200/200c poh 101-590010-127a13 - beechcraft · rudder boost 7-6 instrument panel 7-6 typical...

BEECHCRAFTSuper King Air 200

SECTION VII

SYSTEMS DESCRIPTIONS

TABLE OF CONTENTS

SUBJECT PAGE

AirframeStructure 7-5Seating Arrangements 7-5

Flight ControlsControl Surfaces 7-5Operating Mechanisms 7-5ManualElevator Trim 7-5Electric Elevator Trim 7-5Rudder Boost 7-6

Instrument Panel 7-6Typical Illustrations (Prior to 1978 Model Year)

Pilot's Control Wheel 7-8Overhead Light Control Panel. 7-8Copilot's Control Wheel 7-8Instrument Panel 7-8Fuel Control Panel 7-9Pedestal : 7-9Right SidePanel 7-9

Typical Illustrations (1978 Model Year and After)Pilot's Control Wheet 7-10Overhead Light Control Panel 7~10Copilot's Control Wheel 7-10Instrument Panel '" 7-10FuelControl Panel 7-11Pedestal 7-11RightSidePanel 7-11

Annunciator System 7-13Annunciator Panels 7-14

Ground ContrOl : 7-17Flaps 7-17Landing Gear 7·17

Landing GearWarning System 7-17Serials Prior to 88-324 7-18Serials88-324 thru88-452 7-18Serials B8·453 andAfter, BL·1 andAfter 7-18

ManualLanding Gear Extension 7-18BrakeSystem 7-18Tires 7-19

Baggage Compartment. , 7-19

February, 1979

Section VIISystems Descriptions

SUBJECT


PAGE

7-2

Seats, Seatbelts, and Shoulder HarnessesSeats

Cockpit 7-19Cabin 7-19Foyer 7~20

Aft-Cabin Area 7-20Seatbelts 7-20Shoulder Harnesses

Cockpit 7-20Cabin 7-20Aft-Cabin Area 7-20

Doors, Windows, and ExitsAirstair Entrance Door (200) 7-21Airstair Entrance Door (200C) 7-21Cargo Door (200C) 7-22

Emergency Exit 7-22Interior Doors 7-22Cabin Windows 7-22

Polarized Type 7-22Shade Type 7-22

Control Locks 7-23Engines 7-25

Cutaway View 7-25Propulsion System Controls 7-26

Power Levers 7-26Propeller Levers '" 7-26Condition Levers '" 7-26Propeller Reversing 7-26Friction Locks 7-26

Engine Instrumentation 7-26Propeller Synchrophaser 7-27Engine Lubrication System 7-27

Magnetic Chip Detector 7 -27Starting and Ignition System 7~27

Auto Ignition 7-27Induction Air System 7-27

Ice ProtectionEngine Air Inlet 7-28Ice Vanes (tnertial Separator System) 7-28

Fuel Control 7-28Fire Detection System 7-29

Schematic 7-29Fire Extinguisher System 7-30

Schematic 7-30Propeller System

Description 7-31Primary Low Pitch Stop 7-31Propeller Governors 7-31Autofeather System 7-31

Fuel System 7-31Schematic 7-32Fuel Pumps 7-33Auxiliary Fuel Transfer System 7-33Use of Aviation Gasoline 7-34Crossfeed 7-34Firewatl Shutoff 7-34Fuel Routing in Engine Compartment. 7-34Fuel Drains 7-34

February. 1979


8ectionYIISystems Description

FuelDrainCollector System 7-34FuelGaging System 7-34

Electrical System 7-35PowerDistribution Schematic 7-36External Power 7-37

Lighting SystemsCockpit 7-37Cabin 7-37Exterior 7-37

Environmental System 7-37Schematic 7-38Pressurization System 7-39FlowControl Unit. 7-40Unpressurized Ventilation 7-41Heating 7-41Radiant Heating 7-42AirConditioning System 7-42Environmental Controls 7-42

Heating Mode 7-43Cooling Mode 7-43Automatic ModeControl 7-43ManualMode Control 7-43BleedAirControl 7-44

