounds per minute.



pneumatic thermostat (pneumostat) is locatedon the lower aft side of the fireseal forward of the firewall. The bi-metallic sensing discs of the thermostat are inserted into the cowling in-take. These discs sense ambient temperatureand regulate the size of the thermostat ori-fices. Warm air will open the orifice and coldair will restrict it until, at minus 30°F, the ori-fice will completely close. When the variableorifice is closed, the pressure buildup willcause the modulating valve to close off the am-bient air source. An ambient air shutoff valve,located in the line to the pneumatic thermo-stat is wired to the left landing gear safetyswitch. When the airplane is on the ground, thissolenoid valve is closed, thereby directing thepressure to the modulating valve, causing it toshut off the ambient air source. The exclu-sion of ambient air allows faster cabin warm-up during cold weather operation. An electriccircuit containing a time delay relay is wiredto the above mentioned solenoid valves toll h l f l l

From here, the air, which is also being used for

cooling and heating, flows into the pressurevessel, creating differential, and out throughthe outflow valve (Figure 12-4) located onthe aft pressure bulkhead. To the left of the out-flow valve (looking forward) is a safety valve(Figure 12-5) which provides pressure relief if the outflow valve fails, depressurizes the air-plane whenever the CABIN PRESS switch ismoved into the DUMP position, and keeps the

airplane unpressurized while it is on the groundwith the left landing-gear safety switch com-pressed. A negative-pressure relief function,which prevents outside atmospheric pressurefrom exceeding cabin pressure by more than0.1 psi during rapid descents with or withoutbleed air flow, is also incorporated into bothvalves.

When the BLEED AIR VALVE switches arein the OPEN position, the air mixture (bleed



SCHRADER

VALVE

PLUG

STATIC AIR

MAXIMUM

DIFFERENTIAL

DIAPHRAGM

TO CONTROLLERCONNECTION

UPPER

(CONTROL)

DIAPHRAGM

NEGATIVE

RELIEF

DIAPHRAGM

REAR

PRESSURE

BULKHEAD



SCHRADER

VALVE MAXIMUM

DIFFERENTIAL

DIAPHRAGM

SAFETY VALVE DUMP SOLENOID

UPPER DIAPHRAGM

NEGATIVE RELIEF

DIAPHRAGM

CABIN

AIR

Figure 12-4. Outflow Valve



The CABIN PRESSure switch (Figure 12-9),located to the left of the pressurization con-troller, in the DUMP (forward lever locked)position, opens the safety valve, allowing thecabin to depressurize and stay unpressurizeduntil the switch is placed in the PRESS (cen-ter) position. In the PRESS position, the safetyvalve closes and the pressurization controller takes command of the outflow valve. In theTEST (aft, spring-loaded to the center) posi-tion, the safety valve is held closed, bypass-ing the landing gear safety switch to allowcabin pressurization tests on the ground.

ferential reaches the pressure relief setting of the outflow valve and the safety valve. Either or both valves then override the pressure con-troller in order to limit the cabin to ambient pres-sure differential to the normal working pressuredifferential previously stated. If the cabin pres-sure altitude should reach a value of 12,500 feet,a pressure-sensing switch on the forward pres-sure bulkhead closes, thus illuminating the redALT WARN annunciator light , (Figure 12-10),warning the pilot of operation requiring oxy-gen use. If the auto deployment oxygen systemis installed, a pressure-sensing switch in thecabin wall (copilot’s side) forward of the emer-gency exit also closes, deploying the passen-ger oxygen masks to face level. During cruiseoperation, if the flight plan requires an alti-tude change of 1,000 feet or more, the CABINALT dial should be readjusted.





PREFLIGHT OPERATION

Prior to takeoff, the cabin altitude selector knobis adjusted until the ACFT ALT (inner) scaleon the indicator dial reads an altitude approx-imately 500 feet or 1,000 feet above the plannedcruise pressure altitude. The RATE controlknob is adjusted as desired. When the indexmark is set between the 9 o’clock and 12 o’clockpositions, the most comfortable rate of climbis maintained. The CABIN PRESSure switchis placed in the PRESSure position.

IN-FLIGHT OPERATION

As the airplane climbs, the cabin pressure alti-

DESCENT AND LANDINGOPERATION

During descent and in preparation for land-ing, the cabin altitude selector is set to in-dicate a cabin altitude of approximately 500feet above the landing field pressure alti-tude (Table 12-1). Also, the RATE controlknob is adjusted as required to provide acomfortable cabin altitude rate of descent.The airplane rate of descent is controlle d sothe airplane altitude does not cat ch up with

the cabin pressure altitude until the cabinpressure altitude reaches the selected value

Figure 12-9. CABIN PRESSURE Switch

Figure 12-10. ALT WARN Annunciator

LIMITATIONS

The following limitations have been imposedon the pressurization system:

• CABIN DIFFERENTIAL PRESSUREGAGE (B200)

Green Arc (Approved Operating Range)0 to 6.6 psi

Red Arc (Unapproved Opera t ingRange) 6.6 psi to end of scale

• CABIN DIFFERENTIAL PRESSUREGAGE (200; BB-195 and after)

Green Arc (Approved Operating Range)0 to 6.1 psi

Red Arc (Unapproved Operating Range)6.1 psi to end of scale

• CABIN DIFFERENTIAL PRESSUREGAGE (200; prior to BB-195)



CLOSEST ADD TO

ALTIMETER SETTING AIRPORT ELEVATION

28.00 .......................... + 2,400

28.10 .......................... + 2,300

28.20 .......................... + 2,200

28.30 .......................... + 2,100

28.40 .......................... + 2,000

28.50 .......................... + 1,900

28.60 .......................... + 1,80028.70 .......................... + 1,700

28.80 .......................... + 1,600

28.90 .......................... + 1,500

29.00 .......................... + 1,400

29.10 .......................... + 1,300

29 20 1 200

Table 12-1. PRESSURIZATIONCONTROLLER SETTINGFOR LANDING



GAGE (200; prior to BB 195)

Green Arc (Approved Operating Range)

0 to 6.0 psi

Red Arc (Unapproved Operating Range)6.0 psi to end of scale

• MAXIMUM OPERATING PRESSURE-ALTITUDE LIMITS

Normal Operation .......... 35,000 feet

• MAXIMUM OPERATING PRESSURE-

ALTITUDE LIMITS (Prior to BB-54,except 38, 39, 42, and 44)

Normal Operation .......... 31,000 feet

29.20 .......................... + 1,200

29.30 .......................... + 1,100

29.40 .......................... + 1,000

29.50 . .. .. . .. .. .. .. . .. .. .. . .. .. + 900

29.60 . .. .. . .. .. .. .. . .. .. .. . .. .. + 800

29.70 . .. .. . .. .. .. .. . .. .. .. . .. .. + 700

29.80 . .. .. . .. .. .. .. . .. .. .. . .. .. + 600

29.90 . .. .. . .. .. .. .. . .. .. .. . .. .. + 500

30.00 . .. .. . .. .. .. .. . .. .. .. . .. .. + 400

30.10 . .. .. . .. .. .. .. . .. .. .. . .. .. + 300

30.20 . .. .. . .. .. .. .. . .. .. .. . .. .. + 200

30.30 . .. .. . .. .. .. .. . .. .. .. . .. .. + 100

30.40 . .. .. .. .. .. .. .. .. .. .. .. .. .. . 0

30.50... . .. .. .. .. . .. .. .. .. . .. .. . - 100

30.60... . .. .. .. .. . .. .. .. .. . .. .. . - 200

30.70... . .. .. .. .. . .. .. .. .. . .. .. . - 300

30.80... . .. .. .. .. . .. .. .. .. . .. .. . - 400

30.90... . .. .. .. .. . .. .. .. .. . .. .. . - 500







The information normally contained in this chapter is notapplicable to this particular aircraft.

CHAPTER 14

LANDING GEAR AND BRAKES

CONTENTSPage

INTRODUCTION................................................................................................................. 14-1

LANDING GEAR (ELECTRIC) .......................................................................................... 14-2

General........................................................................................................................... 14-2

Gear Assemblies ............................................................................................................ 14-2

Wheel Well Door Mechanisms ...................................................................................... 14-2

Controls.......................................................................................................................... 14-4

Indicators ....................................................................................................................... 14-5





Indicators ....................................................................................................................... 14 5

Warning System............................................................................................................. 14-6

Operation ....................................................................................................................... 14-7

LANDING GEAR (HYDRAULIC)...................................................................................... 14-9

General........................................................................................................................... 14-9

Gear Assemblies .......................................................................................................... 14-10

Wheel Well Door Mechanisms.................................................................................... 14-13

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

Indicators ..................................................................................................................... 14-15

Operation ..................................................................................................................... 14-17

NOSEWHEEL STEERING................................................................................................ 14-21

BRAKE SYSTEM............................................................................................................... 14-21

Operation ..................................................................................................................... 14-21

Care and Handling in Cold Weather............................................................................ 14-21

Main Gear Safety Switches ......................................................................................... 14-22

LIMITATIONS.................................................................................................................... 14-23

Air Speed Limitations.................................................................................................. 14-23

QUESTIONS....................................................................................................................... 14-25





ILLUSTRATIONS

Figure Title Page

14-1 Nose Gear Assembly.............................................................................................. 14-2

14-2 Main Gear Assembly.............................................................................................. 14-3

14-3 Main Gear Door (Standard Gear)........................................................................... 14-4

14-4 Main Gear Door (High Flotation Gear) ................................................................. 14-4

14-5 Landing Gear Switch Handle and Indicator Lights................................................ 14-5

14-6 Normal Indications Gear Down............................................................................. 14-5

14-7 Nose Gear Not Fully Extended .............................................................................. 14-6

14-8 Normal Indications, Gear Up................................................................................. 14-6

14-9 One or More Gear Not Fully Retracted.................................................................. 14-6

14 10 N l L di G O ti 14 8





14-10 Normal Landing Gear Operation ........................................................................... 14-8

14-11 Hydraulic Power Pack............................................................................................ 14-9

14-12 Components Locations......................................................................................... 14-10

14-13 Hydraulic Landing Gear System.......................................................................... 14-11

14-14 Nose Gear Assembly............................................................................................ 14-11

14-15 Internal Nose Gear Lock...................................................................................... 14-1214-16 Main Gear Assembly ........................................................................................... 14-13

14-17 Main Gear Door Mechanism (Standard Gear).................................................... 14-14

14-18 Main Gear Door Mechanism (High-Flotation Gear) ........................................... 14-14

14-19 Landing Gear Control Handle and Indicator Lights ............................................ 14-15

14-20 Normal Indications Gear Down........................................................................... 14-15

14-24 Normal Retraction................................................................................................ 14-18

14-25 Normal Extension ................................................................................................ 14-19

14-26 Alternate Extension.............................................................................................. 14-20

14-27 Brake System Schematic (Serial Nos. BB-666 and Subsequent) ........................ 14-22

14-28 Brake System Schematic (Serial Nos. BB-453 through BB-665) ....................... 14-23

14-29 Brake System Schematic (Serial Nos. BB-2 through BB-452) ........................... 14-24





CHAPTER 14LANDING GEAR AND BRAKES





INTRODUCTIONThe tricycle landing gear on the Super King Air 200 is actuated either by an electric motor or an electrically-driven hydraulic pump. The gear is controlled with a landing gear con-trol switch handle on the pilot’s right subpanel. On the electrically-actuated gear, motor torque is mechanically transmitted for gear extension and retraction. On the hydraulicgear, three hydraulic actuators provide motive power for gear operation.

Individual gear position lights provide gear position indication and two red lights in thegear control handle. In addition, a warning horn sounds if all three gears are not down and

LANDING GEAR(ELECTRIC)

GENERAL

The landing gear is actuated by a 28-VDCmotor powered by the right generator bus.The motor operates torque tubes and a cha indrive to transmit power to a mechanical actu-ator at each gear. A circuit breaker or currentlimiter and a spring-loaded friction clutch,prevents damage to the motor that may resultfrom a mechanical malfunction.

GEAR ASSEMBLIES

DescriptionThe landing gear assemblies (main and nose)consist of shock struts, torque knees (scissors),drag braces, actuators, wheels and tires, brake

bli d hi d B k

The shimmy damper, mounted on the rightside of the nose gear strut, is a balanced hy-draulic cylinder that bleeds fluid through anorifice to dampen nose wheel shimmy.

The jackscrew actuators retract and extend

the gear and provide a gear uplock due to fric-tion.

WHEEL WELL DOORMECHANISMS

Landing gear doors are mechanically actu-ated by gear movement during extension and

retraction. On airplanes configured with thestandard main gear, rollers on the shock strutcontact cams in the wheel well during re-traction (Figure 14-3).

Cam movement is transmitted through link-age to close the doors. During extension,roller action reverses cam movement to openth d Wh th ll h l ft th





assemblies, and a shimmy damper. Brake as-

semblies are located on the main gear assem-blies; the shimmy damper is mounted on thenose gear assembly (Figures 14-1 and 14-2).

OperationThe upper end of the drag braces and twopoints on the shock struts are attached to theairplane structure. When the gear is extended,

the drag braces are rigid components of thegear assemblies.

Airplane weight is borne by the air charge in theshock struts. At touchdown, the lower portionof each strut is forced into the upper cylinder;this moves fluid through an orifice, further com-pressing the air charge and thus absorbing land-ing shock.

A t k t th d l

the doors. When the rollers have left the

cams, springs drive the linkage overcenter tohold the doors open.

TORQUE

KNEE

SHOCKSTRUT

DRAG

BRACE

ROLLER

(NOSEWHEEL

DOOR)

PIVOT

POINTSHIMMY

DAMPER

DRAG

BRACE





SHOCK

STRUT

TORQUE

KNEE

PIVOT

POINT





On airplanes configured with the high-flota-tion gear, the main gear wheels are larger andthe shock strut shorter than on the standardgear. Since the wheels will not retract com-

pletely into the wheel well, a cutout in thedoors allows part of the wheel to protrude intothe airstream by approximately five inches. Onairplanes so configured, the main gear doorsare mechanically linked to the shock strut andare opened and closed as the gear extends or retracts (Figure 14-4).

Nose gear doors on airplanes with standard or

high-flotation gear are mechanically actuatedin the manner previously described for stan

CAM SPRING

Figure 14-3. Main Gear Door (Standard Gear)

three green gear position lights adjacent tothe switch handle, and two red lights to illu-minate the handle (Figure 14-5).

The switch handle is detented in both the UPand DN positions. A solenoid-operated down-

lock latch (commonly referred to as the “J”hook) engages the handle when the airplaneis on the ground, preventing inadvertent move-ment of the handle to the UP position. Whenairborne, the safety switch on the right maingear completes circuitry to disengage the han-dle latch, and the handle can be positioned toUP. A DN LOCK REL button to the left of thehandle, when pressed, releases the downlock

latch whether the airplane is on the ground or in flight (Figure 14-5).

INDICATORS

When the gear down cycle begins, the redlights illuminate the switch handle. As eachgear locks down, the corresponding green

If any gear does not lock down during exten-sion, its corresponding green light will not beon and the red handle will remain illuminated(Figure 14-7).



WARN HORN

HD LTTEST

LDG GEAR CONTROL

DOWNLOCK REL

SILENCE

UP

DN

LANDINGGEAR

RELAY

5

P R E S

S T O T E

DIM

P R E S S T O

T E

D IM

P R E

S S T O T E

DIM



g p g glight comes on. When all three gear are downand locked, all three green lights a re on andthe handle illumination ceases (Figure 14-6).

DOWNLOCK REL

HD LTTEST

LDG GEAR CONTROL

UP

DN

LANDING

RED

LIGHT

CAP

DOWN

LOCKRELEASE

SWITCH

HANDLE

GREEN

POSITIONLIGHTS

P R E T E

DIM

P R ES O

T E

D IM

P R E

S S T O T E

DIM

Figure 14-6. Normal Indications Gear Down



Super King Air 200, BB-453 andSubsequent, and BL-1 andSubsequent

Super King Air B200/B200C

With the flaps in the UP or APPROACH po-sition and either or both power levers retardedbelow a certain power level, the warning hornwill sound intermittently and the switch han-dle lights will illuminate. The horn can be si-l enced by p ress ing the WARN HORNSILENCE button; the lights in the switch han-dle cannot be cancelled. The warning systemwill be rearmed if the power lever(s) are ad-

vanced sufficiently.

Super King Air 200, Prior to BB-324

With the flaps in the APPROACH positionand either or both power levers retarded belowa certain power level, the warning horn andswitch handle lights will be activated and nei-ther can be cancelled

OPERATION

NormalPull the LDG GEAR CONT switch handle outof detent and position it to UP or DN, as ap-

plicable. This applies DC power from the rightgenerator bus to the applicable field windingof the landing gear motor (Figure 14- 10).

As the motor operates, torque tubes and theduplex chain arrangement from the motor gearbox drive the main and nose gear actu-ators to extend or retract the gear. In addi-tion, a 200-amp remote circuit breaker (Serial

Nos. BB-2 through BB-185) or a 150-amperecurrent limiter (Serial Nos. BB-186 throughBB-1192) protects the motor from overload.

When full extension or retraction has beenachieved, a dynamic brake relay, controlledby up or down limit switches, simultane-ously breaks the motor circuit and completesa circuit through the armature and the unused





ther can be cancelled.

Super King Air 200, BB-324 ThroughBB-452

With the flaps in the APPROACH position or beyond, the switch handle lights will illuminateand, if the airspeed is below 140 knots, thewarning horn will sound intermittently. Neither the horn nor the lights can be cancelled.

Super King Air 200, Prior to BB-324, BB-453 and Subsequent,and BL-1 and Subsequent

Super King Air B200/B200C

With the flaps beyond the APPROACH posi-tion, the warning horn and the switch handlelights will be activated regardless of the power

a circuit through the armature and the unused

field winding to stop the moto r.

Friction in the jackscrew assembly in each ac-tuator holds the gear in the retracted position.

The nose gear is locked down by an over-center condition of the drag brace. The maingears are locked down by a notched hookand plate attachment on the drag braces.

EmergencyReduce speed to 130 knots, place the LDGGEAR CONT switch handle to the DN posi-tion, and pull the LANDING GEAR RELAYcircuit breaker (Figure 14-10).



CROSS-SHAFT

SPRING-L AD

EIDLE S

GEAR BOX

M

A



DETAILA

LANDING GEARMANUAL EXTENSION SYSTEM

Pull up on the emergency engage handle lo-cated on the floor aft or to the left of thepedestal, and turn it clockwise to the stop.This disconnects electrical power from themotor and engages the emergency drive sys-tem. Pump the extension handle until the three

green gear position indicator lights come on.Additional pumping could bind the drivemechanism and prevent subsequent retrac-tion; however, if the green indicator lights donot come on, continue pumping until a defi-nite resistance is felt.

After an emergency landing gear ex-tension has been made, do not stowthe extension handle or move anylanding gear controls or reset anyswitches or circuit breakers until theairplane is on jacks. These precau-tions are necessary because the fail-ure may have been in the gear-up

If a practice emergency extension is made, thegear can be retracted electrically. Rotate theemergency engage handle counterclockwiseand push it down. Stow the extension handleand reset the LANDING GEAR RELAY circuitbreakers. Place the LDG GEAR CONT switch

handle in the UP position to retract the gear.

LANDING GEAR(HYDRAULIC)

GENERAL

On airplanes Serial Nos. BB-1193, BL-73,and subsequent, the landing gear is actuatedby a hydraulic power pack (Figure 14-11).The pack consists mainly of a 28-VDC motor-driven hydraulic pump, a hydraulic reser-voir pressurized by engine ble ed air, filters,a solenoid-operated selector valve, and an up-lock pressure switch. Adjacent to the pack isa service valve used for hand pump actuation

WARNING





ure may have been in the gear up

circuit, in which case, the gear mightretract on the ground. The gear can-not be retracted manually.

a service valve used for hand pump actuationof the gear during ground maintenance op-erations. Figure 14-12 shows the power packand components locations.

SYSTEM FILTER

TO NORMAL EXTEND

SIDE OF SYSTEM

FROM THE EMERGENCY

EXTEND SIDE OF SYSTEM

TO HAND

PUMP

FILTER

28-VDC PUMP

MOTOR

RESERVOIR

FLUID LEVEL SENSORSELECTOR

VALVE

SOLENOID

UPLOCK PRESSURE

SWITCH

TO NORMALRETRACT SIDEOF SYSTEM



NOSE GEARACTUATOR

SERVICEVALVE

OVERBOARDBLEED AIR

VENT

ACCUMULATOR

MAIN GEARACTUATOR

BLEED AIRREGULATOR

POWER PACKASSEMBLY

MAIN GEARACTUATOR

CHECKVALVE

FILL RESERVOIR



The power pack reservoir, serviced with MIL-H-5606 hydraulic fluid, is divided into two sec-tions. One section supplies the electrically-driven hydraulic pump, and the other sectionsupplies the hand pump. A fill reservoir justinboard of the left nacelle and forward of the

main spar (Figure 14-12) features a cap anddipstick assembly for maintaining system fluidlevel.

When reservoir fluid level is low, a sensor onthe reservoir completes a circuit to illuminatean amber HYD FLUID LOW annunciator.Pressing the HYD FLUID SENSOR TESTbut ton on the p i lo t ’ s subpanel tes ts the

annunciator.

GEAR ASSEMBLIES

DescriptionThe landing gear assemblies (main and nose)consist of shock struts, torque knee (scis-

sors), drag braces, actuators, wheels andt i r es , b r ake assembl ies , and a sh immydamper. Brake assemblies are located on themain gear assemblies; the shimmy damper ismounted on the nose gear assembly (Figures14-14 and 14-15).

Operation

The upper end of the drag braces and two

Figure 14-12. Components Locations



NOSEGEAR

ACTUATOR

HAND PUMP

LH MAIN GEAR

ACTUATOR

RH MAIN GEARACTUATOR

HYDRAULICPOWER

PACK

PLUMBING NETWORKFROM POWER PACK



Airplane weight is borne by the air charge inthe shock struts. At touchdown, the lower por-tion of each strut is forced into the upper cylin-der; this moves fluid through an orifice, further compressing the air charge and thus absorb-ing landing shock. Orifice action also reduces

bounce during landing.

At takeoff, the lower portion of the strut ex-tends until an internal stop engages.

A torque knee connects the upper and lower portion of the shock struts. It allows strut com-pression and extension but resists rotationalforces, thereby keeping the wheels aligned

with the longitudinal axis of the airplane. Onthe nose gear assembly the torque knee also

Figure 14-13. Hydraulic Landing Gear System

TORQUE

KNEE

SHOCK

STRUT

DRAG

BRACE

ROLLER

(NOSEWHEEL

DOOR)

PIVOT

POINTSHIMMY

DAMPER



PISTON

INLETPORT

ACTUATORDOWNLOCKSWITCH(UNLOCKED)

LOCKSPRING

BALL

LOCK

LOCK

COLLAR



BALLLOCK

LOCKCOLLAR

INLETPORT

PISTON

UNLOCKED

ACTUATORDOWNLOCKSWITCH

A hydraulic actuator attached to the foldingdrag brace of each gear assembly providesmotive force for gear actuation. Nose gear downlocking is provided by an internal lockmechanism (Figure 14-15) in the hydraulicactuator and by the overcenter condition of the

drag brace.

The main gears are mechanically locked downby a notched hook and plate attachment on themain gear drag braces (Figure 14-16).

WHEEL WELL DOORMECHANISMS

Gear movement during extension and re-traction mechanically actuates landing gear doors. On airplanes configured with the stan-dard main gear, rollers on the shock strutcontact cams in the wheel well during re-traction (Figure 14-17).

Cam movement is transmitted through linkageto close the doors. During extension, roller ac-tion reverses cam movement to open the doors.When the rollers have left the cams, springsdrive the linkage overcenter to hold the doorsopen.

On airplanes configured with the high-flota-tion gear, the main gear wheels are larger andthe shock strut shorter than on the standardgear.Since the wheels will not retract com-pletely into the wheel well, a cutout in thedoors allows part of the wheel to protrude in tothe airstream by approximately five inches. Onairplanes so configured, the main gear doors

are mechanically linked to the shock strut andare opened and closed as the gear extends or retracts (Figure 14-18).

Nose gear doors on airplanes with standard or high-flotation gear are mechanically actuatedin the manner previously described for stan-dard main gear doors.









CONTROLS

The LDG GEAR CONT switch handle on thepilot’s right subpanel controls the landinggear. Gear position is indicated by three greengear position lights adjacent to the switch han-dle, and two red lights to illuminate the han-dle (Figure 14-19).

The switch handle is detented in both the UPand DN positions. A solenoid-operated down-lock latch (commonly referred to as the “J”hook) engages the handle when the airplane

CAM SPRING

Figure 14-17. Main Gear Door Mechanism (Standard Gear)

latch whether the airplane is on the ground or in flight (Figure 14-19). As an additionalsafety factor, control circuitry to the landinggear selector valve is complete only when the



DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEAR

DOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

RED

LIGHT

CAP

DOWN

LOCK

RELEASE

Figure 14-19. Landing Gear Control Handle and Indicator Lights

BEACON STROBE

LIGHTS

LDG GEAR CONTROL



main gear safety switches sense an airbornecondition.

INDICATORS

Landing gear position is indicated by an as-sembly of three green lights in a single unit tothe right of the LDG GEAR CONT switch

handle. Two red parallel-wired lights in thehandle illuminate to indicate that the gear isunlocked or in transit.

When the gear down cycle begins, the redlights illuminate the switch handle. As eachgear locks down, the corresponding greenlight comes on. When all three gear are downand locked, all three green lights are on, and

the handle illumination ceases (Figure 14-20)

When the gear up cycle begins, the handle il-luminates and the three green position lightsgo out The handle remains illuminated until

DOWNLOCK REL

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

TEST

UP

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

Figure 14-20. Normal IndicationsGear Down

Warning System

Th l di i i f



BEACON STROBE

TAILGEAR

LIGHTS

LDG GEAR CONTROL

UP

DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

UP

NOSE

L R

DN

LANDINGGEAR

2

OFF

Figure 14-21. Nose Gear Not Fully Extended

DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

HANDLE LIGHTS

TEST SWITCH

Figure 14-23. One or More Gear NotFully Retracted



If any gear fails to retract completely, the han-dle continues to be illuminated (Figure 14-23).

Pushing on the light capsule tests the green po-sition indicator lights. Test the handle illu-

mination by pressing the HDL LT TEST switch

The landing gear warning system consists of the red lights that illuminate the LDG GEARCONT switch handle and a warning horn thatsounds when the gear is not down and lockedduring certain flight regimes.

With the flaps in the UP or APPROACH po-sition and either or both power levers retardedbelow approximately 85% N1, the warning

horn will sound intermittently and the switchhandle lights will illuminate. The horn can besilenced by pressing the WARN HORN SI-LENCE button; the lights in the switch han-dle cannot be cancelled. The warning systemwill be rearmed if the power lever(s) are ad-vanced sufficient ly. With the flaps beyond theAPPROACH position, the warning horn andthe switch handle lights will be activated re-gardless of the power settings and neither

DOWNLOCK REL

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

TEST

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

Figure 14-22. Normal Indications, Gear Up

OPERATION

Normal RetractionWith the safety switches sensing an airbornecondition, moving the LDG GEAR CONT

switch handle UP completes circuits to thepump motor relay and the up solenoid of thegear selector valve (Figure 14-24).

Power to the pump motor relay pulls in 28VDC to the hydraulic pump motor in the power pack. The gear selector valve is energized tothe gear up position, directing fluid pressureto the retract side of all three gear actuators.

When retraction is complete (approximatelysix seconds), the gear actuators bottom out, andpressure increases rapidly. At 2,775 psi, theuplock pressure switch opens, breaking thecircuit to the pump motor relay, and the pumpmotor deenergizes.

Since there are no gear uplock mechanisms,pressure in the retract side holds the gear re-

t t d Wh t l k d th

The gear selector valve is spring-loaded tothe down position for fail-safe operation in theevent of electrical power loss.

Alternate Extension

In the event of electrical power loss or hy-draulic power pack malfunction, a hydraulichand pump is provided for an alternate gear ex-tension (Figure 14-26). The hand pump, lo-cated on the floor between the pilot’s rightfoot and the pedestal, is labeled LANDINGGEAR EMERGENCY EXTENSION.

To use the alternate extension system, pull

the LANDING GEAR RELAY circuit breaker on the gear control panel, and position theLDG GEAR CONT switch handle DN.

Remove the hand pump handle from the se-curing clip and actuate the hand pump until thethree green gear position lights (NOSE–L–R)illuminate. Place the pump handle in the downposition, and secure in the retaining clip.





tracted. When system leakage drops the pres-sure to 2,475 psi, the uplock pressure switchcloses to reestablish the power circuit to thepump. Automatic cycling of the pump main-tains pressure to keep the gear up and locked.

Normal ExtensionPlacing the LDG GEAR CONT switch handles

in the DN position completes a circuit to thedown solenoid of the gear selector valve andthrough any of three gear downlock switchesto the pump motor relay (Figure 14-25). Theenergized relay pulls in 28 VDC for operationof the hydraulic pump motor in the power pack.

The gear selector valve is energized to thedown position, routing pressure to the extend

side of all three gear actuators As each main

If the green gear position lights do notilluminate, continue pumping untilheavy resistance is felt to ensure thegear is down and locked. Then leavethe handle at the top of the stroke.

NOTE

The landing gear cannot be damagedby continued operation of the handpump.

WARNING

WARNING

1 4

- 1 8

F OR T RAI NI N G

P URP O S E S

ONL Y

S UP E R

KI N G

A I R

2 0 0 / B 2 0 0

P I L O T T R A

CURRENT

LIMITING

RESISTOR

28 VDC2A

DOWN

UP

SERVICE

VALVE

SWITCHES

RIGHT SAFETY

SWITCH

TIME DELAY

IN OUTLOGIC

RELAY

RIGHT

MAIN

NOSE

ACT

LEFT

MAIN

SERVICE

VALVE

HAND PUMP

PUMP

MOTOR

RELAY

60A

5A

CHECK VALVE

OVERBOARD

VENT

ORIFICE

FILL

RESERVOIR

FILL

PORT

HAND

PUMP

SUCTION

PORT

HANDPUMPPRESSUREPORT

SECONDARY

RESERVOIR

PRESSURESWITCH

HAND

PUMP

DUMP

VALVE

SOLENOID

1

2

PRESSURE

CHECKVALVE

ACCUMULATOR

REGULATED ENGNE

BLEED AIR (18 TO 20 PSI)

FILTER

FILTER

THERMAL RELIEF VALVE

VENT VALVE

SOLENOIDUP

AUXILIARY

PRESSURE

PORT

(PLUGGED)

GEAR

DOWN

PORT

GEAR UPPORT

PUMP MOTOR

PRIMARY

RESERVOIRRETURN

FILTER

VENT PORTAUXILIARY RETURN

PORT (PLUGGED) POWER PACK ASSEMBLY

PUMP

PUMP

CHECK

VALVE

SYSTEM

RELIEF

VALVE

PRESSURE FLUID

RETURN FLUID

NOTE:

THE INTERNAL SHUTTLE VALVE IS SPRING LOADED TO A POSITION WHICH ALLOWS FLUID IN THE ACTUATOR TO FLOW OUT THE NORMAL

EXTENDED PORT.

PRESSURE SWITCH CIRCUIT OPENS ON INCREASING PRESSURE AT 2,275 ± 55 PSIG AND CLOSES ON DECREASING PRESSURE AT A DIFFERENTIAL OF 300–400 PSIG.

LEGEND

CONTROLSWITCH

DOWN

FILTER

RELIEF

VALVE

28 VDC

ELECTRIC POWER

SELECTOR VALVE



AINI N G MA N U A L

F l i gh t S a

f e t y

i n t er n a t i on

al

MAIN

DOWNLOCK

SWITCHES

LEFT SAFETY

SWITCHES LEFT MAIN

ACTUATOR

NOSE

ACTUATORRIGHT MAIN

ACTUATOR

CO OSWITCH

Figure 14-24. Normal Retraction

F OR T R

AI NI N G P URP O S E S

O

N

S UP E R

KI N G

A I R

2 0 0 / B 2 0 0

P I L O

CURRENT

LIMITING

RESISTOR

28 VDC 2A

UP

SERVICE

VALVE

SWITCHES

RIGHT SAFETY

SWITCH

TIME DELAY

IN OUT

LOGIC

RELAY

RIGHT

MAIN

NOSESERVICE

VALVE

HAND PUMP

PUMP

MOTOR

RELAY

60A

5A

CHECK VALVE

OVERBOARD

VENT

ORIFICE

FILL

RESERVOIR

FILL

PORT

HAND

PUMP

SUCTION

PORT

HAND

PUMP

PRESSURE

PORT

SECONDARY

RESERVOIR

PRESSURE

SWITCH

HAND

PUMP

DUMP

VALVE

SOLENOID

1

2

PRESSURE

CHECK

VALVE

ACCUMULATOR

FILTER

FILTER


VENT VALVE

SOLENOIDUP

AUXILIARY

PRESSURE

PORT

(PLUGGED)

GEAR

DOWN

PORT

GEAR UPPORT

PUMP MOTOR

PRIMARY

RESERVOIRRETURN

FILTER


PORT (PLUGGED) POWER PACK ASSEMBLY

PUMP

PUMP

CHECK

VALVE

SYSTEM

RELIEF

VALVE

PRESSURE FLUID

RETURN FLUID

NOTE:

THE INTERNAL SHUTTLE VALVE IS

SPRING LOADED TO A POSITION WHICH

ALLOWS FLUID FROM GEAR DOWN

PORT OF POWER PACK TO

FLOW INTO ACTUATOR.

FLUID PRESSURE FROM

PUMP UNLOCKS VALVE.

LEGEND

DOWN

FILTER

RELIEF

VALVE

28 VDC

REGULATED ENGNE

BLEED AIR (18 TO 20 PSI)

ELECTRIC POWER

SELECTOR VALVE



1 4 - 1 9

NL Y

LOT T R A I NI N G MA N U

A L

F l i g

h t S af e t y


DOWN

NOSE

ACT

LEFT

MAIN

DOWNLOCK

SWITCHES

LEFT SAFETY

SWITCHES

VALVE

LEFT MAIN

ACTUATOR

NOSE

ACTUATORRIGHT MAIN

ACTUATOR

CONTROL

SWITCH

Figure 14-25. Normal Extension

1 4 - 2 0

F OR T RAI NI N G

P URP O S E S

ONL Y

S UP E R

KI N G

A I R

2 0 0 / B 2 0 0

P I L O T T R

A I N

CURRENT

LIMITING

RESISTOR

28 VDC2A

DOWN

UP

SERVICE

VALVE

SWITCHES

RIGHT SAFETY

SWITCH

TIME DELAY

IN OUT

LOGICRELAY

RIGHT

MAIN

NOSE

ACT

LEFT

SELECTOR VALVE

REGULATED ENGINE

BLEED AIR (18 TO 20 PSI)CHECK VALVE

ORIFICE

FILL

RESERVOIR

OVERBOARD

VENT28VDC

60A

5A

HAND PUMP

HAND

PUMP

SUCTION

PORT

SERVICE

VALVE

GEAR UP

PORTACCUMULATOR


VENT VALVE

FILTER


PORT (PLUGGED)POWER PACK ASSEMBLY

PRIMARY

RESERVOIR

PUMP

SYSTEM

RELIEF

VALVE

AUXILIARY

PRESSURE

PORT

(PLUGGED)

GEAR

DOWN

SUPPORT

HANDPUMP

PRESSURE

PORT

PRESSURE

SWITCH

HAND

PUMP

DUMPVALVE PRESSURE

CHECK

VALVE

DOWNSOLENOID

RETURN

FILTER

FILL

PORT

HAND PUMP PRESSURE FLUID

RETURN FLUID

HAND PUMP SUCTION

CONDITIONS:1. LANDING GEAR CONTROL HANDLE IN

"DOWN" POSITION

2. 2-AMPERE CONTROL CIRCUIT BREAKER

PULLED

NOTES:

PRESSURE FLUID FROM HAND PUMP

SHUTTLES INTERNAL SHUTTLE

VALVE TO ALLOW FLUID TO FLOW

INTO ACTUATOR.

HAND PUMP PRESSURE

FLUID UNSEATS VALVE.UP

SOLENOID

SECONDARY

RESERVOIR

PUMP

MOTOR

RELAY

PUMPCHECKVALVE

PUMP MOTOR

FILTERRELIEFVALVE

CONTROL

LEGEND

1

2



ANI N G MA N U A L

F l i gh t S a

f e t y

i n t er n a t i on

al

MAIN

DOWNLOCK

SWITCHES

LEFT SAFETY

SWITCHES

LEFT MAIN

ACTUATOR

NOSE

ACTUATOR

RIGHT MAIN

ACTUATOR

CONTROL

SWITCH

Figure 14-26. Alternate Extension

The landing gear cannot be retracted with thealternate extension system.

After a practice alternate extension, the gear may be retracted hydraulically by resettingthe LANDING GEAR RELAY circuit breaker

and moving the LDG GEAR CONT switchhandle to UP.

NOSEWHEEL STEERING

GENERAL

Direct linkage from the rudder pedals to an armnear the top of the shock strut mechanicallyactuates nosewheel steering. The steeringangle is from 14° left of center to 12° right of center, but can be considerably increased whenaugmented by differential braking and/or dif-ferential thrust.

OPERATION

Si ti f th dd d l i t

Pressure from the master cylinders is appliedto the brake assemblies. Each master cylinder supplies pressure to its set of brake assemblies;therefore, differential braking is available.

Prior to BB-666, the initial pressure from a set

of pedals will position a shuttle valve in thebraking system. Brake operation from the op-posite side can then only be accomplished bymoving the shuttle valve.

An optional brake deicing system using bleedair is provided for cold weather operation.This feature is covered in Chapter 10, ICEAND RAIN PROTECTION.

The pilot can set the parking brakes by ap-plying the brakes, then pulling out on thePARKING BRAKE handle on the pilot’s leftor right subpanel. The brakes can be releasedby applying toe pressure on the pedals, thenpushing in the PARKING BRAKE handle.Prior to BB-453, only the pilot can set theparking brake (Figure 14-29).

On some airplanes the PARKING BRAKE





Since motion of the rudder pedals is trans-mitted by cables and linkage to the rudder,deflection of the rudder occurs when force isapplied to any of the pedals. With the nose-wheel stationary on the ground or with theself-centering nose gear retracted, rudder pedalmovement compresses a spring-loaded linkin the system but it is not sufficient to steer the

nosewheel. If the nosewheel is on the groundand rolling, less force is required for steering;therefore, pedal deflection results in steeringthe nosewheel.

BRAKE SYSTEM

OPERATION

On some airplanes the PARKING BRAKEhandle is located on the pilo t’s right subpanel,below the LDG GEAR CONT switch handle.On these airplanes, either the p ilot or the copi-lot can set the parking brakes.

CARE AND HANDLING IN

COLD WEATHERPreflightCheck the brakes and the tire-to-ground con-tact for freeze lockup. Anti-ice solutions maybe used on the brakes and tires if freezeup oc-curs. No anti-ice solution, which contains a lu-bricant, such as oil, should be used on thebrakes. It will decrease the effectiveness of thebrake friction areas



OVERBOARD DRAIN

RESERVOIR

COPILOT’SMASTER

CYLINDER

RIGHT WHEEL BRAKELEFT WHEEL BRAKE

PILOT’SMASTER

CYLINDER

PARKING BRAKE



snow and slush can be forced into the brake

assemblies. Keep flaps retracted during taxi-ing to avoid throwing snow and slush into theflap mechanism and to minimize damage toflap surfaces.

Do not taxi with a flat shock strut.

Left Gear Safety Switch

• Safety valve

• Preset solenoid

• Dump solenoid

• Door seal solenoid

• Ambient air modulating valves

• Lift computer (stall warning)

• Stall warning heat control

CAUTION

RIGHT WHEEL BRAKE

Figure 14-27. Brake System Schematic (Serial Nos. BB-666 and Subsequent)



OVERBOARD DRAIN

RESERVOIR

LEFT WHEEL BRAKE

PILOT’SMASTERCYLINDER

PARKING BRAKE

SHUTTLEVALVE

SHUTTLEVALVE

PARKVALVE

PARKVALVE

RIGHT WHEEL BRAKE

COPILOT’SMASTERCYLINDER



Right Gear Safety Switch

• Landing gear handle lock solenoid

• Landing gear motor

• Landing gear emergency control

• Flight hourmeter

LIMITATIONS

AIRSPEED LIMITATIONS

M i L di G O ti S d

Maximum Landing Gear Extended Speed

VLE

• Do not exceed 182 KCAS/181 KIASwith landing gear extended.

Figure 14-28. Brake System Schematic (Serial Nos. BB-453 through BB-665)





Figure 14-29. Brake System Schematic (Serial Nos. BB-2 through BB-452)

1. Elect ric gear—When full up or downposition of the gear has been achieved,gear travel is limited by:

A. Limit switches only

B. Limit switches and the dynamic brakerelay

C. Physical stops in the drag brace

D. A slip clutch in the motor gearbox

2. Electric gear—Electric motor torque is

transmitted to the gear actuators:A. By a chain to all three gear actuators

B. By torque tubes to-a l l three gear actuators

C. By torque tubes to the nose gear actuator and by a chain to the maingear actuators

D. By torque tubes to the main gear

actuators and by a chain to the nosegear actuator

4. If the wing flaps are beyond the 40% (AP-PROACH) position, the warning hornwill sound if:

A. Both power levers are retarded belowa specified setting

B. Either power lever is retarded belowa specified setting

C. Any one gear is not down and lockedand power levers are below 80% N1position.

D. Any one gear is not down and locked,regardless of power lever setting

5. I f the wing flaps are up or a t 40% (AP-PROACH) position, the warning hornwill sound if:

A. E i t h e r o r b o t h po w e r l ev e r s a reretarded below a specified power setting

B. Both power levers are retarded below



QUESTIONS



gear actuator

3. Electric gear—In the event of electricalfailure, the landing gear is extended:

A. By using the emergency engage handleand the extension handle, located on thefloor aft or to the left of the pedestal

B. By releasing the uplocks and allowingthe gear to free-fall

C. By applying battery power to thelanding gear motor

D. By pull ing the emergency engagehandle and a l lowing the gear tofree-fall

pa specified setting and any one gear is not down and locked

C. Either or both power levers are retardedbelow a specified setting and any onegear is not down and locked

D. There is no other requirement

6. If the rudder pedals are deflected withthe airplane stationary:

A. The nosewheel steers, the rudder doesnot move

B. The spring-loaded link in the systemcompresses, the nosewheel does notsteer

C. The nosewheel does not steer and therudder does not move

7. When the PARKING BRAKE handle ispulled:

A. Two master cylinders are mechanicallyactuated, applying the brakes

B. Tw o m a s te r c y l in d e rs , a l r e ad yactuated, are mechanically held in

that positionC. The parking brake valve is actuated to

trap pressure from that point to brakeassemblies

D. T h e p a rk i n g b ra k e v a lv e i sm e c h a n i c a l l y a c t u a t e d t o b u i l dpressure for brake application

8. Hydraulic gear—The landing gear is heldretracted by:

A. Mechanical uplock mechanisms

B. Co n t in u o us l y a p p li e d h y dr a u l icpressure

C. Internal uplock mechanisms in allthree gear actuators

D. Spring tension

10. H y d r au l i c g ea r — Wi t h t he a i r p la n eairborne, placing the LDG GEAR CONTswitch handle UP:

A. Completes a circuit to the UP solenoidof the gear selector valve

B. Completes a circuit to the pump motor

relay, pulling in 28 VDC to start thepump motor

C. A and B

D. None of the above

11. Hydraulic gear—When the landing gear is fully retracted, the electrically drivenhydraulic pump:

A. Stops, and does not start again

B. Stops, but cycles as required

C. Operates continuously

D. Continues to operate for five minutes,then stops





9. Hydraul ic gear—The land ing gear i slocked down by:

A. Co n t in u o u s ly a p p l i ed h y d r a ul i cpressure

B. Internal downlock mechanisms in allthree gear actuators

C. An internal lock in the nose gearactuator and overcenter drag brace( n o s e g e a r ) o r b y m e c h a n i c a ldownlock mechanisms on the dragbraces (main gear)

D. Bungees

CHAPTER 15

FLIGHT CONTROLS

CONTENTS

Page

INTRODUCTION................................................................................................................. 15-1

GENERAL ............................................................................................................................ 15-1

FLIGHT CONTROL LOCKS............................................................................................... 15-2

ROLL..................................................................................................................................... 15-3

Operation ....................................................................................................................... 15-3

PITCH.................................................................................................................................... 15-3

Operation ....................................................................................................................... 15-3

YAW ...................................................................................................................................... 15-3

Operation 15-3





Operation ....................................................................................................................... 15 3

Rudder Boost ................................................................................................................. 15-3

Yaw Dampening............................................................................................................. 15-5

TRIM SYSTEMS.................................................................................................................. 15-5

Operation ....................................................................................................................... 15-5

Elevator Electric Trim.................................................................................................... 15-6

FLAPS ................................................................................................................................... 15-6

Operation ....................................................................................................................... 15-8

Split Flap Protection .................................................................................................... 15-10

LIMITATIONS.................................................................................................................... 15-12

Airspeed Limitations ................................................................................................... 15-12

Maneuver Limits.......................................................................................................... 15-12

Flight Load Factor Limits at 12,500 Pounds............................................................... 15-12Maximum Operating Pressure-Altitude Limits ........................................................... 15-12

QUESTIONS....................................................................................................................... 15-13





ILLUSTRATIONS

Figure Title Page

15-1 Flight Controls and Trim Tabs ............................................................................... 15-2

15-2 Flight Control Locks .............................................................................................. 15-315-3 Rudder Boost and Yaw Damp Switches ................................................................ 15-4

15-4 Rudder Boost Diagram........................................................................................... 15-5

15-5 Autopilot and Yaw Damp Switches ....................................................................... 15-5

15-6 Trim System Control.............................................................................................. 15-6

15-7 Elevator Electric Trim Controls ............................................................................. 15-7

15-8 Flap System Diagram............................................................................................. 15-8

15-9 Flap Control and Indication ................................................................................... 15-9

15-10 Stall Warning System........................................................................................... 15-11





2 0

2 0 2 0

10

5

5

1 0 10

5

5

L O

C

G S

CHAPTER 15FLIGHT CONTROLS





INTRODUCTION

The Super King Air is equipped with manually-actuated primary flight controls, oper-

ated through cables, bellcranks, and pushrods. The ailerons and rudder are conven-tional; the horizontal stabilizer and elevators are mounted at the extreme top of the verticalstabilizer, conforming to the T-tail configuration. A pneumatic rudder boost system as-sists in directional cont rol in the event of engine failure or a difference in engine bleedair pressure.