Oxygen System 7-44Manual Plug-in System 7-44Auto-deployment System 7-44

Pitot andStatic System 7-45Schematic 7-47

Engine Bleed Air Pneumatic system 7-46BleedAirWarning System 7-46

Automatic Devices in the Control SystemYaw Damp 7-46

Stall Warning system 7-46Ice Protection Systems

Windshield Heat 7-46Pitot andStatic System Schematic 7-47

Propeller Electric DeiceSystem 7-48Schematic 7-48

Surface Deice System 7-49Schematic 7-50

Pitot Mast 7-49Stall Warning Vane 7-49Fuel 7-51

Comfort FeaturesToilet 7-51Relief Tubes 7-51

Cabin FeaturesFireExtinguishers 7-51

Windshield Wipers 7-51CargoRestraint (200C) 7-51

February, 1981 7-3


7·4

INTENTIONALLY LEFT BLANK


October, 1978


AIRFRAME

STRUCTURE

The Cilecchuaft Super King Air 200 is an all-metal, low-wingmonoplane. It utilizes fully cantilevered wings, and aT-tailempennage.

SEATING ARRANGEMENTS

The pilot and copilot seats are mounted in a separate forwardcompartment. Various configurations of passenger chairsand two- or four-place couch installations may be installed onthe continuous tracks mounted on the cabin floor. One or twofold-up seats may be installed in the aft cabin area. The toiletis also equipped for use as a seat. Seating for up to 15persons, including crew, is available. For additionalinformation, refer to the "Cabin Arrangement Diagram" inSection VI, WEIGHT AND BALANCE/EQUIPMENT LIST.

FLIGHT CONTROLS

CONTROL SURFACES

The airplane is equipped with conventional ailerons andrudder. It utilizes aT-tail horizontal stabilizer and elevator,mounted at the extreme top of the vertical stabilizer.

OPERATING MECHANISMS

The airplane is equipped with conventional dual controls forthe pilot and copilot. The ailerons and elevators are operatedby conventional control wheels interconnected by aT-bar.The rudder pedals are interconnected by linkage below thefloor. These systems are connected to the control surfacesthrough push-rod and cable-and-bellcrank systems. Rudder,elevator, and aileron trim are adjustable with controlsmounted on the center pedestal. A position indicator for eachof the trim tabs is integrated with its respective control.

MANUAL ELEVATOR TRIM

Manual control of the elevator trim is accomplished with ahandwheel located on the left side of the pedestal. It is aconventional trim wheel which is rolled forward for nosedown trim, and aft for nose-up trim.

ELECTRIC ELEVATOR TRIM

The electric elevator-trim system, if installed, is controlled byan ELEV TAB CONTROL - ON - OFF switch located on thepedestal, a dual-element thumb switch on each controlwheel, a trim-disconnect switch on each control wheel, and aPITCH TRIM circuit breaker in the FLIGHT group on the rightside panel. The ELEV TAB CONTROL switch must be ON for

October, 1978


the system to operate. Both elements of either dual-elementthumb switch must be simultaneously moved forward toachieve nose-down trim, aft for nose-up trim; when released.they return to the center (off) position. Any activation of thetrim system by the copilot's thumb switch can be overriddenby the pilot's thumb switch. A before take-off check of bothdual-element thumb switches should be made by movingeach of the four switch elements individually. No one switchelement should activate the system; moving the two switchelements on either the pilot's or the copilot's control wheel inopposite directions should not activate the system - only thesimultaneous movement of a pair of switch elements in thesame direction should activate the electric elevator-trimsystem.