All surfaces are manually trimmed from the cockpit; however, optional elevator elec-tric trim is available. Two trailing-edge flaps on each wing are actuated by an electric

t d i i fl ibl d i h ft th h b A f t h i id lit



ELEVATORS

TRIM TABS

RUDDER

TRIM TAB

AILERON

TRIM TAB

FLAPS

FLAPS

GROUND ADJUSTABLE TAB

AILERON



FLIGHT CONTROLLOCKS

The flight control locks consist of a chain, twopins, and a U-shaped clamp (Figure 15-2).

The pin in the control column prevents con-trol wheel rotation and fore-and-aft move-ment of the control column, locking theailerons and elevators. On BB-82 and subse-quent, BL-1 and subsequent, and all B200 air-

l th t l l t b f ll f d

the rudder. The pedals must be centered beforethis pin can be installed.

The U-shaped clamp around the power leversserves as a warning not to start engines withthe control locks installed.

NOTE

The rudder control lock must be re-moved prior to towing the airplaneto prevent damage to the steering

li k

Figure 15-1. Flight Controls and Trim Tabs



bellcranks, and pushrod linkage to move theelevators. Elevator travel is approximately20° up and 14° down, and is limited by adjustable stops.

YAWOPERATION

Yaw control around the vertical axis is main-tained by the rudder, which extends along theentire aft edge of the vertical stabilizer. It isactuated, through cables and bellcranks, by ei-ther set of mechanically-connected rudder

pedals. Rudder travel is approximately 15°either side of neutral, and is limited by adjustable stops. Yaw damping and rudder boostare also activated through the rudder.

RUDDER BOOST

A rudder boost system is provided as an aidin maintaining directional control in the eventof engine failure or a large variation of power between the engines. Two pneumatic boost

Fig re 15 2 Flight Control Locks



ROLL

OPERATION

Roll control around the longitudinal axis ismaintained by conventional ailerons mountedon the trailing edge of each wing, outboard of the flaps. Rotation of either interconnectedcontrol wheel on the control column mech-anically positions the ailerons. Aileron travelis approximately 25° up and 17° down, lim-ited by adjustable stops.

between the engines. Two pneumatic boostservos are incorporated into the rudder cablesystem to provide force for rudder boosting,when required.

Operation

The rudder boost system is armed by placingthe RUDDER BOOST switch to the ON posi-tion, and both the left and right BLEED AIRVALVE switches in either the OPEN or ENVIROFF positions (Figure 15-3).

A differential pressure switch in the system( c o m m o n l y r e f e r r e d t o a s t h e D e l t a Pswitch) senses bleed-air pressure from each

i If b t ti l diff

Figure 15-2. Flight Control Locks



INSTR & ENVIR OFF

ENVROFF

OPENLEFT RIGHT

BLEED AIR VALVES

PARKINGBRAKEOFF

ENGINEANTI-ICE

ON

MAIN

OFF

ACTUATORSTANDBY

L EF T R IG HT

COLLINS

L A ND I NG T AX I I C E N A V R E C OG

OFF

LIGHTS

LEFT R I GHT

I N

P U L L



CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

PR

YAWENG

APENG

SR

I/20

DN

L

UP

R

YAWALT

YAW

HDG

COLLINS

HDG NAV APPR B/C CLIMB

ALT ALT SEL VS IAS DSC

DN

LIFT

LIFT

IDLE

GDFINE

PITCH

TRIM

UP

PROP

ITION

LOWIDLE

CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

PRESS

TEST ERASE

POWER

T ES T O FF O FF



18 PSIPRESSURE

REGULATOR

FILTER15 PSI

PRESSUREREGULATOR

N.C. N.C.


ENVIROFF

INSTR & ENVIR OFF

RUDDERBOOST

OFF

DUAL FED NO. 2

VDC

VDC

VDC

LEFT CONTROL(DUAL FED NO. 1)

RIGHT CONTROL(DUAL FED NO. 2)

PNEUMATICPRESSURE

18 PSI

TO RIGHT BLEED AIRWARNING SYSTEM

TO LEFT BLEED AIRWARNING SYTEM

CHECK VALVECHECK VALVE

60 PSID SWITCH

LEFTRUDDERSERVO

RIGHTRUDDER

SERVO

PNEUMATICSHUTOFF VALVE

PNEUMATICSHUTOFF VALVE

Figure 15-4. Rudder Boost Diagram



The system is tested during engine runup byretarding one engine to idle and advancingpower on the other engine until the rudder pedal on the same side as the high rpm enginemoves forward. Reverse the procedure to checkthe opposite side.

YAW DAMPING

OperationOn all airplanes, a yaw damping system isprovided. It can be activated with a switchlocated on the pedestal or autopilot panel(Figure 15-3 and Figure 15-5). On some in-

ll i i ill i ll i i h

TRIM SYSTEMS

OPERATION

Trim in all three axes is maintained by trim tabs

on the primary flight control surfaces. A tab

YAWALT

YAW

HDG

COLLINS



is located on the trailing edge of the rudder,each elevator, and the left aileron. Movingtrim wheels (Figure 15-6) mechanically trans-mits motion to screwjack actuators that posi-tion the tabs.

ELEVATOR ELECTRIC TRIMMost airplanes have an optional electric ele-vator trim system installed. An electric motor in the fuselage aft section actuates the eleva-tor trim tabs through a system of cables.

OperationThe ELEV TRIM switch must be placed in theON position to arm the system (Figure 15-7).

Electr ical power to the system is routedthrough the PITCH TRIM circuit breaker. DualPITCH TRIM thumb switches on the outboardside of either control wheel must be moved si-multaneously to achieve pitch trim. Either switch alone will not actuate the trim motor.As an option in some airplanes, trim inputs bythe pilot override those made by the copilot.The PITCH TRIM switches are spring loaded

to the center (off) position when released. Themanual elevator trim wheel can be used for trimming, even when the electrical trim sys-tem is switched on.

A bilevel, push-button, momentary-on trimdisconnect switch on each control wheel canbe used to disconnect the system. To initiatea trim disconnect, depress either of theseswitches to the second level. The green ELECTRIM OFF light on the advisory panel comeson when disconnect is selected. To reset thesystem for subsequent operation, cycle theELEV TAB CONTROL switch to OFF, thenback to ON.

FLAPS

Two flaps on each wing are driven by an elec-tric motor through a gearbox and four flexi-ble drive shafts connected to screwjacks ateach flap. The motor incorporates a dynamicbraking system through the use of two sets of field windings. Lowering the flaps results innose pitch-up, lowered stall speed, and re-





CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

P

DN

LIFT

LIFT

IDLE

GDFINE

PITCH

TRIM

UP

PROP

ITION

LOW

IDLE

POWER

ELEVATORTRIM WHEEL

AILERON

TRIM WHEEL

ELEVATOR TAB

CONTROL SWITCH



DEICE ON LDG / TAX I LIGHT

ELEC TRIM O FF

BATTERY CHARGE

PASS O

AIR COND

EXT

FLIGHT

5

ALT

ALERT

5

COPLT

5

TURN &SLIP

3

PILOTPILOT

ALTMAIR DATA

1

COPLT

ENCD ALTM

5

PITCH

TRIM

5

RUDDER

BOOST

5

OUTSIDE

AIRTEMP

PITCH TRIM CIRCUIT BREAKER ELECT TRIM OFF LIGHT



CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

P

R

ELEVATOR TRIM SWITCHN O S E D N

N O S E U P

P I

T C

H T R I

M

TRIM

DISCONNECT

SWITCH

duced airspeed. The flaps limit switches andflaps position transmitter are located under theright inboard flap.

OPERATION

Flap movement is initiated by positioning theFLAP handle (Figure 15-8).

Placing the FLAP handle fro m the UP to theAPPROACH (40%) position connects No. 3d u a l - f e d b u s p o w e r t h r o u g h t h e F LA PMOTOR circuit breaker to the flap motor (Figure 15-9). The flaps are driven to the40% (14° ± 1°) position, as indicated on the

flap position indicator. For BB-1439, 1444and subsequent, the flaps cannot be stoppedat any intermediate point during this travel.

Placing the handle to the DOWN position andleaving it there results in full 100%, (35° +1°, –2°) flap extension. For BB-1439, 1444 andsubsequent only the UP, APPROACH (or take-off), and DN positions are selectable. However,they are follow-up flaps which allows the flapsto extend or retract to achieve the selectedflap handle position. The flaps cannot bestopped in any intermediate position.

Prior to BB-1439, 1444 and subsequent, if anyposition between 40% and 100% is desired (for example, 60%), place the handle to DOWNuntil the desired position is attained, then re-turn it to the APPROACH position. The flapswill stop at 60%. In like manner, the flaps maybe raised to any position between DOWN andAPPROACH by placing the handle in the UPposition beyond the APPROACH detent until



POSITIONTRANSMITTER

RH

FLAP CONTROLC/B

FLAP MOTORC/B

DUAL FEDBUS NO. 3

TAKEOFFAND

APPROACH

UP

FLAPS20

60



FLAPDYNAMIC

BRAKERELAY

UP

APP

DOWN

STALL

FLAP

MOTOR

SPLITFLAPPROTECTION

LH

RH

FUSES ORCAM SWITCHES

U P

A P P R O A

C H

D O W N

F L A P

DOWN80

LIMIT SWITCHES

POSITIONINDICATOR



the desired position is reached, then returningit to the APPROACH detent. Moving the han-dle from DOWN to APPROACH will not retractthe flaps. When the handle is moved from theAPPROACH position to the UP position, theflaps retract completely and cannot be stoppedin between the APPROACH and UP positions.

A safety mechanism is provided to discon-nect power to the electric flap motor in theevent of a malfunction, which would causeany flap to be 3° to 6° out of phase with theother flaps.

The flap-motor power circuit is protected by a20-ampere flap-motor circuit breaker placarded FLAP MOTOR, located on the leftcircuit-breaker panel below the fuel controlpanel. A 5-ampere circuit breaker for thecontrol circuit (placarded FLAP CONTROL) isalso located on this panel.

Super King Air 200, BB-453 andSubsequent, BL-1 and Subsequent and

Super King Air B200/B200CWith the flaps in the UP or the APPROACHposition and either or both power levers re-

Protection is provided between each pair of flaps on that side of the airplane. There is nosplit flap protection between the left pair of flaps from the right.

STALL WARNINGOPERATION

The stall warning system senses angle of at-tack through a lift transducer actuated by avane mounted on the leading edge of the leftwing (Figure 15-10).

Angle of attack from the lift transducer andflap position signals are processed by the liftcomputer to sound the stall warning hornmounted on the copilot’s side of the cockpit.The horn sounds when the following condi-tions are present:

1. Airspeed is 5 to 13 knots above stall,flaps are fully retracted.

2. Airspeed is 5 to 12 knots above stall ,flaps are in the APPROACH (40%)position.





p ptarded below a certain power level, the warn-ing horn will sound intermittently and thelanding gear switch handle lights will illumi-nate. The horn can be silenced by pressingthe WARN HORN SILENCE button; the lightsin the switch handle cannot be cancelled. The

landing gear warning system will be rearmedif the power lever(s) are advanced sufficiently.

SPLIT FLAP PROTECTION

A split flap sensing system (Figure 15-8) pro-vides protection in the event any flap panel isout of phase with the other panel. Airplane se-

3. Airspeed is 8 to 14 knots above stall,flaps are fully extended.

The system can be tested prior to flight byplacing the STALL WARN TEST switch, lo-cated on the copilot’s left subpanel, in the

TEST position. This simulates a stall condi-tion and sounds the warning horn.

The heating elements protect thelift transducer vane and faceplatefrom ice. However, a buildup of ice

WARNING





LIFT TRANSDUCER VANE COPILOT'S LEFT SUBPANEL

LIMITATIONS

For complete limitations information, refer to the LIMITATIONS section of the Pi lot’sOperating Manual.

AIRSPEED LIMITATIONSManeuvering Speed

VA (12,500 pounds)

• Do not make full or abrupt control move-ments above 182 KCAS/181 KIAS.

Maximum Flap Extension/Extended Speed

VFE

APPROACH Position—40%• Do not extend flaps or operate with 40%

flaps above 200 KCAS/KIAS.

Full DOWN Position—100%

• Do not extend flaps or operate with100% flaps above 144 KCAS/146 KIAS(King Air 200) or 155 KCAS/157 KIAS(King Air B200).

Maximum Landing Gear Operating SpeedVLO

• Do not extend landing gear above 182

NOTE

Super King Air B200/B200C, Super King Air 200 Serial No. BB-199 andsubsequent, BL-1 and subsequent,and any earlier airplanes modifiedby Beechcraft Kit Number 101-5033-

1 in compliance with BeechcraftService Instruction Number 0894.

Maximum Operating Speed

VM0

MM0

• Do not exceed 270 KCAS/269 KIAS

(.48 Mach) in any operation.

NOTE

Super King Air 200 Serial No. BB-2 through BB-198, except airplanesmodified by Beechcraft Kit Number 101-5033-1 in compl iance wi thBeechc ra f t Se rv i ce Ins t ruc t i on

Number 0894.

MANEUVER LIMITS





Do not extend landing gear above 182KCAS/181 KIAS.

• Do not retract landing gear above 164KCAS/163 KIAS.

Maximum Landing Gear Extended Speed

VLE• Do not exceed 182 KCAS/181 KIAS

with landing gear extended.

Air Minimum Control Speed

VMCA

• The lowest airspeed at which the air-plane is directionally controllable whenone engine suddenly becomes inopera-

MANEUVER LIMITS

The Super King Air 200 and B200 are NormalCategory Aircraft. Acrobatic maneuvers, in-cluding spins, are prohibited.

FLIGHT LOAD FACTOR LIMITSAT 12,500 POUNDS

Flaps Up

• Do not exceed 3.17 positive Gs, or 1.27negative Gs.

Flaps Down

• Do not exceed 2.00 positive Gs, or 0

1. What is the maximum allowable altitudewith yaw damping inoperative?

A. 10,000 feet

B. 17,000 feet

C. 20,000 feet

D. 25,000 feet

2. What happens when the FLAP handle ismoved f rom the DOWN to the AP-PROACH position?

A. The flaps will bypass the APPROACHposition and retract fully.

B. The flaps will not retract.

C. Th e f la p s w il l r et r ac t t o t heAPPROACH position.

D. The flaps will retract completely, thenreturn to the APPROACH position.

3. O n a i rp l an e s w i th op t io n al el e va t orelectric trim, how is trim initiated?

A. Either the pilot or the copilot mov-i i th l t f hi PITCH

4. W hy sh ou ld th e r ud de r c on tr ol lo ckbe removed prior to towing the airplane?

A. So the airplane can be steered with therudder pedals

B. So the brakes can be applied

C. To prevent damage to the steeringlinkage

D. I t is not necessary to remove therudder control lock prior to towing.

5. H ow ca n th e ru dd er bo os t sy st em bechecked for proper operation during enginerunup?

A. Increasing power on an engine untilthe rudder pedal on the same sidemoves forward

B. Increasing power on an engine untilthe rudder pedal on the opposite sidemoves forward

C. Rudder boost operation cannot bechecked during engine runup

D. Reducing power on an engine and



QUESTIONS



ing ei ther element of his PITCHTRIM switch.

B. Both the pilot and the copilot movingboth elements of their PITCH TRIMs w i t c h e s i n t h e s a m e d i r e c t i o nsimultaneously.

C. Either the pilot or the copilot movesboth elements of his PITCH TRIMswitch simultaneous ly.

D. Both the pilot and copilot movingeither element of their PITCH TRIMs w i t c h e s i n t h e s a m e d i r e c t i o nsimultaneously.

g p gnot ing that nei ther rudder pedalmoves forward

6. How is the stall warning system normallytested prior to flight?

A. By manua l ly ac tua t ing the l i f ttransducer vane

B. B y m a n ua l l y a ct u a t in g t h e l if ttransducer vane and simultaneouslyplacing the STALL WARN TESTswitch in the TEST position

C. The system cannot be tested prior toflight

CHAPTER 16

AVIONICS

CONTENTS

Page

INTRODUCTION................................................................................................................. 16-1

COLLINS PROLINE II......................................................................................................... 16-1

Audio System................................................................................................................. 16-1

Communication Radios.................................................................................................. 16-3ADF Equipment........................................................................................................... 16-13

Transponder Equipment............................................................................................... 16-17

FLIGHT INSTRUMENTS.................................................................................................. 16-20

Pitot and Static System................................................................................................ 16-20

Outside Air Temperature Gage.................................................................................... 16-23

AUTOFLIGHT SYSTEM................................................................................................... 16-24





AUTOFLIGHT SYSTEM................................................................................................... 16 24

Yaw Damper ................................................................................................................ 16-24

STALL WARNING SYSTEM ............................................................................................ 16-25

COMMUNICATION SYSTEM.......................................................................................... 16-26

Static Discharging Description.................................................................................... 16-26

LIMITATIONS.................................................................................................................... 16-27

Airspeed Indicator ....................................................................................................... 16-27

Outside Air Temperature Gage.................................................................................... 16-27

ILLUSTRATIONS

Figure Title Page

16-1 B200 Audio Panel Controls ................................................................................... 16-2

16-2 VHF-22 COMM Radio Controls/Displays ............................................................ 16-5

16-3 VHF-22 COMM Radio Self Test Displays ............................................................ 16-6

16-4 VIR-32 NAV Receiver Controls/Displays ............................................................. 16-8

16-5 DME-42 Systems IND-42 Displays....................................................................... 16-9

16-6 Typical Pro Line II Dual DME Installation.......................................................... 16-11

16-7 IND-42 Self Test Displays................................................................................... 16-13

16-8 VIR-32 NAV Receiver Self-Test Displays........................................................... 16-13

16-9 ADF-60A ADF Receiver Controls/Displays ....................................................... 16-14

16-10 ADF-60A ADF Receiver Self-Test Displays....................................................... 16-17

16-11 TDR-94 Transponder Controls/Displays ............................................................. 16-17

16-12 TDR-94 Transponder Self-Test Displays............................................................. 16-20

16 13 16 21





16-13 B200 Antenna Locations...................................................................................... 16-21

16-14 Pitot and Static System Diagram ......................................................................... 16-22

16-15 Pitot Mast Location.............................................................................................. 16-23

16-16 Static Ports Location............................................................................................ 16-23

16-17 Pilot’s Static Air Source Valve Switch................................................................. 16-23

16-18 Typical OAT Gage and Probe .............................................................................. 16-24

16-19 YAW Damp Switch.............................................................................................. 16-25

TABLES

Table Title Page

16-1 CTL-22 COMM Control, Controls and Indications................................................ 16-4

16-2 CTL-32 NAV/DME Control, Controls and Indications .......................................... 16-7

16-3 IND-42A/C DME Indicator, Controls and Indications ......................................... 16-10

16-4 CTL-62 ADF Control, Controls and Indications .................................................. 16-15

16-5 CTL-92 ATC Control, Controls and Indications................................................... 16-18

16-6 Airspeed Indicator Markings................................................................................. 16-27





CHAPTER 16AVIONICS





INTRODUCTION

The Super King Air utilizes an avionics package which consists of, but is not limitedto, the navigation system, the weather radar system, the autoflight system, the stallwarning system, and the communication system.

COLLINS PROLINE II

AUDIO SYSTEM

1 6 - 2


ONL Y

S UP E R

KI N G

A I R

2 0 0 / B 2 0 0

P I L O T T R A

I NI N G MA

F l i

VOL VOL V O L

VOL VOL

AUTOCOMM

COMM

OFF

AUTOCOMM

OFFPILOT AUDIO OFF

NAV

AUDIOSPKR

SIDE-TONE

INTPHSENS

OFF

AUDIOSPKR

OFFNORM

HI

LO

DME

RANGE

VOICE PAGING INTPHAUDIOEMER

OFF

HOTINTPH

MKR BCN DME

1 12 2

1 2

2 2 ADF1 1

COMM

COPILOT AUDIO OFF

NAV DME

1 12 2 2 2 ADF1 1

BOTH

RANGE

VOICE BOTH

MKR BCN1 & 2

VOL

COMM 1

COMM 2

CABIN

VO L

COMM 1

COMM 2

CABIN

PUSHON/OFF

GNDCOMMPWR

ANNPUSH BRT

SIDE-TONE

INTPHSENS

ENCDALTM1

2

AVIONICS BY BEECHCRAFT

MKR BCN

D I M

COPILOTSMICROPHONESELECTORSWITCH

PILOTSMICROPHONE

SELECTORSWITCH

COPILOT AUDIO OUTPUTSOURCE SELECT SWITCHES

PILOT AUDIO OUTPUTSOURCE SELECT SWITCHES

PASSENGERADDRESSVOLUME

CONTROL

COPILOT NAV/ADF

VOICE/MORSEFILTER SWITCH

PILOT NAV/ADFVOICE/MORSEFILTER SWITCH

COPILOTSIDETONE

VOLUME POTPILOT

SIDETONEVOLUME POT

COPILOTINTERPHONE

THRESHOLD POT

PILOTINTERPHONE

THRESHOLD POT

COPILOTSPEAKERSWITCH

PILOTSPEAKER

SWITCH

*COPILOT AUTOCOMM SWITCH

*PILOT AUTOCOMM SWITCH

*AUTOMATICALLY TURNSON AUDIO OUTPUT FROMSELECTED TRANSMITTER

*AUTOMATICALLY TURNSON AUDIO OUTPUT FROMSELECTED TRANSMITTER

INTERPHONEVOLUME

CONTROL

COPILOT

MASTER

VOLUME

CONTROL

PILOTMASTERVOLUME

CONTROL

MARKER BEACONVOLUME CONTROL

MARKER BEACONSENSE SWITCH

HOT INTERPHONEON/OFF SWITCH

DME VOLUMECONTROLS

*AUDIO AMPSBYPASS SWITCH

*WHEN IN EMER POSITION THE FOLLOWINGAUDIO ONLY BYPASSES THE AMP AND ALLARE MIXED TOGETHER:COMM 1, COMM,2 SIDETONE 1, SIDETONE 2,AND AURAL WARNING INPUTS

GROUND COMMPOWER SWITCH



AN U A L

i gh t S af e t y


Figure 16-1. B200 Audio Panel Controls



OperationThe two audio control amplifiers operate in-dependently for the pilot and copilot systems.If the pilot moves a source select switch up,the pilot’s headset will receive audio fromthat source. If the “Audio Speaker” switch is

selected up, then the audio will also be broad-cast over the pilot’s speaker. If this audio con-tains both a Morse code (Range) and a voicemessage, these can be listened to simultane-ously or individually with the switch labeledVOICE–BOTH–RANGE. All of the above se-lections are identical on the copilot’s side.

A microphone selector switch allows trans-

mission through either Communication RadioNo. 1 (COMM 1), Radio No. 2 (COMM 2), or to the “Cabin” speakers. For the COMM 1 andCOMM 2 selection an AUTO COMM switchwill automatically send the selected COMMradio’s received audio to the headset (andspeaker, if selected). This eliminates havingto select the COMM radio on the source selectpanel unless audio from the opposite radio is

desired (e.g., microphone selector switch onCOMM 1 for enroute communication but ATISon COMM 2).

E h COMM di h i di id l l

through the use of the microphone selector switch mentioned above. A knob labeled PAG-ING allows volume control of cabin audio.

The aural warning tone generator generatesaural warning tones to the cockpit headphonesand speaker for stall warning, landing gear

warning, autopilot disconnect warning, and al-titude alert warning (other optional equip-ment may be included). The aural warningtone generator operates on command fromfault detection equipment and does not havea volume control.

An AUDIO EMER switch is provided shouldthere be a failure of both amplifier systems

(Figure 16-1). When selected to EMER, theamplifier selector panel is no longer func-tional. COMM 1, COMM 2, SIDETONE 1,SIDETONE 2, and aural warning tones are allsimultaneously output to the headsets only,wi thout individual se lect ion capabi l i ty(Sidetone = transmitted audio sent to the head-set to moni tor qual i ty of t ransmission) .Therefore each radio volume control should

be turned down to listen to the appropriateunit for that point in time. A Morse code iden-tifier check will not be possible on the re-maining NAV radios during the EMER mode.



Each COMM radio has an individual volumecontrol knob. However, a knob is provided oneach microphone selector switch for master volume control (this does not affect incomingwarning tones).

A GND COMM PWR button is connected tothe hot battery bus of the aircraft and allowsactivation of headsets, speakers and handheldmicrophones and COMM 1 prior to turning onthe main aircraft battery switch. This allowsfrequencies such as clearance delivery andATIS to be retrieved without excessive batteryuse. Once communication is complete, turn off

g g

COMMUNICATION RADIOS

General

Two communications transceivers (VHF-22A) are installed and are each controlledby a CTL-22 control. This control is shownin Figure 16-2 and detailed operation/de-scription are shown in Table 16-1.

Operation



CONTROL OR DISPLAY FUNCTION/DESCRIPTION

ACTIVE FREQUENCY DISPLAY The active frequency (frequency to which the VHF-22A is tuned) and dashes

(----) on dIAG during self test.

PRESET FREQUENCY DISPLAY The preset (inactive) frequency and diagnostic messages are displayed in thelower window.

COMPARE ANNUNCIATOR ACT momentarily illuminates when frequencies are being changed. If ACT

continues to flash, the actual radio frequency is not identical to the frequency

shown in the active window.

ANNUNCIATORS The COMM control contains MEM (memory), and TX (transmit) annunciators.

The MEM annunciator illuminates whenever a frequency is displayed in the

lower window. The TX annunciator illuminates whenever the VHF-22A is

transmitting.

POWER AND MODE SWITCH The OFF–ON positions switch system power to turn the system on or off. The SQ OFF position disables the receiver squelch circuits, so you should hear

noise. Use this position to set the volume control or, if necessary, to try to

receive a very weak signal which cannot break squelch.

LIGHT SENSOR The built-in light sensor automatically controls the display brightness. The ANN

PUSH BRT control knob/push button can be used to override the automatic dim

controls and force the display to go to full bright.

XFR/MEM SWITCH This switch is a three-position, spring-loaded toggle switch. When moved to the

XFR position, the preset frequency is transferred up to the active display and the

VHF-22A retunes. The previously active frequency becomes the new presetfrequency and is displayed in the lower window. When this switch is moved to

the MEM position, one of the six stacked memory frequencies is loaded into the

preset display. Successive pushes cycle the six memory frequencies through the

display ( 2 3 4 5 6 1 2 3 )

Table 16-1. CTL-22 COMM CONTROL, CONTROLS AND INDICATIONS



display (...2, 3, 4, 5, 6, 1, 2, 3...).

FREQUENCY SELECT KNOBS Two concentric knobs control the preset or active frequency display. The larger

knob changes the three digits to the left of the decimal point in 1-MHz steps. The

smaller knob changes the two digits to the right of the decimal point in 50-kHz

steps (or in 25-kHz steps for the first two steps after the direction of rotation has

been reversed). Numbers roll over at the upper and lower frequency limits.

ACT BUTTON Push the ACT button for approximately two seconds to enable the frequency

select knobs to directly retune the VHF-22A. The bottom window will display

dashes and the upper window will continue to display the active frequency. Push

the ACT button a second time to return the control to the normal two-display

tune/preset mode of operation. The active tuning feature is not affected by power

removal. If active tuning is selected (one push of the ACT button) and power is

removed from the control, active tuning will still be enabled the next time power

is reapplied to the control.

STO BUTTON The STO button allows up to six preset frequencies to be selected and entered



Moving the mode switch (Figure 16-2) to SQOFF and then adjusting the background noiselevel can help set the radio volume. After acomfortable level has been established, re-turn the mode switch to ON to return thesquelch back to normal.

Whenever a new frequency is selected in theactive window, an ACT annunciator (com-pare annunciator) illuminates to indicate thetransceiver is changing frequencies and thenextinguishes after the process is complete. If it continues to flash, then the selected fre-quency on the CTL-22 is not the frequencybeing used by the transceiver. A recycling of the frequency should be accomplished (veri-

fying the ACT annunciator extinguishes) or theuse of a different radio.

For frequency selection and storing refer toTable 16-1.

Stuck MIC Protection

Each time the Push-to-Talk (PTT) button is

pushed the microprocessor in the transceiver starts a two-minute timer. If the transmitter isstill on at the end of two minutes the micro-processor turns it off. This protects the ATCchanne l f rom long- t e rm in te r f e r ence .Transmission is indicated by a TX annuncia-tor on the CTL-22 control and continuous il-lumination without pressing the PTT buttonwould indicate this malfunction (Figure 16-2

and Table 16-1). This annunciator should ex-tinguish at the two-minute time limit.

COMPAREANNUNCIATOR

PRESET COMM

FREQUENCYDISPLAY

ACTIVE COMMFREQUENCYDISPLAY

TRANSFER/ MEMORYSWITCH



MEMORY

ANNUNCIATOR

POWER ANDMODE SELECT

SWITCH

COMM VOLUMECONTROL

LIGHT TEST MEMORY

ACTIVE TUNEBUTTON(ACTIVE TUNING/ PRESET TUNING)

FREQUENCYSELECTKNOBS (2)

TRANSMITANNUNCIATOR

Overtemperature ProtectionThe temperature of the transmitter is contin-uously monitored by a microprocessor. If thetemperature exceeds +160°C (+320°F) thetransmitter is shut down. This immediatelyeliminates sidetone. A release of the PTT will

sound two beeps, and as long as the tempera-ture remains above the limit, the transmitter will not operate. If you must transmit, youcan override the protection by rapidly keyingthe PTT button twice and holding it on thesecond push.

Self Test

An extensive self-test diagnostic routine can beinitiated in the transceiver by pushing the TESTbutton on the radio. During the self test, theupper and lower displays modulate from mini-mum to maximum lighting intensity to indicatethe self test is in progress. Several audio toneswill be heard from the audio system during thistest. At the completion of the self test the radiousually displays four dashes (- - - -) in the upper

display and 00 in the lower display (Figure 16-3). If any out-of-limit conditions are found, theupper display will read dIAG and the lower dis-play will contain a two-digit diagnostic code(Figure 16-3).

Navigation/DME EquipmentTwo navigation receivers (VIR-32) are installedand are each controlled by a CTL-32 control.This control is shown in Figure 16-4 and detailedfunction/descriptions are listed in Table 16-2.

Two DME transceivers (DME-42) are also in-stalled and they show information on the re-spective indicator (IND-42). These are shownin Figure 16-5 with a detailed function/de-scription in Table 16-3.

OperationThe navigation radios provide VOR, LOC,

GS and marker beacon capabilities. On-sideor cross-side course display information canthen be selected on the pilot and copilot dis-plays via push buttons or an EFIS system.

Selection of frequencies and storing are iden-tical to that described in the communicationssection above.

The DME equipment will indicate the slant-range distance when the on-side navigationradio contains a DME assoc iated frequency.The DME identifier is sent once every 30

d d h i d th i di t





(Figure 16 3). seconds and when received, the indicator

DIAGNOSTIC

ANNUNCIATOR

DIAGNOSTICCODE



Table 16-2. CTL-32 NAV/DME CONTROL, CONTROLS AND INDICATIONS


ACTIVE FREQUENCY DISPLAY Active frequency is the frequency to which the DME-42 channel 1 is tuned.In DME or NAV self test, diagnostic messages are displayed in the upper

window.

PRESET FREQUENCY DISPLAY The preset (inactive) frequency is the frequency to which the DME-42channel 2 is tuned. In DME or NAV self test, diagnostic messages are

displayed in the lower window.

COMPARE ANNUNCIATOR ACT momentarily illuminates when frequencies are being changed. If ACT

continues to flash, the actual tuned frequency is not identical to thefrequency shown in the active display.

ANNUNCIATORS

MEM

HLD

The NAV control contains MEM (memory) and HLD (hold) annunciators.

The MEM annunciator illuminates when a frequency is displayed in thelower window.

The HLD annunciator indicates the DME is in DME hold. In this mode it isnormally tuned to the frequency displayed in the active window at the time

of selection. After selecting hold, the upper window displays the NAV

frequency and the lower window displays the DME hold frequency. Tuningof the active frequency can take place during this time. When completed,

the unit will always revert back to display of the DME hold frequency in thelower window.

VOLUME CONTROL The volume control is concentric with the power and mode switch. Itcontrols only the NAV receiver volume.

POWER AND MODE SWITCH The NAV control power and mode switch contains three detented

positions. The positions are: OFF–ON–HLD.



The OFF–ON positions switch system power.

The HLD position allows the NAV frequency to be changed but holds the

DME to the current active NAV frequency.

LIGHT SENSOR The built-in light sensor automatically controls the display brightness. TheANN PUSH BRT control knob/push button can be used to override theautomatic dim controls and force the display to go to full bright.

XFR/MEM SWITCH This switch is a three-position, spring-loaded toggle switch. When moved to

the XFR position, the preset frequency is transferred up to the activedisplay and the NAV/DME retunes. The previously active frequency

becomes the new preset frequency and is displayed in the lower window.

When this switch is moved to the MEM position, one of the four stackedmemory frequencies is loaded into the preset display. Successive pushes



Table 16-2. CTL-32 NAV/DME CONTROL, CONTROLS AND INDICATIONS (Cont)

MARKER BEACONVOLUME CONTROL

MARKER BEACONSENSITIVITY SWITCH


ACT BUTTON Push the ACT button for approximately two seconds to enable the frequencyselect knobs to directly retune the VIR-32 and DME. The bottom window will

display dashes and the upper window will continue to display the active

frequency. Push the ACT button a second time to return the control to the normaltwo-display tune/preset mode of operation. The active tuning feature is not

affected by power removal. If active tuning is selected (one push of the ACTbutton) and power is removed from the control, active tuning will still be enabled

the next time power is reapplied to the control.

STO BUTTON The STO button allows up to four preset frequencies to be selected and entered

into the control’s nonvolatile memory. To store a frequency, simply toggle theMEM switch until the upper window displays the desired channel number (CH 1

through CH 4), rotate the frequency select knobs until the lower window displaysthe frequency to be stored, and push the STO button twice within five seconds.

After approximately five seconds, the control will return to the normal two-displaytune/preset mode of operation.

TEST BUTTON Push the TEST button to initiate the radio self test diagnostic routine. (In the case

of the VIR-32 NAV receiver, self-test is active only while the TEST button ispushed or about 15 seconds maximum. In the case of the DME-442 transceiver,

the self test routine requires about 10 seconds for completion.)



MEMORYANNUNCIATOR DME HOLD

ANNUNCIATOR

COMPAREANNUNCIATOR


SWITCH

FREQUENCYSELECT

ACTIVE VOR/LOCFREQUENCYDISPLAY

PRESET VOR/LOCFREQUENCY

DISPLAY

TRANSFER /MEMORYSWITCH



VOL VOL

DME1 2

ENCD

ALTM1

COLLINS

SEL PWR

18.8.8.8 8 8

1 2 32 3 WPTPT NMM8

8 8 8

HLDLD KTT MININ IDD

DISPLAYGROUND SPEEDTIME TO STATIONSTATION IDENT

No. 2 DMEAUDIO VOLUMENo. 1 DME

AUDIO VOLUME

DISTANCE DISPLAY

CHANNELANNUNCIATORS

DISTANCELABEL

IND-42A (PILOT'S DISPLAY)

DISPLAY

ANNUNCIATORS

X X X X



COLLINS

CH SEL PWR

18.8.8.8 8 8

1 2 32 3 WPTPT NMM

8 8 8 8

HLDLD KTT MININ IDD

LIGHT SENSOR

POWERSWITCH

CHANNELSWITCH(1, 2, 3)

X X X X

ANNUNCIATOR DESCRIPTION

CONTROL/INDICATOR FUNCTION/DESCRIPTION

NUMERIC DISPLAY The numeric display presents the NM (distance) and diagnostic code.

ALPHANUMERIC DISPLAY The alphanumeric display presents the KT (velocity), MIN (time-to-station), ID

(2-, 3-, or 4-letter station identifier), and diagnostic identifier.POWER (PWR) SWITCH The latching push-on/push-off PWR switch controls the power applied to the

IND-42.

MODE SELECTOR (SEL) SWITCH

ALPHANUMERICThe non latching pushbutton SEL switch selects the information to be displayedin the display. (When power is initially applied, NM (distance) is shown in the

numeric display and ID (DME station identifier) is shown in the alphanumericdisplay.) Pressing the SEL switch will sequentially select KT (velocity), MIN

(time-to-station), and ID (2-, 3-, or 4-letter station identifier).

KT, MIN, and ID are shown in the alphanumeric display and NM (distance) is

continuously shown in the numeric display, provided the DME is locked on asignal.

CHANNEL (CH) SWITCH

(IND-42A ONLY)The momentary pushbutton CH switch sequentially selects the information from

the next DME channel and lights the appropriate channel annunciator 1, 2, or 3.The copilot’s IND-42C will always power up on channel 2

ANNUNCIATORS The annunciators provide an indication of which DME channel is selected,

system operational information, and units of measure. The following listdescribes the annunciators.

1 2 3

Sequentially controlled by the channel (CH) button to indicate which DME

channel is providing the information being displayed in the numeric andalphanumeric displays.



Table 16-3. IND-42A/C DME INDICATOR, CONTROLS AND INDICATIONS



NMAutomatically illuminates after power on when valid DME data is available.Indicates that the numbers displayed in the numeric display are slant range

DME distance in nautical miles.

HLDIndicates that DME hold has been selected on the CTL-32 NAV Control. When

in HLD, KTS and MIN will revert to ID after approximately 5 seconds.

KTIndicates that the value displayed in the alphanumeric display is the computedrate of change of DME distance.

MINIndicates that the value displayed in the alphanumeric display is the computed

time-to-station in minutes.

ID

Automatically illuminates after power on. The DME ident is transmitted onceevery 30 seconds and it is possible that 2 minutes could elapse before the

station ident is displayed in the alphanumeric display. The station identifier is

using the DME from a VOR). To operate in theHOLD mode, first select the desired DME sta-tion and then move the power and mode switchto HOLD. The frequency being held will appear in the preset window. The CTL-32 and DME in-dicator will now show HLD in the annunciator section as a reminder of the HOLD selection.

(KT and MIN can be selected on the DME in-dicator during HOLD operations; however, theindicator will default to ID after five seconds.)Although, the preset frequency is showing the“held” station, any previously selected fre-quency still remains in memory and a movementof the XFR/MEM switch to XFR will move thatpreset frequency to the active window. If fre-quency selection is done by a movement of the

frequency select knobs, only the active win-dow will change. To return to preset frequencyselection the power and mode switch must beturned to ON. Frequencies can still be retrievedfrom the memory during HOLD operations asdiscussed in Table 16-2.

In a dual DME installation the copilot’s indi-cator will usually allow selection of different

DME channels. By repeatedly pushing the CHbutton, these channels can be cycled. The cur-rent selected option will be indicated on thedisplay as 1, 2, or 3 (Figure 16-5). A typicalinstallation of channel usage is shown in Figure

test sequence. During the self test, the upper and lower displays modulate from minimumto maximum lighting intensity to indicate theself test is in progress. The DME will be placedin self test at the same time.

VOR Self TestSelect a VOR frequency on the CTL-32 NAVcontrol. (108.20 MHz will do. A specific fre-quency is not required for test.) A signal on thefrequency will not interfere with the self test.

• Select VOR-1 or -2 (as required) as theactive course sensor on the EHSI.

• Rotate the Course Select knob to ap-proximately 0°.

• Push and hold the TEST button on theCTL-32.

• The active course sensor VOR1 or 2annunciator on the EHSI will turn red.

• After approximately two seconds, theVOR1 or 2 annunciator will turn green,

the EHSI lateral deviation bar will ap-proximately center, and a TO indicationwill appear. The RMI pointers con-nected to the VIR-32 will indicate ap-proximately 0° magnetic bearing.





g g16-6. If the pilot does not have channel se-lection then channel 1 is the default indication.

Self TestLike the communication radios, the NAV ra-dios provide the ability for an extensive self-

• Release the TEST button. (The VIR-32will return to normal operation after ap-proximately 15 seconds, even if theTEST button is held.)

CH1 - NAV 1 ACTIVE2 - NOT USED

CHDME 2

ILS (Localizer and Glideslope)Self-TestSelect a localizer frequency on the CTL-32NAV control. (108.10 MHz will do. A specificfrequency is not required for test.)

• Select LOC1 or 2 (as required) as the ac-tive course sensor on the EHSI.

• Push and hold the TEST button on theCTL-32.

• The active course sensor LOC1 or 2annunciator on the EHSI will turn redand the red GS flag will come into view.

• After approximately three seconds,

the LOC1 or 2 annunciator will turngreen and the GS flag will go out of view, the EHSI lateral deviation bar will deflect right approximately two-thirds of full scale, and the glides-l o p e p o i n t e r w i l l d e f l e c t d o w napproximately two-thirds of full scale.

• Release the TEST button. (The VIR-32will return to normal operation after ap-

proximately 15 seconds, even if theTEST button is held.)

Marker Beacon Self-Test

• Read the distance to the station on theIND-42A/C and the left side of the EHSIdisplay. Verify the station ID next to thedistance display.

NOTE

The DME can require at least 30 sec-

onds, and as much as two minutes, toproperly decode the station ident.

• On the CTL-32, push TEST. On the IND-42 the following happens:

• Initially the IND-42A/C display mod-ulates in intensity between maximumand minimum brightness.

• LH display on IND-42A/C shows atest distance of 100.0(nm). After about10 seconds, the RH display shows anAOK (Figure 16-7).

• Listen to DME audio and note thata u d i o i s a Mo r se c o d e A O K(• – – – – – • –).

• Push SEL to annunciate KT and read100 (knots) in RH window.

NOTE

If the 10-second self test expires be-





The marker beacon assembly is tested auto-matically when the TEST button on the CTL-32 is pushed and either a VOR or localizer frequency is selected. For No. 1 NAV receiver

proper operation of the marker beacon as-sembly is indicated by all three-marker an-nunciators on the EADI cycling through inorder. For No. 2 NAV receiver the indicationwill be the three marker annunciators flick-ering at 30Hz. In addition, a tone will also bepresent in the marker beacon audio output.

If the 10 second self test expires before reaching this point, select TSTagain and continue with the test.

If there are any detected faults in the system onthe IND-42A/C a diagnostic code will appear in place of the AOK display (Figure 16-7). TheEFIS display will only show dashes for a fault.

The diagnostic routines are intended as an ex-tension of the self-test capability. The opera-tor should first observe the deviation indicatorsand associated flags for the proper self-test re-

with the code 00, indicating normal operation,no trouble found (Figure 16-8). If an out-of-limitcondition is detected during self test, that two-di i d ill l b di l d h CTL



IND-42A SELF-TEST DISPLAYNO FAULTS FOUND

IND-42C SELF-TEST DISPLAY 03(TRANSMITTER) FAULT FOUND

2

Figure 16-7. IND-42 Self Test Displays

TEST DISPLAY NO FAULT PRESENT

TEST DISPLAY ABNORMAL OPERATION PRESENT

FLAGANNUNCIATOR

DIAGNOSTICCODE

DIAGNOSTIC

ANNUNCIATOR

DIAGNOSTICCODE



digit code will also be displayed on the CTL-32 along with the word dIAG (diagnostic) or FLAG. FLAG will be displayed along with atwo-digit code when something is abnormal but

a failure has not occurred (i.e., low signal level,etc.) (Figure 16-8). dIAG is displayed alongwith a two-digit code to indicate a failure hasbeen detected in the VIR-32 (Figure 16-8).

Completion of self test is indicated when theNAV control displays either the normal activeand preset frequencies in the upper and lower windows, respectively, or a two-digit code.

is controlled through the CTL-62 control. Thiscontrol is shown in Figure 16-9 and detailedfunction/descriptions are listed in Table 16-4.

The ADF receives transmissions from a se-

TEST DISPLAY FAILURE PRESENT

Figure 16-8. VIR-32 NAV ReceiverSelf-Test Displays

ANT mode, the ADF receiver functions as an

aural receiver, providing only an aural outputof the received signal. In ADF mode, it func-tions as an automatic direction finder receiver in which bearing-to-the-station is presented onan associated bearing indicator and an aural

NOTE

If the RMI pointer remains parked,the system may not be receiving areliable signal. In this case, try twoor three other stations if possible.



MEMORYANNUNCIATOR

COMPAREANNUNCIATOR

NAV VOLUME

CONTROL


SWITCH

MEMORYSTORE

BUTTON

TESTBUTTON

ACTIVE TUNEBUTTON(ACTIVE TUNING/ PRESET TUNING)

FREQUENCYSELECTKNOBS (2)

ACTIVE ADFFREQUENCYDISPLAY

PRESET ADFFREQUENCY

DISPLAY

LIGHTSENSOR

TRANSFER/ MEMORYSWITCH

CTL-62 ADF CONTROL

Figure 16-9. ADF-60A ADF Receiver Controls/Displays



output of the received signal is provided. ATONE mode provides a 1000-Hz aural outputtone when a signal is being received to allowidentification of keyed CW signals.

ADF Self Test• Apply power to the ADF and RMI, and

EFIS systems.

• Select appropriate bearing pointer on theRMI and EHSI for the ADF to be tested.

• Set the control mode to ADF. This ap-

• Push and hold the self-test switch. Notethe RMI and EHSI bearing pointer ro-tates 90° counterclockwise. Releaseself-test switch.

NOTE

If the signal is weak or of poor qual-ity, the bearing pointer can rotaterather slowly. Degraded receiver sensitivity might give the same re-sponse.



Table 16-4. CTL-62 ADF CONTROL, CONTROLS AND INDICATIONS


ACTIVE FREQUENCY DISPLAY The active frequency; the frequency to which the ADF-60A is tuned. In self-test

mode and if an out-of-tolerance condition is detected, the word “dIAG” isdisplayed in the upper window while the diagnostic code is displayed in the

lower window.

PRESET FREQUENCYDISPLAY

The preset frequency is displayed in the lower window. In self-test mode and ifan out-of-tolerance condition is detected, the diagnostic code is displayed in

the lower window.

COMPARE ANNUNCIATOR ACT momentarily illuminates when frequencies are being changed. If the ACT

annunciator continues flashing, the receiver is not tuned to the displayed

active frequency.

ANNUNCIATORSThe ADF control contains a MEM (memory) annunciator. The MEM

annunciator illuminates whenever a frequency is displayed in the lower window

VOLUME CONTROL The volume control, is concentric with the power and mode switch and

controls ADF audio volume.

LIGHT SENSOR The built-in light sensor automatically controls the display brightness. The annpush brt control knob/push button can be used to override the automatic dim


POWER AND MODE SWITCH

OFF

ANT

The power and mode switch contains four detented positions.

The OFF position interrupts system power ( Turns the ADF off). Selecting ANT,ADF, or TONE applies power to the ADF system and establishes the systemmode of operation.

In ANT mode, the ADF receiver functions as an aural receiver, providing only

an aural output of the received signal



ADF

TONE

g

In ADF mode, it functions as an automatic direction finder receiver in which

bearing-to-the-station is presented on an associated bearing indicator and anaural output of the received signal is provided.

TONE mode provides a 1000-Hz aural output tone when a keyed CW signal is

being received.

XFR/MEM SWITCH This switch is a 3-position, spring-loaded toggle switch. When moved to the

XFR position, the preset frequency is transferred up to the active display and

the ADF-60 retunes. The previously active frequency becomes the new presetfrequency and is displayed in the lower window. When this switch is moved to

the MEM position, one of the four stacked memory frequencies is loaded intothe preset display. Successive pushes to the MEM position cycles the four

memory frequencies through the display ( 2 3 4 1 2 3 ) The frequency



Table 16-4. CTL-62 ADF CONTROL, CONTROLS AND INDICATIONS (Cont)


TUNING Normally, tuning is accomplished by entering a frequency into the presetwindow and then either storing that frequency in memory (STO) or entering it

into the active window (XFR) to tune the receiver. An alternate method is to

press the ACT button for at least 2 seconds (this gives direct tuning access tothe upper window) and insert the desired frequency directly into the active

window.

FREQUENCY SELECT KNOBS Two concentric knobs control the preset or active frequency displays. The

larger knob changes the 1000’s and 100’s kHz digits. The smaller knob

changes the 10’s, units, and tenths kHz digits. Each detent of the larger knobchanges the frequency in 100-kHz steps. Each detent of the smaller knob

changes the frequency in 1-kHz steps with the exception that the first twodetent positions following a change in rotational direction will cause a 0.5-kHz

change. Rapid rotation of the smaller knob will cause frequency changesgreater than 1 kHz as a function of the rate of rotation. Frequencies roll over

at the upper and lower limits. The two frequency select switches are

independent of each other such that the upper and lower limit rollover of the10-kHz digit will not cause the 100-kHz digit to change.

ACT BUTTONPush the ACT button for approximately 2 seconds to directly change the

active display window with the frequency select knob. The bottom windowwill display dashes. Push the ACT button a second time for about 2 seconds

to return the control to the normal 2-display tune/preset mode of operation.

The active tuning feature is not affected by power removal. If active tuning isselected (one push of the ACT button) and power is removed from thecontrol, active tuning will still be enabled the next time power is reapplied to

the control.

STO BUTTON The STO button allows up to four preset frequencies to be selected and



p p q

entered into the control’s nonvolatile memory. To store a frequency, simply

toggle the MEM switch until the upper window displays the desired channelnumber (CH 1 through CH 4), rotate the frequency select knobs until the

lower window displays the frequency to be stored, and press the STO button

twice within 5 seconds. After approximately 5 seconds, the control will returnto the normal 2-display tune/preset mode of operation.

TEST BUTTON Push the TEST button to initiate the radio self-test routine. Self-test is active

only while the TEST button is pushed. The display modulate in intensity whilethe TEST button is pushed.

When the CTL-62 is used with the CAD-62to control the ADF-60A system, certain di-agnostics codes can be displayed on theCTL-62 in the self-test mode. The diagnos-tic display appears when the self-test buttonis pressed as above for the ADF self test. If the diagnostics detect no faults, the CTL-62

displays four dashes (- - - -) and code 00(Figure 16-10). If the diagnostics detect afault dIAG will be displayed along the faultcode on the CTL-62 display (Figure 16-10).

TRANSPONDER EQUIPMENT

GeneralTwo TDR-94 transponders are installed in theB200 aircraft with only one operating at anyone time. The TDR-94 is a mode-A, mode-Cand mode-S transponder and is an integralpart of the Air Traffic Control Radar BeaconSystem. These transponders are controlled bya CTL-92 and is shown in Figure 16-11 withdetailed function/descriptions in Table 16-5.



TEST DISPLAY NO FAULT PRESENT TEST DISPLAY FAILURE PRESENT

DIAGNOSTICCODE

DIAGNOSTICANNUNCIATOR

Figure 16-10. ADF-60A ADF Receiver Self-Test Displays



COMPAREANNUNCIATOR

CODESELECT

ACTIVE CODEDISPLAY

NO. 1 OR NO. 2SELECT SWITCH

IDENTDISPLAY

(DISPLAYED WHENIDENT BUTTON USED)

TRANSPONDER REPLYANNUNCIATOR

ENCODINGALTIMETER

SOURCE SELECTSWITCH



FUNCTION/DESCRIPTION

ANNUNCIATOR

CONTROL ORDISPLAY

UPPER DISPLAY WINDOW

LOWER DISPLAY WINDOW

The ATC code (code with which the active transponder replies) and diagnostic

messages are displayed in the upper display window. During normal operation, the

CTL-92 has only a single display (the transponder code) shown in the upper window.

The lower display window is normally blank. It is active only during self test. If a

fail/warn condition is detected, dIAG will be displayed. Press the TEST button to

view the diagnostic code.

COMPARE ANNUNCIATORACT momentarily illuminates when codes are being changed. If ACT flashes, the

actual reply code is not identical to the code shown in the active code display.

The ATC control annunciator contains a TX (transmit) annunciator. The TX

annunciator illuminates when the transponder replies to an interrogation.

POWER AND MODE

SWITCH

The ATC control power and mode switch contains four detented positions. Available

positions are: OFF–STBY–ON–ALT.

Power is removed in the OFF position and is applied when any of the other modes

is selected.

In the STBY mode, power is applied to the transponder but it is prevented from

transmitting replies. STBY should be used only during taxi or when requested by ATC.

The ALT position is the normal operating position and allows the transponder to

reply to the interrogation pulses, as well as transmitting uncorrected barometricaltitude when the transponder is interrogated in mode C.

The ON position deletes the altitude code and is normally used when requested by ATC.

1/2 SWITCH The 1/2 switch selects which of two transponders is active.

The built-in light sensor automatically controls the display brightness. The ANN

Table 16-5. CLT-92 ATC CONTROL, CONTROLS AND INDICATIONS



LIGHT SENSORThe built in light sensor automatically controls the display brightness. The ANN

PUSH BRT control knob/push button can be used to override the automatic dim


CODE SELECT KNOBS

Two concentric knobs control the active code display. The larger knob changes the

two more significant digits, and the smaller knob changes the two less significantdigits. The less significant digit is incremented or decremented for each detent of

the smaller knob if the knob is slowly turned. Rapid rotation of either knob will cause

changes proportional to the rate of rotation. Rollover of the less significant digits will

occur at 0 and 7, and will cause the more significant digits to be incremented or

decremented. The left two digits and the right two digits are independent of each

other. The various codes used for normal operation are listed in the Aeronautical

Information Manual . Codes 7600 or 7700 are selected for in-flight emergency

operation and will be annunciated by the codes flashing in the active code display

for a couple of seconds before transmission begins

OperationBoth transponders provide identification(mode-A) of the ai rcraft on the ATC groundcontroller’s plan position indicator. TheTDR-94 transponder also provides aircraftpressure altitude to the ground controller’sindicator (mode-C). The transponders aresent altitude data informat ion from the pilot’saltimeter or the copilot’s altimeter (selectionis via an encoding altimeter switch on theaudio panel) (Figure 16-11). In normal mode-A d C ti th TDR94 i i t

operates the “on ground” or “in air” state of themode-S information.

Self Test

To carry out the transponder self-t est posi-

tion the mode switch to ON and select the de-sired transponder to be tested. Set the desiredcode using the code select knobs. Push theTEST button on the CTL-92. During self testthe CTL-92 display flashes from minimumto maximum brightness. If there are no di-

i di i d d d



TEST BUTTONPush theTEST button to initiate the radio self-test routine. In dual version units, the

1/2 switch determines which transponder responds to the test command.

ENCODING ALIMETERSELECT SWITCH

This switch selects which altimeter, the pilot’s (ALTM1) or copilot’s (ALTM2), willprovide encoding altimeter information to the transponders.

SELF-TEST DISPLAY

NO FAILURE

During self test, the active code display intensity will modulate from minimum to

maximum. If the transponder is functioning properly and an altitude encoder is

connected to the CTL-92 and operating, AL will be displayed in the upper window

and the altitude in thousands of feet in100-foot increments will be displayed in the

lower window.

FAILURE

If an out-of-tolerance condition is detected, the upper window shows the word

DIAG while the lower window shows a two-character diagnostic code.

FUNCTION/DESCRIPTIONCONTROL OR

DISPLAY

Table 16-5. CLT-92 ATC CONTROL, CONTROLS AND INDICATIONS (Cont)



A or mode-C operation, the TDR94 is inter-rogated by radar pulses from a ground stationand replies automatically with a series of pulses. The TDR-94/94D can operate inmode-S and provide a unique aircraft iden-tification code as well as air-to-air and air-to-ground interrogation replies. The unitalso has data link capability, which allows itto perform additional air traffic control andaircraft separation assurance functions. TheTDR-94 transponder can also transmit anident pattern when requested to squawk ident

agnostic conditions detected, uncorrectedbarometric altitude is displaye d on the bot-tom display line in hundreds of feet an d the

annunciator AL on the top display l ine(Figure 16-12).

If no altitude data is present the display willshow dashes (- - - -) without a diagnosticcode (Figure 16-12).

If a diagnostic condition is detected in theTDR-94/94D during self test, the upper win-

If during normal operation a “fail/warn” con-dition is detected the CTL-92 will display thedIAG message in the lower window withouthaving pressed the TEST button (Figure 16-12). When this occurs, press the TEST but-ton to view the associated diagnostic code.

AntennasA typical installation of antennas is shownin Figure 16-13.

EFIS and other AvionicsFor specific EFIS and avionics equipment notdiscussed here refer to the appropriate pilotguides, AFM supplements and other appro-priate information.

FLIGHT INSTRUMENTS

PITOT AND STATIC SYSTEM

The pitot and static system (Figure 16-14)provides a source of impact air and static air for operation of the flight instruments. Two

heated pitot masts (Figure 16-15) are locatedon each side of the lower portion of the nose.Tubing from the left pitot mast is connectedto the pilot’s airspeed indicator, and tubingfrom the right pitot mast is connected to thecopilot’s airspeed indicator.




DIAGNOSTICCODE

ALTITUDEANNUNCIATOR

CURRENT AIRCRAFTALTITUDE IN 100FT

INCREMENTS



TEST DISPLAY FAULT PRESENT

FOUR DASHES

TEST DISPLAY NO FAULT PRESENT

ACTIVE CODEDISPLAY




C O M M N o . 1 A N T E N N A

1 A N T E N N A

V O R / L O C A N T E N N A

( L E F T & R I G H T S I D E )



E L T A N T E N N A

( R I G H T S I D E O F F I N )

E L T A N T E N N A

R A D I O A L T I M E T E R A N T E N N A

R A D I O A L T I M E T E R A N T E

N N A

D M E N o .

1 A N T E N N A

D M E N o . 1 A N

T E N N A

N o . 1 T R A N S P O N D E R A N

T E N N A

T R A N S P O N D E R A N T E N N A S

N o . 2 T R A N S P O N D E R A N

T E N N A

A D F A N T E N N A

H F L O

N G W I R E A N T E N N A

F T S I D E V I E W

T O P V I E W

1 6 - 1 3 .

B 2 0 0

A n t e n n a L o c a t i o n s



N T E N N A

E C O N E )

C O M M N o .

N N A G

P S A N T E N N A

2 A N T E N N A

A N T E N N A

N o . 1 A D F A N T E N N A

N O . 2 A D F A N T E N N A

T E N N A

N T E N N A

A AN N A

O N E )

L E

F i g u r e



PILOT'S STATIC

AIR SOURCENORMAL ALTERNATE

+ + +

SEE FLIGHT MANUAL PERFORM-ANCE SECTION FORINSTR CAL ERROR

(POSITIONED ABOVE S1)

S1 S2

DRAIN VALVE

PILOT'S ALTERNATE

STATIC AIR

COPILOT'S STATIC AIR

PILOT'S STATIC AIR

PILOT'S ALTERNATE

STATIC AIR

TO COPILOT'S

INSTRUMENTS

PILOT'S PITOT

COPILOT'S PITOT

DIFFERENTIAL PRESSURE SWITCH

(FOR LANDING GEAR WARNING

ON BB-324 THRU BB-452

THAT ARE NOT IN COMPLIANCE

WITH SI 1047)

A

TO PILOT'S

INSTRUMENTS

PILOT'S STATIC AIR

COPILOT'S STATIC AIR

PRESSURE BULKHEAD

NOTE: ALTIMETERS AND

VERTICAL INDICATORS OMITTED

FROM THIS VIEW FOR CLARITY

PILOT'S STATIC AIR

SOURCE CONTROL VALVE

(VALVE IN "NORMAL" POSITION)

TO

COPILOT'S INSTRUMENTS

TO PILOT'S INSTRUMENTS

COPILOT'S AIRSPEED

INDICATOR

PILOT'S AIRSPEED

INDICATOR

REAR PRESSURE

BULKHEAD

R/H PITOT

MAST

R/H STATIC

PORTS

S1 S2P2

FORWARD PRESSURE

BULKHEAD

BOTTOM TOP



ADS-65

AUTOPILOT

AIR

DATA

SENSOR

MANIFOLD MANIFOLD

ASI

IVSI

ALT

ALT

PPI

CDPI

DRAIN

DRAIN

DRAIN DRAIN

ALTERNATE

STATIC

SELECTOR

VALVE

ALTERNATE

STATIC

PORT

ALT = ALTIMETER

IVSI = INSTANTANEOUS VERTICAL SPEED

INDICATOR

PILOT'S PITOT

COPILOTS PITOT

LEGEND

PNEUMATIC

PRESSURE

CABIN

PRESSURE

APC-65

AUTOPILOT

COMPUTER

The normal static system provides two sepa-

rate sources of static air: one for the pilot’sflight instruments and one for the copilot’sflight instruments. Each of the two static air lines open to the atmosphere through two staticports (Figure 16-16) on each side of the aftfuselage.

An alternate static air line is also provided for the pilot’s flight instruments. In the event of

a failure of the pilot’s normal static air source,which could be caused by ice accumulationsobstructing the static ports (the static portsare not heated), the alternate static source maybe selected by lifting the spring-clip retainer off the PILOT’S STATIC AIR SOURCE valve

of the rear pressure bulkhead from inside theunpressurized area of the fuselage.

The pilot’s airspeed and altimeter nor-mal indications are changed when thealternate static air source is in use.Refer to the Airspeed Calibration-Alternate System, and the Altimeter Correction Alternate System graphsin the Flight Manual (PerformanceSection) for operation when the alter-nate static air source is in use. (The ver-tical speed indicator is also affected,

WARNING



Figure 16-17. Pilot’s Static Air SourceValve Switch

Figure 16-15. Pitot Mast Location



switch (Figure 16-17), located under the copi-lot’s right side circuit-breaker panel, and plac-ing the switch in the ALTERNATE position.

This connects the alternate line to the pilot’sflight instruments only. It obtains static air aft

tical speed indicator is also affected,but no correction table is available.)

When the alternate static air source is not re-quired, the pilot should ensure the PILOT’SSTATIC AIR SOURCE valve switch is held inthe NORMAL (forward) posi t ion by thespring-clip retainer.

OUTSIDE AIR TEMPERATUREGAGE

On later models, the OAT gage (Figure 16-18)is located on the left sidewall panel below thepilot’s left arm. The probe is mounted belowthe pilot’s side window and directly oppositeof the gage. The ON–OFF button for the pos tlights is located next to the gage on the sidepanel.

On BB-1439, 1444 and subsequent, a digitaldisplay is located on the sidewall, and it indi-cates the free air temperature in Celsius. Whenthe adjacent button is depressed, Fahrenheit isdisplayed. The probe is located on the lower fuselage under the pilot position (Figure 16-18).

AUTOFLIGHT SYSTEM

YAW DAMPERA yaw damper function aids the pilot in main-taining directional control of the airplane. Thefunction may be used at any altitude; how-

ever, it is required for flight above 17,000feet. Yaw damping should be deactivated for takeoff and landing.

If the airplane has an autopilot system, theoperation of the yaw damper is covered inthe applicable Flight Manual supplement






(vendor’s manual). If an autopilot system isnot installed, yaw damping functions as anindependent system. The components of thissystem consist of a yaw sensor, amplifier,and a control valve (Figure 16-19).

STALL WARNING

SYSTEM

Th f i f i h

The system has preflight test capability throughthe use of the STALL WARN TEST switch(Figure 16-21) on the copilot’s left subpanel.This switch, held in the TEST position, raisesthe transducer vane, which actuates the warn-ing horn for preflight test purposes.

In the ICE group located on the pilot’s right

subpanel, a STALL WARN switch (Figure16-22) controls electrical heating of the trans-ducer vane and mounting plate.

WARNING



YAW

ENG

AP

ENG

SR

I/20

DN

L

UP

R

YAWALT

YAW

HDG

COLLINS



Figure 16-19. YAW Damp Switch

Figure 16-20. Stall WarningTransducer Vane



The formation of ice at the trans-ducer vane results in erroneous in-dications during flight.

The stall warning system consists of a trans-ducer, a lift computer, a warning horn, and a testswitch. Angle of attack is sensed by air pressureon the transducer vane (Figure 16-20) locatedon the left wing leading edge. When a stall isimminent, the transducer output is sent to a liftcomputer which activates a stall warning horn

COMMUNICATIONSYSTEM

STATIC DISCHARGINGDESCRIPTION

A static electrical charge, commonly referredto as “P” (precipitation static), builds up onthe surface of an airplane while in flight andcauses interference in radio and avionicsequipment operation. The charge is also dan-gerous to persons disembarking after land-i n g , a s w e l l a s t o p e r so n s p e r f o r mi n gmaintenance on the airplane. Fifteen staticwicks (Figure 16-23) are installed on the trail-ing edges of the flight surfaces and the wingtips. The wicks aid in the dissipation of theelectrical charge. Nineteen are installed andonly three may be broken or missing.



Figure 16-22. STALL WARN Heat Switch

(Pilot’s Right Subpanel)



LIMITATIONS

AIRSPEED INDICATOR

Refer to Table 16-6 for airspeed indicator limitations.

OUTSIDE AIR TEMPERATUREGAGEDo not operate the airplane when the outsideair temperature is beyond the following limits:

• Minimum limit at all altitudes is –53.9°C(–65.02°F) for Super King Air 200 and –60°C (–76°F) for Super King Air B200.

• Maximum limit as follows:

1. Sea level to 25,000 feet—ISA +37°C

2. Above 25,000 feet—ISA + 31°C

AUTOPILOTRefer to the applicable FAA-approved Fl ight Ma nu a l supp lement fo r FAR Par t 91Operational Limitations for the autopilot.Except for minimum altitude, refer to the samesupplement for limitations imposed by FARPart 135, Operations, which establishes these

two limitations as well:

1. Enroute—500 feet above terrain is mini-mum altitude.

2. Coupled Approach—Observe decisionheight (DH) or minimum descent altitude(MDA).



MARKING KCAS VALUE KIAS VALUE SIGNIFICANCEOR RANGE OR RANGE

RED LINE 91 86 Air Minimum Control Speed (VMCA)

WHITE ARC 80 to 144 75 to 146 Full-flap Operating Range †80 to 155 †75 to 157

WIDE WHITE ARC 80 to 102 75 to 99 Lower limit is the stalling speed at(VSO) maximum weight with full flaps(ALL AIRPLANES) (100%) and idle power

Table 16-6. AIRSPEED INDICATOR MARKINGS*



(ALL AIRPLANES) (100%) and idle power.

NARROW WHITE ARC 102 to 144 99 to 146 Lower limit is the stalling speed (VS) at†102 to 155 †99 to 157 maximum weight with Flaps Up (0%) and

idle power. Upper limit is the maximumspeed permissible with flaps extendedbeyond approach (more than 40%).

WHITE TRIANGLE 200 200 Maximum flaps to/at approach (40%)(ALL AIRPLANES) speed.

BLUE LINE (ALL AIRPLANES) 122 121 One engine-inoperative best rate ofclimb speed.

RED & WHITE HASH- ‡270 KCAS (269 KIAS) or

1. What is the purpose of the static wicks?

A. To dissipate static electricity

B. To collect static electricity

C. To function as an aerodynamic aidD. To dissipate lightning strikes

2. What is the minimum number of staticwicks required for the Super King Air?

A. 20

B. 25

C. 15

D. 16

3. What instrument does the right pitot mastsupply?

A. Copilot’s airspeed indicator

B. Air data command display

C. Pilot’s airspeed indicator

D. Air data computer

4. When is the yaw damper required?

A. Above 20,000 feet

B. 20,000 feet or above

C 17 000 f b

6. Collins Proline II—The ACT annuncia-tor on the COMM radios:

A. Indicates a malfunction with the ra-dios if it extinguishes.

B. Illuminates only during the self test.

C. Illuminates only when both radios aretuned to the same frequency.

D. Indicates the transceiver is changingfrequencies.

7. Collins Proline II—The AUTO-COMM

switch:A. Allows the NAV radios to automati-

cally identify the frequency whenwithin range.

B. Allows reception of currently selectedCOMM radio without moving its se-lect switch.

C. Only operates during the AUDIO-EMERG mode.

D. Should always be selected for proper COMM radio operation.

8. Collins Proline II—To get proper bearinginformation from the ADF, the selector should be on:



QUESTIONS



C. 17,000 feet or above

D. Above 17,000 feet

5 . Col lins Prol ine I I—The GND COMMPWR button is used to:

A. Speak to line personnel while push-ing/towing the aircraft.

B. Act iva te the r adios whi l e on theground.

C Communicate through COMM 1 while

should be on:

A. ADF or ANT

B. ADF onlyC. ANT only

D. None of the above

9. Collins Proline II—The two transponders:

A. Get information from different en-coding altimeters.

B A b th l i t ATC t ll ti

10. Col l ins P rol ine I I—The DME Holdfunction:

A. Is indicated by the annunciator HLDon the NAV radio.

B. Operates only on the pilot’s side.

C. Eliminates the use of any previouslystored frequencies.

D. Is not possible when an ILS frequencyhas been tuned in the active window.

11. Collins Proline II—The AUDIO EMERG-NORM switch is used:

A. When the GND COMM PWR switchis no longer functional.

B. To bypass the amplifier on the pilotside only.

C. To bypass the amplifier on the copi-lot side only.

D. B and C





CHAPTER 17

MISCELLANEOUS SYSTEMS

CONTENTS

Page

INTRODUCTION................................................................................................................. 17-1

OXYGEN SYSTEMS ........................................................................................................... 17-1

Manual Plug-In System ................................................................................................. 17-4

Autodeployment System................................................................................................ 17-4

TOILET................................................................................................................................. 17-8

RELIEF TUBES.................................................................................................................... 17-8

QUESTIONS....................................................................................................................... 17-10





ILLUSTRATIONS

Figure Title Page

17-1 Oxygen System Diagram (BB-1439, 1444 and Subsequent)................................. 17-2

17-2 Oxygen System Diagram (Prior to BB-1444, Except 1439).................................. 17-3

17-3 Oxygen Mask Stowed ............................................................................................ 17-4

17-4 O2 Mask Selector (BB-1439, 1444 and After—Puritan Bennett) ......................... 17-5

17-5 O2 Mask Selector (Prior to BB-1444, Except 1439) ............................................. 17-5

17-6 First Aid Mask Access Panel ................................................................................. 17-6

17-7 Oxygen System Push-Pull Handles (BB-1439, 1444 and Subsequent) ................ 17-617-8 Oxygen System Push-Pull Handles (Prior to BB-1439) ....................................... 17-6

17-9 Oxygen Bottle and Shutoff Valve ......................................................................... 17-6

17-10 Oxygen System Annunciators ............................................................................... 17-7

17-11 Passenger Oxygen Mask Deployed ....................................................................... 17-7

17-12 Oxygen Available with Partially Full Bottle.......................................................... 17-8

17-13 Toilet ...................................................................................................................... 17-8

17-14 Relief Tube............................................................................................................. 17-9





TABLES

Table Title Page

17-1 Average Time of Useful Consciousness................................................................. 17-5

17-2 Oxygen Duration—200 and B200 ......................................................................... 17-9

INTRODUCTION

The miscellaneous systems include the oxygen system toilet and the relief tubes

CHAPTER 17MISCELLANEOUS SYSTEMS





The miscellaneous systems include the oxygen system, toilet, and the relief tubes .

OXYGEN SYSTEMS

The Super King Air has two oxygen systemsavailable:

1)A plug-in system for SNs BB-2 throughBB-54, and

SLPM (Standard Liters Per Minute). The di-luter-demand crew mask is the only excep-t ion when used in the 100% mode . For computation purposes, each diluter-demandcrew mask being used in the 100% mode counts

t k t 3 7 SLPM



FORWARD PRESSURE BULKHEADCOCKPIT OXYGENPRESSURE GAGE

DILUTER DEMANDCREW MASK

PASSENGER MANUALOVERRIDE HANDLE

CABLE

PASSENGER MANUALOVERRIDE SHUTOFFVALVE

SOLENOID

OFF

ON

BAROMETRICPRESSURESWITCH

CONTROLCABLEPASSENGER 2 MASK OUTLET

(TYPICAL 5 PLACES)

OXYGENPRESSURE GAGE FILL VALVE

ANNUNCIATOR PASS OXYGEN ON

DILUTER DEMAND CREW MASK

CONSOLEPULL ON SYSTEMREADY CONTROL

DETAIL A

DETAIL B

DETAIL C

DETAIL D

B

AD



OXYGEN PRESSURESENSE SWITCH

PASSENGER SINGLE MASK OUTLET

FIRST AID OXYGEN MASK STOWEDIN MANUALLY OPERATED BOX

CONTROL CABLECOMPOSITE OXYGEN CYLINDER

HIGH PRESSURE OVERBOARD RELIEF

AFT PRESSURE BULKHEAD

OPTIONAL OXYGEN MASKCONTAINER, LINES AND

OUTLET FOR FOLD-UP SEATS

CLEGEND



A

OXYGEN PRESSUREGAGE

FILL VALVE

DETAIL B

DETAIL ABAROMETRICPRESSURESWITCH

OFF

ON

PASSENGER MANUALOVERRIDE SHUTOFFVALVE

SOLENOID

OXYGEN OUTLET

PASSENGER MANUALOVERRIDE CONTROL

COCKPIT OXYGENPRESSURE GAGE

FORWARD PRESSURE BULKHEADANNUNCIATOR PASS OXYGEN ON

DILUTER DEMAND CREW MASK

PULL ON SYSTEMREADY CONTROL

OXYGEN OUTLET

PASSENGER 2 MASK OUTLET(TYPICAL 5 PL ACES)

OPTIONAL OXYGEN MASK

IN

OUT

NOTICE: AVIATORS BREATHING OXYGENKEEP FILL AREACLEAN, DRY & FREE FROM OILPRESSURE TO 1850 PSI @ 14.7 PSI & 70°F

O X Y G E N

P R E C HA

R G E

HIGH PRESSURE LINE

LOW PRESSURE LINE

CONTROL CABLE

FLEXIBLE HOSE

LEGEND



B

CONTROL CABLE

FIRST AID OXYGEN MASK STOWEDIN MANUALLY OPERATED BOX

PASSENGER SINGLE OXYGEN MASK

OXYGEN PRESSURESENSE SWITCH

OPTIONAL OXYGEN MASKCONTAINER, LINES AND

OUTLET FOR FOLD-UP SEATS

AFT PRESSURE BULKHEAD

HIGH PRESSURE OVERBOARD RELIEF

STEEL OXYGEN CYLINDER

NOTE:BB-55-309 311–342 344–382

putation is identical with that used on theSuper King Air 200. At cabin altitudes above20,000 feet, the 100% mode is required.

MANUAL PLUG-IN SYSTEM

(Super King Air 200)The manual plug-in system (SNs BB-2 throughBB-54) is the constant-flow type with eachmask plug having its own regulating orifice.The crew oxygen masks are stowed under thepilot’s and copilot’s seats. Oxygen outlets arelocated on the forward cockpit sidewalls. Thepassenger masks are stowed in pockets be-hind the seat backs. However, with respect to

the couch, the masks are stowed underneath.The cabin outlets, located on the cabin head-liner at the top center at the forward and aftends of the cabin, are protected by accessdoors when not in use. Pushing the plug in

firmly and then turning clockwise one-quar-ter turn easily connects the masks. Reversingthis procedure unplugs the mask.

AUTODEPLOYMENT SYSTEM

The autodeployment system (Figures 17-1 and

17-2) is available for all Super King Air air-planes after SN BB-54 and is factory installedon all Super King Air B200 airplanes.

The crew utilizes diluter-demand, quick-don-ning oxygen masks (Figure 17-3) which areheld in the overhead panel. (Prior to BB-1444,except 1439, they hang on the aft cockpit par-tition behind and outboard of the crew seats.)

Since these masks deliver oxygen only upon in-halation, there is no oxygen loss when the masksare plugged in and the PULL ON–SYS READYhandle is pulled out. For BB-1439, 1444 and sub-sequent this is located to the left of the power








quadrant. For prior aircraft it is located aft of the overhead lighting control panel.

The PULL ON–SYS READY han-

dle shall be pulled out to arm theoxygen system prior to flight. Thisis mandatory since the oxygen bot-t le cable or l inkage may freeze.Should this cable or linkage freezewhen the handle is in the OFF po-sition (pushed in), the handle can-not be pulled out, and oxygen wouldnot be available.

Table 17-1 sets forth the average time of use-ful consciousness (time from onset on hy-poxia until loss of effective performance) atvarious altitudes.

On BB-1439, 1444 and subsequent, the crewmask has three modes of operation—normal,100%, and emergency (Figure 17-4). The nor-mal position mixes cockpit air with oxygensupplied through the mask. This mode reducesthe rate of oxygen depletion. When 100% isselected, only oxygen directly from the oxy-gen bottle is breathed by the crewmember.The emergency position supplies a positivepressure to the face piece and should be used

Prior to BB-1444 except BB-1439, a smalllever (Figure 17-5) on each crew mask permitsthe selection of two operational modes— NORMAL and 100%. In the NORMAL posi-tion, cockpit air is mixed with the oxygensupplied through the mask. This mode reducesthe oxygen depletion, plus it is more com-

fortable to use than 100% oxygen. However,when smoke or contaminated air is in the cock-pit, 100% oxygen must be used. The selector levers must be kept in the 100% posi tion whenthe masks are stowed so that no adjustment isnecessary when the masks are donned.

When the primary oxygen supply line is

WARNING

Figure 17-4. O2 Mask Selector (BB-1439, 1444and After—Puritan Bennett)



if smoke and/or fumes are present in the cabin. When the pr imary oxygen supply l ine ischarged, oxygen can be obtained from the first

Table 17-1. AVERAGE TIME OF USEFULCONSCIOUSNESS

35,000 feet ............................... 1/2 to 1 minute

30 000 feet 1 to 2 minutes

aid oxygen mask located in the toilet area.The first aid mask is actuated by manuallyopening the overhead access panel (Figure17-6) marked FIRST AID OXYGEN–PULLand opening the on-off valve inside the box.There is a placard which reads: NOTE: CREWSYS MUST BE ON to remind the user that thePULL ON–SYS READY handle in the cock-pit must be armed before oxygen flows throughthe first aid mask.

The PULL ON–SYS READY push-pull han-dle (Figure 17-7) is located to the left of the

d t (P i t BB 1444 t 1439

next to the PULL ON–SYS READY handle inthe overhead panel; Figure 17-8). Both are op-erated the same way. Pushing in the handledeactivates the selected function, while pullingout the handle actuates the desired function.

The system ready handle operates a cable thatopens and closes the shutoff valve on the oxy-gen bottle (Figure 17-9) in the aft fuselage, be-hind the aft pressure bulkhead. When thehandle is pushed in, no oxygen supply is avail-able anywhere in the airplane. It must be pulledout before engine starting to ensure oxygen isavailable any time it is needed. If the oxygenbottle is not empty when the handle is pulledout, the primary oxygen supply line chargeswith oxygen. This supply line, when charged,delivers oxygen to the two-crew oxygen out-lets, to the first aid oxygen mask, and to themanual override shutoff valve. The crew canmonitor the pressure in the oxygen bottle by

di th th il t’ i ht



Figure 17-6. First Aid Mask Access Panel

Figure 17-8. Oxygen System Push-PullHandles (Prior to BB-1439)



power quadrant (Prior to BB-1444, except 1439it is aft of the overhead light control panel;Figure 17-8). The PASSENGER MANUALO’RIDE (override) push-pull handle (Figure17-7) is located on the right side of the power quadrant (prior to BB-1444, except 1439 it is

reading the oxygen gage on the copilot’s rightsubpanel. It should be noted the filler gage alsoreads bottle and system pressure.

The passenger oxygen system is the constantflow type. Any time the cabin-pressure altitudeexceeds approximately 12,500 feet, a baro-

metric pressure switch automatically energizesa solenoid causing passenger manual overrideshutoff valve to open. Also, the crew can man-ually open this valve any time by pulling out thePASSENGER MANUAL O’RIDE handle. Oncethe shutoff valve is opened, either automati-cally or manually, oxygen flows into the pas-senger oxygen supply line. When this happens,

a pressure-sensitive switch in the supply linecauses the PASS OXY ON advisory annuncia-tor to illuminate (Figure 17-10). On SNs BB-310, -343, -383, -415, -416, -418, -448, -450 andsubsequent (including the B200 airplanes), andwith all serial numbers in the 1979 model year,this switch also causes the cabin lights (whichincludes all fluorescent lights, the vestibulelight, and the center baggage compartment light)to illuminate in the full bright mode.

This occurs regardless of the position of theCABIN LIGHTS switch on the copilot’s leftsubpanel.

Automatic deployment of the passenger con-stant-flow oxygen masks is accomplishedwhen the pressure of the oxygen in the supply

line causes a plunger to extend against eachof the mask dispenser doors, which forces thedoor open (Figure 17-11). When the doorsopen, the masks drop down approximatelynine inches below the doors.

NOTE

The lanyard valve pin at the top of theoxygen mask hose must be pulledout in order for oxygen to f lowthrough the mask.

A lanyard valve pin is connected to the mask

with a flexible cord. When the mask is pulleddown for use, the cord pulls the pin out of thelanyard valve. When this occurs, oxygen willflow continuously from the mask until thepassenger shutoff valve is closed. If the PAS-SENGER MANUAL O’RIDE handle is pushedin (and cabin altitude is below 12,500 feet),or the oxygen control circuit breaker in the en-vironmental group is pulled (regardless of

cabin altitude), this will isolate the remainingoxygen for the crew and first aid outlets. Refer to Table 17-2 for the oxygen duration and seeFigure 17-12 for oxygen bottle capacity.

NOTE (200 AND B200)

For duration time with crew usingdiluter-demand, quick-donning oxy-

gen masks with selector on 100%,increase computation of NUMBEROF PEOPLE USING by a factor of two (e.g., with four passengers, enter this table at eight).



1 EC TRIM OFF



NOTE (200)

Oxygen duration is computed for aPuri tan-Zep oxygen system thatuses either the red, color-coded,

plug-in-type or the autodeployed-t y p e m a s k , b o t h r a t e d a t 3 . 7Standard Liters Per Minute (SLPM)flow. Both are approved for alti-tudes up to 31,000 feet.

NOTE (B200)

area and is enclosed by the cargo partition. Thetoilet may be either the chemical type or theelectrically-flushing type. In either case, thetwo-hinged lid half-sections must be raised togain access to the toilet. A toilet tissue dis-penser is contained in a slide out compartmenton the forward side of the toilet cabinet.

If a Monogram electrically-flushing

CAUTION



Figure 17-12. Oxygen Available withPartially Full Bottle Figure 17-13. Toilet



NOTE (B200)

Oxygen duration is computed for anautodeployed-type mask, 3.9 LitersPer Minute (LPM-NTPD), color-coded orange and white, and ap-proved for altitudes up to 35,000feet.

TOILET

toilet is installed, the sliding knife

valve should be open at all times,except when actually servicing theunit. The cabinet below the toiletmust be opened in order to gain ac-cess to the knife valve actuator han-dle.

RELIEF TUBES



OXYGEN DURATION—200

CYLINDER NUMBER OF PEOPLE USINGVOLUME 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15CU FT DURATION IN MINUTES

22 150 72 48 36 30 24 21 18 16 15 13 12 11 10 *

49 336 168 108 84 66 54 48 42 37 33 30 27 25 24 2264 438 216 144 108 84 72 60 54 48 43 39 36 33 31 28

76 552 261 173 130 104 87 74 66 57 52 47 43 40 37 34

115 792 396 264 198 158 132 113 99 88 79 72 66 60 56 52

OXYGEN DURATION—B200

CYLINDER NUMBER OF PEOPLE USING

VOLUME 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17CU FT DURATION IN MINUTES

22 143 71 47 35 28 23 20 17 15 14 13 11 11 10 * * *49 320 160 106 80 64 53 45 40 35 32 29 26 24 22 21 20 1866 431 215 143 107 86 71 61 53 47 43 39 35 33 30 28 26 2576 496 248 165 124 99 82 70 62 55 49 45 41 38 35 33 31 29115 751 375 250 187 150 125 107 93 83 75 68 62 57 53 50 46 44

OXYGEN DURATION—BB-1439, 1444 and Subsequent

CYLINDER NUMBER OF PEOPLE USINGVOLUME 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 **16 **17

CU FT DURATION IN MINUTES22 144 72 48 36 28 24 20 18 16 14 13 12 11 10 * * *50 317 158 105 79 63 52 45 39 35 31 28 26 24 22 21 19 1877 488 244 162 122 97 81 69 61 54 48 44 40 37 34 32 30 28115 732 366 244 183 146 122 104 91 81 73 66 61 56 52 48 45 43

Table 17-2. OXYGEN DURATION—200 AND B200

* Will not meet oxygen requirements.

** For oxygen duration computations, count each diluter-demand crew mask in use as 2(e g with 4 passengers and a crew of 2 enter the table at 8 people using)



A valve lever is located on the side of the relief tube horn. This valve lever must be depressedat all times while the relief tube is in use. Eachtube drains into the atmosphere through its ownspecial drain port, which protrudes from thebottom of the fuselage. Each drain port atom-izes the discharge to keep it away from the skin

(e.g. with 4 passengers and a crew of 2, enter the table at 8 people using).

1. Where are the crew oxygen masks stowedin the autodeployment system (Serial Nos.BB-55 and subsequent and the B200)?

A. In the side panels

B. In the overhead

C. Under the pilot and copilot seats

D. On the aft partition, outboard andbehind the crew seats

2. W he re a re t he c re w ox yg en o ut le tslocated in the manual plug-in system

(Serial Nos. BB-2 through BB-54)?A. On the forward cockpit sidewalls

B. In the overhead

C. Under the crew seats

D. On the pedestal

3. When do the crew diluter-demand, quick-donning masks deliver oxygen?

A. At all times

B. Upon exhalation

C. Upon inhalation

D. When the hose is plugged-in

4. Why must the PULL ON–SYS READYh dl b ll d i fli h

5. At what cabin-pressure altitude will theautodeployment system operate?

A. 10,500 feet

B. 12,500 feetC. 20,000 feet

D. 31,000 feet

6 . What opens the mask dispenser doorsduring automatic deployment of the pas-senger oxygen masks?

A. Pneumatically-operated solenoid

B. Electrically-operated solenoidC. Oxygen pressure in the supply line

D. Re l e a se o f m ec h a n ic a l l oc k s b ypulling the PASSENGER MANUALO’RIDE handle

7. When selected to the 100% mode, thenumber of crew masks in use should befor computing oxygen duration.

A. Counted once

B. Tripled

C. Halved

D. Doubled

8 Wh h ld h lidi k if l



QUESTIONS



handle be pulled out, prior to flight, arm-

ing the system?A. In case o f oxygen bo t tl e l i nkage

freeze-up

B. To prevent oxygen mask icing

C. To prevent fill valve freeze-up

D. To deenergize the barometric pres-sure switch

8. When should the sliding-knife valve on

a Monogram toilet be open?A. At all times except when actually ser-

vicing the unit

B. At all times including when servicingthe unit

C. Only when servicing the unit

D. Only when in actual use

CHAPTER 18

WEIGHT AND BALANCE/

PERFORMANCE

CONTENTSPage

INTRODUCTION................................................................................................................. 18-1

SYMBOLS, ABBREVIATIONS, AND TERMINOLOGY ................................................. 18-1

General Airspeed Terminology...................................................................................... 18-1

Meteorological Terminology ......................................................................................... 18-3

Power Terminology........................................................................................................ 18-3

Control and Instrument Terminology ............................................................................ 18-4

Graph and Tabular Terminology.................................................................................... 18-4

Weight and Balance Terminology.................................................................................. 18-5

WEIGHT AND BALANCE.................................................................................................. 18-6

Weight and Balance Computation ................................................................................. 18-6

PERFORMANCE.................................................................................................................. 18-9

QUESTIONS 18 16





QUESTIONS....................................................................................................................... 18-16

ILLUSTRATIONS

Figure Title Page

18-1 Basic Empty Weight and Balance Form ................................................................ 18-7

18-2 Weight and Balance Loading Form........................................................................ 18-8

18-3 Loading Data (Passenger) ...................................................................................... 18-9

18-4 Cabin Loading...................................................................................................... 18-10

18-5 Loading Data Cargo Configuration...................................................................... 18-11

18-6 Density Variation of Aviation Fuel ...................................................................... 18-12

18-7 Useful Load Weights and Moments Usable Fuel................................................. 18-13

18-8 Moment Limits vs. Weight................................................................................... 18-14

18-9 Moment Limits vs. Weight with CG.................................................................... 18-15





INTRODUCTION

It is the responsibility of the airplane operator to ensure the airplane is properly loaded.At the time of delivery the manufacturer provides the necessary weight and balance data

CARGO

FUEL

CHAPTER 18WEIGHT AND BALANCE/

PERFORMANCE





At the time of delivery, the manufacturer provides the necessary weight and balance data

to compute individual loading. All subsequent changes in airplane weight and balanceare the responsibili ty of the airplane owner and/or operator. Weight and balance com-putation and considerations are covered in the first portions of this chapter.

Information in this chapter begins with a list of symbols, abbreviations, and terminol-ogy. The weight and balance covers loading a typical airplane and uses data to completea typical computation. Additional information is provided in the airplane flight manual.

GS—Ground Speed is the speed of an airplanerelative to the ground.

IAS—Indicated Airspeed is the speed of an air-plane as shown on the airspeed indicator whencorrected for instrument error. IAS valuespublished in this handbook assume zero in-strument error.

KCAS—Calibrated Airspeed expressed inknots.

KIAS—Indicated Airspeed expressed in knots.

M—Mach Number is the ratio of true airspeedto the speed of sound.

TAS—True Airspeed is the airspeed of an air-plane relative to undisturbed air, which is theCAS corrected for altitude, temperature, andcompressibility.

V—Takeoff Decision Speed.

V2 —Takeoff Safety Speed.

VA —Maneuv eri ng Speed is the max imu mspeed at which application of full availableaerodynamic control will not overstress theairplane.

VF —Des ign Flap Speed is the highest speedpermissible at which wing f laps may beactuated.

VLOF —Lift-off Speed.

VMCA —Air Minimum Control Speed is theminimum flight speed at which the airplane isdirectionally controllable as determined in ac-cordance with Federal Aviation Regulations.The airplane certification conditions includeone engine becoming inoperative and wind-

milling (or inoperative with the autofeather sys-tem armed when Har tzel l propel lers areinstalled), a 5° bank towards the operative en-gine, takeoff power on operative engine, land-ing gear up, flaps in takeoff position, and mostrearward CG. For some conditions of weight andaltitude, stall can be encountered at speedsabove VMCA as established by the certificationprocedure described above, in which event stall

speed must be regarded as the limit of effectivedirectional control.

VMCG —Ground Minimum Control Speed.

VM0 /MM0 —Maximum Operating Limit Speedis the speed limit that may not be deliberatelyexceeded in normal flight operations. V is ex-pressed in knots and M in Mach Number.

VR —Rotation Speed.

V S — S t a l l i n g S p e e d o r t h e m i n i m u msteady f l ight speed at which the airplaneis control lable.

VS0 —Stall ing Speed or the minimum steadyfli ht d t hi h th i l i t l





VFE —Maximum Flap Extended Speed is thehighest speed permissible with wing flaps ina prescribed extended position.

VLE —Maximum Landing Gear ExtendedSpeed is the maximum speed at which an air-plane can be safely flown with the landinggear extended.

flight speed at which the airplane is control-

lable in the landing configuration.

VSSE —Inten tional One-Engine-InoperativeSpeed is speed above both VMCA and stall speed,selected to provide a margin of lateral and di-rectional control when one engine is suddenlyrendered inoperative. Intentional failing of oneengine below this speed is not recommended.



VXSE —One-Engine-Inoperat ive Best Angle-of-Climb Speed is the airspeed, which deliv-ers the greatest gain in altitude in the shortestpossible horizontal distance with one engineinoperative.

Vy —Be st Rate -of -Cli mb Sp eed is the ai r-speed which delivers the greatest gain in al-

titude in the shortest possible time with gear and flaps up.

VYSE —One-Engine- Inoperat ive Best Ra te-of-Climb Speed is the airspeed which deliv-ers the greatest gain in altitude in the shortestpossible time with one engine inoperative.

METEOROLOGICALTERMINOLOGY

Altimeter Setting—Barometric Pressure cor-rected to sea level.

Indicated Pressure Altitude—The number ac-tually read from an altimeter when the baro-metric subscale has been set to 29.92 inches

of mercury (1013.2 millibars).

IOAT—Indicated Outside Air Temperature is thetemperature value read from an indicator.

ISA—International Standard Atmosphere inwhich:

(1) The air is a dry perfect gas.

Pressure Altitude—Altitude measured fromstandard sea-level pressure (29.92 in Hg) bya pressure (barometric) altimeter. It is the in-dicated pressure altitude corrected for positionand instrument error. In this handbook, al-timeter instrument errors are assumed to bezero. Position errors may be obtained fromthe Altimeter Correction graphs.

Station Pressure—Actual atmospheric pres-sure at field elevation.

Temperature Compressib ility Effects—Ane r r o r i n t h e i n d i c a t i o n o f t e m p e r a t u r ecaused by airf low over the temperatureprobe. The error varies, depending on alti-tude and airspeed.

Wind—The wind velocities recorded as vari-ables on the charts of this handbook are to beunderstood as the headwind or tailwind com-ponents of the reported winds.

POWER TERMINOLOGY

Beta Range—The region of the Power Lever control which is aft of the Idle Stop and for-ward of reversing range where blade pitchangle can be changed without a change of gasgenerator rpm.

Cruise Climb—Is the maximum power ap-proved for normal climb. These powers aretorque or temperature (ITT) limited.



(2) The temperature at sea level is 15°C (59°F).

(3) The pressure at sea level is 29.92 inchesof mercury (1013.2 millibars).

(4) Th e t e m p e r a t u r e g r a d i e nt f r o m se alevel to the al t i tude at which the t em-p e r a t u r e i s – 5 6 . 5 ° C ( – 6 9 . 7 ° F ) i s

H i g h I d l e — O b t a i n e d b y p l a c i n g t h eCondition Lever in the High Idle position.This limits the power operation to a minimumof 70% of N1 rpm.

Low Idle—Obtained by placing the ConditionLever in the Low Idle position. This limitsthe power operation to a minimum of 52%

Maximum Cruise Power—Is the highest power rating for cruise and is not time limited.

Propeller Ground Fine—Propeller ground fineoperation is used to provide deceleration onthe ground during landing and accelerate-stopconditions by taking advantage of the maxi-mum available propeller drag without creat-

ing negative thrust.

Reverse—Reverse thrust is obtained by lift-ing the Power Levers and moving them aft of the Beta range.

SHP—Shaft Horsepower.

Takeoff Power—Is the maximum power rat-

ing and is limited to a maximum of five min-utes operation. Use of this rating should belimited to normal takeoff operations and emer-gency situations.

CONTROL AND INSTRUMENTTERMINOLOGY

Condition Lever (Fuel Shutoff Lever)—Thefuel shutoff lever actuates a valve in the fuelcontrol unit, which controls the flow of fuelat the fuel control outlet and regulates the idlerange from Low to High Idle.

ITT (Interstage Turbine Temperature)—Eightprobes wired in parallel indicate the tem-perature between the compressor and power

Propeller Governor—This governor will main-tain the selected speed requested by the pro-peller control lever, except on reverse selectionwhere the power lever interconnection to theintegral pneumatic area of the governor willselect a lower speed. The pneumatic area dur-ing normal selection will act as an overspeedlimiter.

Torquemeter—The torquemeter system de-termines the shaft output torque. Torque val-ues are obtained by tapping into two outletson the reduction gear case and recording thedifferential pressure from the outlets. The re-lationship between torquemeter pressure andpropeller shaft power is shown in the LIMI-TATIONS section. Instrument readout is in

foot-pounds.

GRAPH AND TABULARTERMINOLOGY

Accelerate-Go—The distance to accelerate toTakeoff Decision Speed (V1), experience anengine failure, continue accelerating to lift-off

speed (VLOF) then climb and accelerate inorder to achieve Takeoff Safety Speed (V2) at35 feet above the runway.

Accelerate-Stop—The distance to accelerateto Takeoff Decision Speed (V1) and stop, usingbrakes and propeller reversing on the opera-tive engine. V1 speed is equal to the takeoff rotation speed (VR).





p p p

turbines.

N1 Tachometer (Gas Generator RPM)—Thetachometer registers the rpm of the gas gen-erator with 100% representing a gas genera-tor speed of 37,500 rpm.

Power Lever (Gas Generator N1 rpm)—Thislever serves to modulate engine power from

p ( R)

AGL—Above Ground Level.

Best Angle-of-Climb Speed—The best angle-of-climb speed is the airspeed which deliversthe greatest gain of altitude in the shortestpossible horizontal distance with gear andflaps up.



Clearway—A clearway is an area beyond theairport runway not less than 500 feet wide, cen-trally located about the extended centerline of the runway, and under the control of the airportauthorities. The clearway is expressed in termsof a clear plane, extending from the end of therunway with an upward slope not exceeding1.25%, above which no object nor any terrain

protrudes. However, threshold lights may pro-trude above the plane if their height above theend of the runway is 26 inches or less and if theyare located to each side of the runway.

Climb Gradient—The ratio of the change inheight during a portion of a climb, to the hor-izontal distance traversed in the same timeinterval.

Demonstrated Crosswind—The maximum 90°crosswind component for which adequate con-trol of the airplane during takeoff and landingwas actually demonstrated during certification.

MEA—Minimum Enroute Altitude.

Net Gradient of Climb—The gradient of climb

with the flaps in the takeoff position, and thelanding gear retracted. Net indicates that theactual gradients of climb have been reducedby .8% to allow for turbulence and pilot tech-nique. The Net Gradient of Climb graphs areconstructed so the value(s) obtained using theairport pressure altitude and outside air tem-perature will be the average gradient from 35feet above the runway up to 1,500 feet above

WEIGHT AND BALANCETERMINOLOGY

Approved Loading Envelope—Those combi-nations of airplane weight and center of grav-ity which define the limits beyond whichloading is not approved.

Arm—The distance from the center of grav-ity of an object to a line about which momentsare to be computed.

Basic Empty Weight—The weight of an emptyairplane including full engine oil and unusablefuel. This equals empty weight plus the weightof unusable fuel, and the weight of all the en-gine oil required to fill the lines and tanks.

Basic empty weight is the basic configurationfrom which loading data is determined.

Center of Gravity—A point at which theweight of an object may be considered con-centrated for weight and balance purposes.

CG limits—The extreme center of gravity lo-cations within which the airplane must be op-

erated at a given weight.

Datum—A vertical plane perpendicular to theairplane longitudinal axis from which fore-and-aft (usually aft) measurements are madefor weight and balance purposes.

Empty Weight—The weight of an empty air-plane before any oil or fuel has bee n added



feet above the runway up to 1,500 feet above

the runway.

Route Segment—A part of a route. Each endof that part is identified by either:

(1) A geographic location, or

(2) A point at which a definite radio fix canbe established.

plane before any oil or fuel has bee n added.

This includes a l l permanent ly- insta l ledequipment, fixed ballast, full hydraulic fluid,full chemical toilet fluid, and all other fulloperating fluids, except that the engines,tanks, and lines do not contain any engine oilor fuel.

Engine Oil—That portion of the engine oil

Leveling Points—Those points which areused during the weighing process to levelthe airplane.

Maximum Weight—The largest weight al-lowed by design, structural, performance, or other limitations.

Moment—A measure of the rotational ten-dency of a weight, about a specified line,mathematically equal to the product of theweight and the arm.

Payload—Weight of occupants, cargo andbaggage.

PPH—Pounds Per Hour.

Ramp Weight—The airplane weight at enginestart assuming all loading is completed.

Standard Empty Weight—The basic emptyweight of a standard airplane as specified bythe manufacturer.

Station—The longitudinal distance from some

point to the zero datum or zero-fuselage station.

Takeoff Weight—The weight of the airplaneat lift off from the runway.

Tare—The weight which may be indicated byscales before any load is applied.

Unusable Fuel—The fuel remaining after con-

WEIGHT AND BALANCE

A current record of airplane basic weight andbalance must be maintained at all times. Thismay require periodic weighing. The need for weighing is determined through maintenanceprocedures and practices.

A Basic Empty Weight and Balance form isprovided by the airplane manufacturer tomaintain the record in a current condition(Figure 18-1). A careful check should bemade each time this form is used to ensurethe information is current. A sample basicempty weight and moment of a typical air-plane is used in this chapter.

WEIGHT AND BALANCECOMPUTATION

Another form is provided by the airplanemanufacturer to compute weight and bal-ance (Figure 18-2). This computation shouldbe completed before each flight to ensureproper airplane operation, both on the ground

and in flight.

One or two pilots’ weights and moments/100are recorded on REF line 2. The moment/100is the product of p i lo t weight t imes theCentroid Fuselage Station (F.S.) of the crew(Figure 18-3). This same method is used tocompute the moment/100 of Pilot’s Baggageand Extra Equipment. The F.S. used will de-





Unusable Fuel The fuel remaining after con

sumption of usable fuel.

Usable Fuel—That portion of the total fuelwhich is available for consumption as deter-mined in accordance with applicable regula-tory standards.

Useful Load—The difference between air-

and Extra Equipment. The F.S. used will de

pend on where the items are actually placedin the airplane and can be computed from theairplane loading data.

Payload computations are made in the leftcolumn. Both passenger and cargo can beadded in this column. The F.S. components for p a s se n g e r s a r e p r o v i d e d o n t h e Ca b i n









Figure 18-2. Weight and Balance Loading Form



The Zero-Fuel Condition is the total of weightsand moments/100 in REF lines 1 through 5.This weight must not exceed 10,400 poundsin the 200, and must not exceed 11,000 pounds

in the B200.

Fuel is added by first referring to the DensityVariation of Aviation Fuel chart (Figure 18-6)to determine the fuel density (Note: For aircraftflight manual purposes 1kg/L = 8.345 lb/gal).Then the weight and moment of the fuel loadedis determined from the Useful Load Weights andMoments Useful Fuel chart (Figure 18-7).

The total weight and moment/100 for RampCondition is the sum of Zero-Fuel conditionplus the Fuel Loading. The fuel weight and mo-ment/100 for start, taxi, and takeoff is listedat the bottom of the form. The total weight for Takeoff Condition must not exceed 12,500pounds. The computed Takeoff Condition mo-ment/100 should be checked to be within lim-

Fuel to destination is computed during theperformance functions of flight planning.This fuel figure is used in REF line 11 to com-pute the Landing Condition. The moment/100

for fuel to destination is computed by sub-tracting the moment/100 for the fuel re-m a i n i n g ( a t l a n d i n g ) f r o m t h e f u e lmoment /100 used for fuel loading. UseFigure 18-7 for this computation.

Landing Condition is computed by subtract-ing the fuel to be used from Takeoff Condition.The moment/100 for Landing Condition should

be checked using Figure 18-8 to ve rify that itis within limits.

PERFORMANCE

The material for performance computationtraining for the Super King Air begins on the

Figure 18-3. Loading Data (Passenger)



its. (Figures 18-8 and 18-9.) When using theMoment Limits vs. Weight Graph, move hor-izontally with the weight of the airplane to thepoint where it meets the airplane’s moment/100which is presented diagonally. The point wherethe two meet shows the center of gravity ininches aft of the datum.

following page.

The material in the Performance section of this manual is not copyrighted by FlightSafetyInternational, Inc. Because of the critical na-ture of this data, the material has been repro-duced di rect ly f rom the manufacturer ’spublication.

























1. What is the maximum takeoff weight of the Super King Air?

A. 12,590 pounds

B. 12,500 pounds

C. 12,350 poundsD. 12,200 pounds

2. What is the maximum zero-fuel weight for the Super King Air 200?

A. 10,100 pounds

B. 10,200 pounds

C. 10,300 pounds

D. 10,400 pounds

3. What is the maximum zero-fuel weight for the Super King Air B200?

A. 10,800 pounds

B. 10,960 pounds

C. 11,100 pounds

D. 11,000 pounds

4. W ha t m om en t/ 10 0 f ig ur e s ho ul d b echecked within limits on each flight?

A. Takeoff, zero fuel

B. Takeoff, ramp

C. Takeoff, landing

D Takeoff in flight

GIVEN:An airplane with basic weight of 8,087pounds and moment/100 of 15,041. Theairplane is loaded with two 170-poundpilots, four 170-pound passengers (two

in seats at F.S. 212 and two in seats atF.S. 259), 160 pounds of baggage, and40 pounds of refreshment in the for-ward cabinet.

5. What is the zero-fuel weight and mo-ment/100 of the given airplane?

6. If the fuel gages read a total of 603pounds after the previous flight, howmuch fuel can be added for flight ingallons?

7. Wil l the takeoff weigh t and mo-ment/100 be within limits with thefuel loading in Question 6?



QUESTIONS



D. Takeoff, in flight


CONTENTS

Page

FLIGHT MANEUVERS AND PROFILES...................................................................... GEN-1

Takeoff....................................................................................................................... GEN-1

FLIGHT PROFILES......................................................................................................... GEN-1

LANDING ...................................................................................................................... GEN-18

Flaps-Up Approach and Landing ............................................................................ GEN-18

Single-Engine Approach and Landing .................................................................... GEN-18

Crosswind Approach and Landing.......................................................................... GEN-18

WINDSHEAR................................................................................................................. GEN-18

General .................................................................................................................... GEN-18

Microbursts.............................................................................................................. GEN-19

Acceptable Performance Guides ............................................................................. GEN-19

COCKPIT RESOURCE MANAGEMENT.................................................................... GEN-20





ILLUSTRATIONS

Figure Title Page

GEN-1 Normal Takeoff and Departure................................................................... GEN-2

GEN-2 Engine Loss at or Above V1 ....................................................................... GEN-3

GEN-3 Rejected Takeoff......................................................................................... GEN-4

GEN-4 Steep Turns ................................................................................................. GEN-5

GEN-5 Approach to Stall—Clean........................................................................... GEN-6

GEN-6 Approach to Stall—Takeoff Configuration ................................................ GEN-7

GEN-7 Approach to Stall—Landing Configuration ............................................... GEN-8

GEN-8 Emergency Descent .................................................................................... GEN-9

GEN-9 Standard Holding Pattern—Direct Entry ................................................. GEN-10

GEN-10 Standard Holding Pattern—Teardrop Entry ............................................. GEN-11

GEN-11 Standard Holding Pattern—Parallel Entry ............................................... GEN-12

GEN-12 Visual Approach and Landing.................................................................. GEN-13

GEN-13 One Engine Inoperative—Visual Approach and Landing........................ GEN-14

GEN-14 ILS Approach—Landing in Sequence from an ILS................................. GEN-15

GEN-15 Non-Precision Approach—Procedure Turn ............................................. GEN-16

GEN-16 Circling Approach and Landing............................................................... GEN-17

GEN 17 Si i l A i h C k i GEN 20





GEN-17 Situational Awareness in the Cockpit....................................................... GEN-20

GEN-18 Command and Leadership........................................................................ GEN-20

GEN-19 Communication Process........................................................................... GEN-21

GEN-20 Decision-Making Process......................................................................... GEN-21

FLIGHT MANEUVERSAND PROFILES

TAKEOFF






TAKEOFF

Crosswind TakeoffFollow procedures for normal takeoff except:

• Hold aileron into wind.

Obstacle Clearance Takeoff

Follow procedures for normal takeoff except:

• Maintain V2 until clear of obstacle.



TAKEOFF

IN POSITION

BEFORE TAKEOFF

VYSE OR ABOVE

CLIMB-OUT

1. CHECKLIST — COMPLETE

1. HOLD BRAKES

2. PROPS — 2,000 RPM

(ON GOVERNORS)

3. RELEASE BRAKES

4. SET TORQUE

TAKEOFF ROLL

1. RECHECK TORQUE/ITT

2. ANNUNCIATORS — CHECK

1. ROTATE AT V1 TO

APPROX 7˚ NOSE UP

2. ESTABLISH POSITIVE

RATE OF CLIMB

3. LANDING GEAR — UP

1. FLAPS — UP

2. YAW DAMPER — ON

3. CLIMB POWER — SET

1. ACCELERATE TO

160 KIAS

2. LANDING/TAXI

LIGHTS — OUT

3. COMPLETE CLIMB

CHECKLIST

AREA DEPARTURE/CLIMBPROFILE

1. 160 KIAS TO 10,000 FT

2. 140 KIAS 10,000 - 20,000 FT

3. 130 KIAS 20,000 - 25,000 FT

4. 120 KIAS 25,000 - 35,000 FT

CRUISE

1. ACCELERATE TO

CRUISE SPEED

2. SET CRUISE POWER

3. COMPLETE CRUISE

CHECKLIST



2. RECHECK V1 AND V2

Figure GEN-1. Normal Takeoff and Departure



V2

TAKEOFF

BEFORE TAKEOFF

1,000 FT AGL

ENGINE LOSS

1. FOLLOW NORMAL TAKEOFF

PROCEDURES UNTIL AT OR

ABOVE V1

1. ROTATE AT V1 TO

APPROX 7˚ NOSE UP

2. ESTABLISH POSITIVE

RATE OF CLIMB

3. LANDING GEAR — UP

1. MAINTAIN RUNWAY HEADING

1. CHECK MAX POWER

2. AIRSPEED AT V2

3. VERIFY PROP FEATHERED

1. VYSE (BLUE LINE)

2. FLAPS — UP

CLIMB

1. COMPLETE ENGINE FAILURE

CHECKLIST CLEAN-UP ITEMS

2. LAND AS SOON AS PRACTICAL

NOTE: IT MAY BE NECESSARY TO BANK AS MUCH AS 5˚

INTO THE GOOD ENGINE TO MAINTAIN RUNWAY

HEADING. IT WILL TAKE ALMOST FULL RUDDERON THE SIDE OF THE GOOD ENGINE TO KEEP

NOTE:

DO NOT RETARD FAILED ENGINE POWER LEVERUNTIL THE AUTOFEATHER SYSTEM HAS COMPLETELY

STOPPED PROPELLER ROTATION.



THE BALL SLIGHTLY OFF CENTER.

Figure GEN-2. Engine Loss at or Above V1



BEFORE TAKEOFF

EMERGENCY OR MALFUNCTIONAT OR BELOW V1

CLEAR OF RUNWAY

1. FOLLOW NORMAL TAKEOFF

PROCEDURES UNTIL INITIATING

ABORT AT OR BELOW V1

1. RECOGNIZE REASON FOR

REJECTING TAKEOFF

2. POWER LEVERS — IDLE

3. BRAKING — AS NECESSARY

4. REVERSE — AS NECESSARY

5. MAINTAIN RUNWAY HEADING

1. COMPLETE AFTER

LANDING CHECKLIST

NOTE:

IF REJECTED TAKEOFF IS DUE TO REASONS

OTHER THAN ONE ENGINE POWER LOSS,

REVERSE IS MOST EFFECTIVE AT HIGH SPEEDS;

BRAKING IS MOST EFFECTIVE AT LOW SPEEDS.

Figure GEN-3. Rejected Takeoff





ROLLOUT

1. RETURN TO AND

HOLD ENTRYPARAMETERS

THROUGH 30° BANK

1. ADD APPROX 100 LBS TORQUE

2. ONE UNIT NOSEUP TRIM

3. SMALL PITCH INCREASE

THROUGH 30° BANK

1. REDUCE TORQUE 100 LBS2. REDUCE PITCH

3. TAKE OUT TRIMROLL INTO TURN

1. MAINTAIN INITIALALTITUDE

HOLD 45° BANK

1. SMALL PITCH CORRECTIONS

2. MAINTAIN AIRSPEED

ROLL OUT OF TURN

1. START ROLLOUT 25° PRIORTO ROLLOUT HEADING



INITIAL ENTRY

1. AIRSPEED — 180 KNOTS

2 TORQUE APPROX 1 000 1 200 LBS



BEGINNING OF MANEUVER STALL AND RECOVERY COMPLETION OF MANEUVER

HORNOR BUFFET

INITIAL CONDITION:

1. TORQUE — 200 LBS

2. PROPELLERS — 1,700 RPM

3. MAINTAIN INITIAL HEADING

4. MAINTAIN INITIAL ALTITUDE5. PITCH ATTITUDE PRIOR TO HORN

OR BUFFET MAY REACH 10˚-15˚,

DEPENDING ON TECHNIQUE

6. HORN WILL SOUND APPROX

10 KTS ABOVE BUFFET

AT HORN OR BUFFET — RECOVER:

1. SIMULTANEOUSLY ADVANCE THE POWER

LEVERS TOWARD MAX TORQUE, REDUCE

THE PITCH ATTITUDE AS NECESSARY TO

STOP THE STALL WARNING, AND ROLLTHE WINGS LEVEL

2. ESTABLISH POSITIVE RATE OF CLIMB

COMPLETION:

1. LEVEL OFF AT NEW

ALTITUDE AND INITIAL HEADING

2. RESET POWER AS REQUIRED

V2



OR BUFFET

Figure GEN-5. Approach to Stall—Clean




HORNOR BUFFET


1. REDUCE THE PITCH ATTITUDE AS

NECESSARY TO STOP THE STALL

WARNING, AND ROLL THE WINGS LEVEL

2. ESTABLISH POSITIVE RATE OF CLIMB3. FLAPS — UP, AT OR ABOVE VYSE (BLUE LINE)

COMPLETION:

1. LEVEL OFF AT NEW



V2

INITIAL CONDITION:



3. MAINTAIN INTITIAL HEADING

4. MAINTAIN INTITIAL ALTITUDE 5. FLAPS — APPROACH

(BELOW TRIANGLE)

6. AT 110 KIAS OR LESS,

SIMULTANEOUSLY SET THE

TORQUE TO 1,100 LBS

(SIMULATED 100% TORQUE),

ESTABLISH A BANK ANGLE

OF 20˚ (NO MORE THAN 30˚), AND

RAISE THE NOSE AND CLIMB

7. STUDENT MAY BE REQUIRED TO

PERFORM THIS MANEUVER

WHILE MAINTAINING 15˚ - 30˚ ANGLE OF BANK OR WHILE

MAINTAINING A HEADING

8. CLEAR AREA IN DIRECTION

OF TURN

9. DECREASE SPEED APPROX

1 KT PER SECOND

10. PITCH ATTITUDE PRIOR TO

HORN OR BUFFET MAY

REACH 15˚ - 25˚, DEPENDING

ON TECHNIQUE



OR BUFFET

Figure GEN-6. Approach to Stall—Takeoff Configuration




HORNOR BUFFET


1. SIMULTANEOUSLY ADVANCE THE POWER

LEVERS TOWARD MAX TORQUE, PROPELLER

LEVERS FULL FORWARD, REDUCE THE PITCH

ATTITUDE AS NECESSARY TO STOP THE

STALL WARNING, AND ROLL THE WINGS LEVEL

2. ESTABLISH POSITIVE RATE OF CLIMB

3. FLAPS — UP, AT OR ABOVE 100 KIAS

4. GEAR — UP

COMPLETION:

1. LEVEL OFF AT NEW



V2

INITIAL CONDITION:



3. MAINTAIN INTITIAL HEADING

4. MAINTAIN INTITIAL ALTITUDE

5. FLAPS — APPROACH

(BELOW TRIANGLE)

6. GEAR — DOWN (BELOW VLE)

7. FLAPS — DOWN 100%

(BELOW TOP OF WHITE ARC)8. PITCH ATTITUDE PRIOR TO

HORN OR BUFFET MAY

REACH 10˚ - 15˚, DEPENDING

ON TECHNIQUE

9. HORN WILL SOUND APPROX

10 KTS ABOVE BUFFET

Figure GEN 7 Approach to Stall Landing Configuration



Figure GEN-7. Approach to Stall—Landing Configuration



INITIAL

1. OXYGEN SYSTEM — VERIFYARMED

2. CREW MASK — ON

3. MIC SWITCH — OXYGEN MASKPOSITION

4. SPEAKER (AS REQUIRED)

5. PASSENGER OXYGEN (ASREQUIRED)

6. POWER LEVERS — IDLE

7. PROP LEVERS — SMOOTHLYFULL FORWARD

8. FLAPS — APPROACH (BELOW

TRIANGLE)

9. GEAR — DOWN (BELOW VLE)

DESCENT

1. INITIAL PITCH ATTITUDE — 20° NOSEDOWN

2. PRIOR TO VLE, REDUCE PITCHATTITUDE TO APPROXIMATELY 14° NOSEDOWN

3. MAXIMUM IAS SHOULD BE VLE

4. ADVISE ATC

5. RESET ALTIMETER AND ALTITUDEALERTER TO LEVEL-OFFALTITUDE

LEVEL-OFF

1. APPROXIMATELY 500 FEETBEFORE LEVEL-OFF ALTITUDE,SMOOTHLY REDUCE RATE OFDESCENT

2. FLAPS — UP

3. GEAR — UP (BELOW VLO

RETRACTION)

4. ADD POWER AS REQUIRED

5. MIC SWITCH — NORMALPOSITION

6. REMOVE MASK

7. SET PROP RPM

8. COMPLETE DESCENT CHECKLIST

REDUCE RATE OF DESCENTAPPROXIMATELY 500 FEETABOVE LEVEL-OFF ALTITUDE

20° NOSEDOWN

VLE — APPROXIMATELY14° NOSEDOWN

LEVEL OFF



9. GEAR DOWN (BELOW VLE)

NOTE: IF INITIAL INDICATED AIRSPEED IS

ABOVE VLE, MAINTAIN THE INITIALALTITUDE UNTIL THE IAS IS AT ORBELOW VLE.

NOTE: DESCENT FROM 35,000 TO 12,500

FEET REQUIRES APPROXIMATELYSIX MINUTES



TURN INBOUND

1. ADJUST LEG LENGTH TO PROVIDE 1MINUTE AT 14,000 FEET AND BELOWOR 1.5 MINUTES ABOVE 14,000 FEET

70°

110°

INITIAL

1. SLOW TO HOLDING AIRSPEED* —

160 KIAS WITHIN 3 MINUTES OF FIX

2. TORQUE — APPROX 800-1,000 LBS



ENTERING HOLDING PATTERN

1. REPORT ENTERING HOLD

2. TURN TO PARALLEL OUTBOUND COURSE

3. START TIMING OVER OR ABEAM FIX,

WHICHEVER OCCURS LATER

*MAX HOLDING SPEEDS

• 6,000 FEET & BELOW — 200 KIAS

• 6,001-14,000 FEET — 230 KIAS

• 14 001 & ABOVE 265 KIAS



TURN INBOUND

1. ADJUST LEG LENGTH TO PROVIDE 1MINUTE AT 14,000 FEET AND BELOW

OR 1.5 MINUTES ABOVE 14,000 FEET

70°



2. TURN 30° FROM OUTBOUND COURSE

3. START TIMING OVER FIX



*MAX HOLDING SPEEDS


• 6,001-14,000 FEET — 230 KIAS

14 001 & ABOVE 265 KIAS

INITIAL

1. SLOW TO HOLDING AIRSPEED* —

160 KIAS WITHIN 3 MINUTES OF FIX

2 TORQUE APPROX 800 1 000 LBS



TURN INBOUND

1. ADJUST LEG LENGTH TO PROVIDE 1

MINUTE AT 14,000 FEET AND BELOWOR 1.5 MINUTES ABOVE 14,000 FEET

110°

*MAX HOLDING SPEEDS


• 6,001-14,000 FEET — 230 KIAS

• 14,001 & ABOVE — 265 KIAS


1 REPORT ENTERING HOLD




2. TURN TO PARALLEL OUTBOUND COURSE

3. START TIMING OVER OR ABEAM FIX,WHICHEVER OCCURS LATER

INITIAL



INITIAL

1. OBTAIN ATIS

2. DESCENT CHECKLIST —

COMPLETE

ARRIVAL

1. TORQUE — APPROX 800 LBS

2. 150 - 175 KIAS (TYPICAL)

3. START BEFORE LANDING

CHECKLIST

DOWNWIND

1. FLAPS — APPROACH2. 130 - 140 KIAS

ABEAM TOUCHDOWN POINT

1. GEAR — DOWN

2. BEFORE LANDING CHECKLIST —

COMPLETE

BASE

1. 130 KIAS (MIN REC)

FINAL

1. 130 - 140 KIAS (VYSE MIN)

WHEN LANDING ASSURED:

2. FLAPS — DOWN

3. TRANSITION TO VREF

4. YAW DAMPER — OFF

LANDING

1. PROPS — FULL FORWARD

2. BETA OR REVERSE

3. BRAKES — AS NECESSARY

THRESHOLD

1. GEAR — RECHECK

DOWN

2. AIRSPEED — VREF

3. POWER — IDLE

REJECTED LANDING

1. POWER — MAX

2. PITCH — 10˚ NOSE UP

3. AIRSPEED — 100 KIAS

4. ESTABLISH NORMAL CLIMB

WHEN CLEAR OF OBSTACLES

5. FLAPS — UP

6. GEAR — UP

CAUTION

O S CO S S G



TO ENSURE CONSTANT REVERSING

CHARACTERISTICS, THE PROPELLER

CONTROL MUST BE IN FULL INCREASE

RPM POSITION.

NOTE: REVERSE IS MOST EFFECTIVE AT

CAUTION

IF POSSIBLE, PROPELLERS SHOULD BE MOVED OUT OF



INITIAL

1. OBTAIN ATIS


COMPLETE

ARRIVAL

1. TORQUE — APPROX 1,600 LBS


3. START ONE-ENGINE-INOPERATIVE

APPROACH AND LANDING CHECKLIST

DOWNWIND


2. 130 - 140 KIAS

ABEAM TOUCHDOWN POINT

1. GEAR — DOWN

2. PROP — FULL FORWARD

BASE

FINAL

1. 130 - 140 KIAS (VYSE MIN)


2. FLAPS — DOWN

LANDING

1. BETA OR REVERSE —

AS NECESSARY2. BRAKES — AS NECESSARY

THRESHOLD

1. GEAR — RECHECK

DOWN


3. POWER — IDLE

GO-AROUND

1. POWER — MAX

2. GEAR — UP

3. FLAPS — UP

4. AIRSPEED — INCREASE TOVYSE (BLUE LINE)



1. 130 KIAS (MIN REC) 3. TRANSITION TO VREF


5. ONE-ENGINE-INOPERATIVE

APPROACH AND LANDING

CHECKLIST — COMPLETE



INITIAL

ARRIVAL

1. OBTAIN ATIS

2. REVIEW APPROACH AND

MISSED APPROACH

3. NAVAIDS — TUNE/IDENT


COMPLETE

1. TORQUE — APPROX800 LBS


3. FD — AS DESIRED

4. START BEFORE

LANDING CHECKLIST

APPROACH INBOUND


2. 130 - 140 KIAS

APPROACHING GLIDE SLOPE

1. GEAR — DOWN

2. COMPLETE BEFORE LANDING CHECKLIST

DH-VISUAL AND LANDING ASSURED

1. FLAPS — DOWN



DH

MM

OM

THRESHOLD

1. GEAR — RECHECK DOWN

LANDING


DH-MISSED APPROACH

1. POWER — MAX

2. PITCH — 7˚ - 8˚ NOSE UP (FD-GA)

3. FLAPS — UP

4. GEAR — UP

5. COMPLETE MISSED APPROACH PROCEDURE

GLIDE SLOPE INTERCEPT

1. TORQUE — APPROX

600 - 800 LBS

2. 130 - 140 KIAS (VYSE MIN)



1. GEAR RECHECK DOWN


3. POWER — IDLE

1. PROPS FULL FORWARD

2. BETA OR REVERSE


CAUTION CAUTION



FAF

FAF

INITIAL

1. OBTAIN ATIS

2. REVIEW APPROACH AND

MISSED APPROACH

3. NAVAIDS — TUNE/IDENT


COMPLETE

ARRIVAL




4. START BEFORE LANDING

CHECKLISTSTATION PASSAGE

1. START TIMING

2. SET ALTITUDE ALERTER

1. START TIMING


3. 130 - 140 KIAS

PROCEDURE TURN OUTBOUND


2. RESET ALTITUDE ALERTER

PROCEDURE TURN INBOUND

INTERCEPT FINAL APPROACH

1. COURSE INBOUND

APPROACH INBOUND

1. RESET ALTITUDE ALERTER

FINAL APPROACH FIX

1. START TIMING

2. GEAR — DOWN


4. COMPLETE BEFORE

LANDING CHECKLIST

5. 130 - 140 KIAS

MINIMUM DESCENT ALTITUDE (MDA)

1. LEVEL OFF AT MDA AT LEAST 1 MILE

MDA

MAP

MAP-MISSED APPROACH

1. POWER — MAX

2. PITCH — 7˚ - 8˚ NOSE UP (FD-GA)

3. FLAPS — UP

4. GEAR — UP

5. COMPLETE MISSED APPROACH

PROCEDURE

MAP-LANDING ASSURED

THRESHOLD



3. POWER — IDLE

LANDING



PRIOR TO MAP, IF POSSIBLE

2. TORQUE — 1,100 - 1,300 LBS

3. 130 - 140 KIAS (VYSE MIN)

1. FLAPS — DOWN




2. BETA OR REVERSE


CAUTION CAUTION



1 NM

MDAMAP

ARRIVAL

1. PLAN CIRCLING MANEUVER

2. FOLLOW NORMAL APPROACH

PROCEDURES TO MDA

MINIMUM DESCENT ALTITUDE (MDA)

1. LEVEL OFF AT MDA AT LEAST 1

MILE PRIOR TO MAP, IF POSSIBLE

2. TORQUE — 1,100 - 1,300 LBS

3. 130 - 140 KIAS (VYSE MIN)

4. MANEUVER WITHIN VISIBILITY

CRITERIA

5. MAINTAIN MDA

MAP AND DURING CIRCLING MANEUVER

1. DETERMINE THAT VISUAL CONTACT WITH

THE RUNWAY ENVIRONMENT CAN BE

MAINTAINED AND A NORMAL LANDING CAN

BE MADE FROM A CIRCLING APPROACH,

OR INITIATE A MISSED APPROACH

2. MAINTAIN MDA DURING CIRCLING MANEUVER

BASE

1. COMMENCE DESCENT FROM

A POINT WHERE A NORMAL

1. 130 - 140 KIAS (VYSE MIN)


2. FLAPS — DOWN



FINAL

THRESHOLD



3. POWER — IDLE

NOTE: THIS IS A CATEGORY B AIRCRAFT, BUT

AIRSPEEDS OF 121 THROUGH 140 KIAS

REQUIRE USING CATEGORY C MINIMUMS.



CAUTION

TO ENSURE CONSTANT REVERSING CHARACTERISTICS

CAUTION

IF POSSIBLE PROPELLERS SHOULD BE MOVED OUT OF

A POINT WHERE A NORMAL

LANDING CAN BE MADE



LANDING

FLAPS-UP APPROACH ANDLANDING

Follow normal approach and landing proce-dures except:

• Complete the flaps up landing checklist.

• Refer to the flaps up VREF.

• Airspeed 140 knots until established onfinal.

• When landing assured—reduce the air-speed to the flaps up VREF.

SINGLE-ENGINE APPROACHAND LANDING


• Complete the one-engine-inoperativeapproach and landing checklist.

• The target torque settings are approxi-mately doubled.

• Smoothly push the propeller lever fullforward (2,000 rpm) prior to the IAF or downwind.

• Maintain the airspeed at least 10 knotsabove VREF until landing assured.

• Cautiously use reverse, if necessary.• If performance is limited when accom-

plishing a circling approach, circle with

CROSSWIND APPROACH ANDLANDING


• Crab into the wind to maintain the de-sired track across the ground.

• Immediately prior to touchdown, lower the upwind wing by use of the aileronand align the fuselage with the runwayby use of the rudder. During the rollout,hold the aileron control into the wind andmaintain directional control with therudder and brakes.

WINDSHEAR

GENERAL

The best windshear procedure is avoidance.Recognize the indications of potential wind-shear and then:

AVOID AVOID AVOID

The key to recovery from windshear is to flythe aircraft so it is capable of a climb gradientgreater than the windshear-induced loss of per-formance. Normally, the standard wind/gustcorrection factor 1/2 gust will provide a suf-ficient margin of climb performance. If a shear is encountered that jeopardizes safety, initiate

a rejected landing procedure. If the sink rateis arrested, continue with the procedure for microbursts.



p g g pp ,the flaps positioned for approach andthe gear up until it is certain the field canbe reached with the gear down.

MICROBURSTS

If a microburst is encountered, the first indi-cation will be a rapid increase in the rate of descent accompanied by a rapid drop belowglide path (visual or electronic).

1. Initiate normal rejected landing pro-cedures (10° pitch).

2. Do not change the aircraft configurationuntil a climb is established.

3. If the aircraft is not climbing, smoothlyincrease pitch until a climb is estab-lished or stall warning is encountered.If stall warning is encountered, de-crease pitch sufficiently to depart thestall warning regime.

4. When positively climbing at a safe al-titude, complete the rejected landingmaneuver.

NOTE

The positive rate of climb should beverified on at least two (2) instru-ments. Leave the gear down until

you have this climb indication, as itwill absorb some energy on impactshould the microburst exceed your capability to climb.

If a decision is made to rotate to the

stall warning, extreme care should beexercised so as not to over rotate be-yond that point as the aircraft is onlya small percentage above the stall

ACCEPTABLE PERFORMANCEGUIDELINES

• Understand that avoidance is primary.

• Ability to recognize potential windshear situations.

• Ability to fly the aircraft to obtain op-timum performance.

WARNING





a small percentage above the stallwhen the aural warning activates.

COCKPIT RESOURCE MANAGEMENT



REMEMBER

2 + 2 = 2

— OR —

2 + 2 = 5(SYNERGY)

IT'S UP TO YOU!

CLUES TO IDENTIFYING:• Loss of Situational Awareness

• Links in the Error Chain

H

U M A N

O P E R A T I O N A L

1. FAILURE TO MEET TARGETS2. UNDOCUMENTED PROCEDURE3. DEPARTURE FROM SOP4. VIOLATING MINIMUMS OR LIMITATIONS5. NO ONE FLYING AIRPLANE6. NO ONE LOOKING OUT WINDOW7. COMMUNICATIONS8. AMBIGUITY9. UNRESOLVED DISCREPANCIES

10. PREOCCUPATION OR DISTRACTION11. CONFUSION OR EMPTY FEELING12.

GROUPS/A

CAPTAININDIVIDUAL

S/A

COPILOTINDIVIDUAL

S/A

SITUATIONAL AWARENESS IN THE COCKPIT

LEADERSHIP STYLES

AUTOCRACTIC AUTHORITARIAN DEMOCRATICLAISSEZ-

FAIRE

COMMAND AND LEADERSHIP

Figure GEN-17. Situational Awareness in the Cockpit



STYLE

(EXTREME)

LEADERSHIP

STYLE

LEADERSHIP

STYLE

FAIRE

STYLE

(EXTREME)

PARTICIPATION



EVALUATE

RESULTRECOGNIZE

NEED

HINTS:• Identify the problem: • Communicate it • Achieve agreement

• Obtain commitment

DECISION MAKING PROCESS

FEEDBACK

NEED SEND RECEIVE

INTERNALBARRIERS

EXTERNALBARRIERS

INTERNALBARRIERS

OPERATIONALGOAL

THINK:• Solicit and give

feedback• Listen carefully• Focus on behavior,

not people

• Maintain focus onthe goal• Verify operational

outcome is achieved

ADVOCACY: To increase other's S/A

• State Position

• Suggest Solutions

• Be Persistent and Focused

• Listen Carefully

INQUIRY: To increase your own S/A

• Decide What, Whom, How to ask

• Ask Clear, Concise Questions

• Relate Concerns Accurately

• Draw Conclusions from Valid Information

• Keep an Open Mind

— REMEMBER —Questions enhance communication flow. Do not give in to the temptation to ask questionswhen advocacy is required. Use of advocacy or inquiry should raise a "red flag."

COMMUNICATION PROCESS

Figure GEN-19. Communication Process



RESULT

IMPLEMENT

RESPONSE

NEED

IDENTIFY

AND

DEFINE

PROBLEM

• Obtain commitment

• Consider appropriate SOPs

• Think beyond the obvious alternatives



Chapter 21. B or D

2. C3. A

4. C

5. C

6. B

7. C

8. A

9. C10. B

11. B

12. D

13. D

14. B15. D

16. B17. C

18. D

19. B

20. A

Chapter 31. B

2. A

3. D

4. C

5. A

6. B

7. C8. A

9. A

Chapter 41 C

Chapter 5 (cont.)6. A

7. C8. A

Chapter 71. B

2. A

3. C or D

4. B5. A

6. D

7. B

8. A

9. A10. D

Chapter 81. A

2. C

3. A

4. D

Chapter 91. C

2. B

3. D

4. D

5. A

6. B7. D

8. C

Chapter 101 A

Chapter 10 (cont.)13. B

14. C15. A

16. D

17. B

18. D

19. A

20. B

21. C22. B

23. B

24. C

25. C

Chapter 111. A

2. C

3. A

4. D

5. D

6. B

7. A8. A

9. A

10. B

11. C

12. A

Chapter 121. A

2. B3. B

4. D

5 A

Chapter 14 (cont.)9. C

10. C11. B

Chapter 151. B

2. B or C

3. C4. C

5. A

6. D

Chapter 16

1. A2. D3. A

4. D

5. C

6. A

7. B

8. B

9. C10. A

11. D

Chapter 171. B or D

2. C

3. C4. A

5. B6. C

7. D

8 A

ANSWERS TO QUESTIONS



1. C

2. D

3. B4. A

5 C

1. A

2. B

3. D4. A

5 C

5. A

6. A, B or D

Chapter 141 B

8. A

Chapter 181. B

2 D



WALKAROUND

The following section is a pictorial walkaround. It shows each itemcalled out in the exterior power-off preflight inspection.

The general location photographs do not specify every checklistitem. However, each item is portrayed on the large-scale photographsthat follow.





1. CABIN DOOR SEAL—CHECK

2. FLAPS—CHECK

4. LEFT MAIN GEAR, STRUT, TIRES, BRAKES—CHECK

5. CHOCK—REMOVE

6. BRAKE DEICE LINE (IF INSTALLED)—CHECK



7. FIRE EXTINGUISHER PRESSURE (IF INSTALLED)—CHECK

WALKAROUND INSPECTION





9. AILERON, AILERON TAB, STATIC WICKS (4)—CHECKED

10. FLUSH OUTBOARD DRAIN—DRAIN 13. STALL WARNING VANE—CHECK

12. MAIN FUEL TANK CAP—SECURE



11. NAVIGATION, RECOGNITION, STROBE LIGHT—CHECKED

14. TIEDOWN—REMOVED

16. STALL STRIP—CHECK 19. WING LEADING EDGE TANK SUMP—DRAIN

15. OUTBOARD DEICE BOOTS—CHECKED 18. RAM SCOOP FUEL VENT AND HEATED FUEL VENT—CLEAR





17. ICE LIGHT—CHECK 20. GRAVITY LINE DRAIN—DRAIN



21. FUEL FILTER STRAINER DRAIN AND STANDBY PUMPDRAIN—DRAIN

22. LANDING GEAR DOORS—CHECK 25. ENGINE OIL CAP—SECURE

24. ENGINE OIL—CHECK



23 WHEEL WELL CHECK 26 ENGINE COMPARTMENT DOOR (OUTBOARD)

29. NACELLE COOLING RAM AIR INLETS—CLEAR

27. EXHAUST STACK (OUTBOARD)—CHECK FOR CRACKS

28. TOP COWLING LOCKS (OUTBOARD)—SECURE

31. ENGINE AIR INTAKE—CLEAR



32. ENGINE COMPARTMENT DOOR (INBOARD)—SECURE,BLEED VALVE EXHAUST CLEAR

33. TOP COWLING LOCKS (INBOARD)—SECURE

34. EXHAUST STACK (INBOARD)—CHECK FOR CRACKS





36. AUXILIARY FUEL TANK CAP—SECURE

37. HYDRAULIC GEAR SERVICE DOOR—SECURE 40. HYDRAULIC LANDING GEAR VENT LINES—CLEAR

39. INBOARD DEICE BOOTS—CHECKED



38. HEAT EXCHANGER (INLET & OUTLET)—CLEAR 41. AUXILIARY FUEL TANK SUMP—DRAIN

42. LOWER ANTENNAS AND BEACON—CHECKED



43. OAT PROBE—CHECK

44. AVIONICS PANEL—SECURE

45. CONDENSER BLOWER OUTLET—CLEAR

48. NOSE GEAR STEERING STOP BLOCK—CHECK

49. NOSE GEAR WHEEL WELL—CHECK





51. PITOT MASTS—COVER REMOVED, CLEAR

52. WINDSHIELD, WINDSHIELD WIPERS—CHECK

53. RAM AIR INLET—CLEAR

54. RADOME—CHECK

63. TOP COWLING LOCKS (INBOARD)—SECURE

64. EXHAUST STACK (INBOARD)—CHECK FOR CRACKS

65. NACELLE COOLING RAM AIR INLETS—CLEAR

For Right Wing, A Close Up Is Given For Those Components

Different Than Left Wing (See Foldout Page For Specific Locations)



55. AVIONICS PANEL—SECURE

56. AUXILIARY FUEL TANK CAP—SECURE

57. INBOARD DEICE BOOTS—CHECKED

66. PROPELLER—CHECK FOR NICKS, DEICE BOOTSECURE

67. ENGINE AIR INTAKE—CLEAR



69. BATTERY BOX DRAIN—CLEAR

71. ENGINE COMPARTMENT DOOR (OUTBOARD)—SECURE

72. EXHAUST STACK (OUTBOARD)—CHECK FOR CRACKS

73. TOP COWLING LOCKS (OUTBOARD)—SECURE

74. GENERATOR COOLING INLET—CLEAR

75. FUEL FILTER STRAINER DRAIN AND STANDBY PUMPDRAIN—DRAIN

76. LANDING GEAR, DOORS, STRUT, TIRES, BRAKES—CHECKED

77. CHOCK—REMOVE

78. FIRE EXTINGUISHER PRESSURE (IF INSTALLED)—CHECK

68. BATTERY AIR EXHAUST (NICKEL-CADMIUM)—CLEAR

See Foldout Page ForSpecific Locations



70. BATTERY AIR INLET (NICKEL-CADMIUM)—CLEAR,VALVE FREE

79. EXTERNAL POWER DOOR—SECURE



80. RAM SCOOP FUEL VENT AND HEATED FUEL VENT—CLEAR

81. GRAVITY LINE DRAIN—DRAIN

82. INVERTER COOLING LOUVERS—CLEAR

83. WING LEADING EDGE TANK SUMP—DRAIN

84. ICE LIGHT—CHECK

85. OUTBOARD DEICE BOOTS—CHECKED86. TIEDOWN—REMOVE

87. FLUSH OUTBOARD DRAIN—DRAIN

88. MAIN FUEL TANK CAP—SECURE

89. NAVIGATION, RECOGNITION, STROBE LIGHT—CHECKED

91. AILERON, FLAPS—CHECKED

92. BRAKES—CHECK

93. BRAKE DEICE (IF INSTALLED)—CHECK

94. OIL BREATHER VENT—CLEAR

See Foldout Page For Specific Locations





96. LOWER ANTENNAS—CHECKED 100. CABIN AIR EXHAUST—CLEAR

97. VENTRAL FIN DRAIN HOLES—CLEAR

98. TIEDOWN—REMOVE

101. OXYGEN SERVICE ACCESS DOOR—SECURE

102. RIGHT STATIC PORTS—CLEAR





105. VENTRAL FIN, STATIC WICK (1)—CHECKED

104. ACCESS PANEL—SECURE

108. ELEVATOR, ELEVATOR TAB, STATIC WICKS (3 EACHSIDE)—CHECKED

109. POSITION LIGHT—CHECK

110. TAIL FLOODLIGHTS (LEFT AND RIGHT IFINSTALLED)—CHECKED

107. RUDDER TAB—CHECK





112. ACCESS PANEL—SECURE

113. RELIEF TUBE DRAIN—CLEAR

116. LEFT STATIC PORTS—CLEAR



114. AFT COMPARTMENT DRAIN TUBE—CLEAR



91959192939496999897100105

107

106

108 111

110104 103 101 1 02

747578817776838279808687

89

88 85 90 7184 72 73



ANN-1FOR TRAINING PURPOSES ONLY



ANNUNCIATOR PANELS

The Annunciator section presents a color representation of all theannunciator lights in the plane.







Documents

B200 Pilot Training Manual