A bi-Ievel, push-button, momentary-on, trim-disconnectswitch is located inboard of the dual-element thumb switchon the outboard grip of each control wheel. The electricelevator-trim system can be disconnected by depressingeither of these switches. If an autopilot is installed,depressing either trim-disconnect switch to the first of the twolevels disconnects the autopilot and the yaw damp system:depressing the switch to the second level disconnects theautopilot, the yaw damp system, and the electric elevatortrim system. If an autopilot is not installed, depressing theswitch to the first level does not do anything, since the yawdamp system is controlled by a separate YAW DAMP switchon the pedestal; depressing the switch "to the second leveldisconnects the electric elevator-trim system. A greenannunciator on the caution advisory annunciator panel,placarded ELEC TRIM OFF, alerts the pilot whenever thesystem has been disabled with a trim-disconnect switch andthe ELEV TAB CONTROL SWitch is ON. The system can bereset by cycling the ELEV TAS CONTROL switch on thepedestal from ON to OFF, then to ON again. The manual-trimcontrol wheel can be used to change the trim anytime,whether or not the electric-trim system is in the operativemode.

RUDDER BOOST

A rudder boost system is provided to aid the pilot inmaintaining directional control in the event of an enginefailure or a large variation of power between the engines.Incorporated into the rudder cable system are two pneumaticrudder-boosting servos that actuate the cables to providerudder pressure to help compensate for asymmetrical thrust.

During operation, a differential pressure valve accepts bleedair pressure from each engine. When the pressure variesbetween the bleed air systems, the shuttle in the differentialpressure valve moves toward the low pressure side. As thepressure difference reaches a preset tolerance, a switch onthe low pressure side closes, activating the rudder boostsystem. This system is designed only to help compensate forasymmetrical thrust. Appropriate trimming is to beaccomplished by the pilot. Moving either or both of the bleedair valve switches on the copilot's subpanel to the INSTR &ENVIR OFF position will disengage the rudder boost system.

7-5


The system is controlled by a toggle switch, placardedRUDDER BOOST - ON - OFF, located on the pedestal belowthe rudder trim wheel. The switch is to be turned ON beforeflight. A preflight check of the system can be performedduring the run-up by retarding the power on one engine toidle and advancing power on the opposite engine until thepower difference between the engines is great enough toclose the switch that activates the rudder boost system.Movement of the appropriate rudder pedal (left engine idling,right rudder pedal moves forward) will be noted when theswitch closes, indicating the system is functioning properlyfor low engine power on that side. Repeat the check withopposite power settings to check for movement of theopposite rudder pedal.

7·6


INSTRUMENT PANEL

The floating instrument panel design allows the flightinstruments to be arranged in a group directly in front of thepilot and the copilot. Complete pilot and copilot flightinstrumentation is installed, including dual naviqationsystems, two course indicators, dual gyro horizons, and dualturn and slip indicators.

The operation and use of the instruments, lights, switches,and controls located on the instrument panel is explainedunder the systems descriptions relating to the subject items.

October. 1978

BEECHCAAFTSuper King Air 200

October, 1978



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Systems Descriptions

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7·12



October, 1978


ANNUNCIATOR SYSTEM

The annunciator system consists of a warning annunciatorpanel (with red readout) centrally located in the glareshield,a caution/advisory annunciator panel (caution - yellow;advisory - green) located on the center subpanel, two yellowMASTER CAUTION flashers (located just inboard of theMASTER WARN ING flashers), a CAUTION· ON - OFFswitch located on the copilot's left subpanel, and a PRESSTO TEST switch located immediately to the right of thewarning annunciator panel.

The annunciators are of the word-readout type. Whenever afault condition covered by the annunciator system occurs, asignal is generated and the appropriate annunciator is illum-inated. .

If the fault requires the immediate attention and reactionof the pilot, the appropriate red warning annunciator in thewarning annunciator panel illuminates and both MASTERWARNING flashers begin flashing. Any illuminated lens inthe warning annunciator panel will remain on until the faultis corrected. However, the MASTER WARNING flashers canbe extinguished by depressing the face of either MASTERWARNING flasher, even if the fault is not corrected. Insuch a case, the MASTER WARNING flashers will again beactivated if an additional warning annunciator illuminates.When a warning fault is corrected, the affected warningannunciator will extinguish, but the MASTER WARNINGflashers will continue flashing until one of them isdepressed.

Whenever an annunciator-covered fault occurs that requiresthe pilot's attention but not his immediate reaction, theappropriate yellow caution annunciator in the caution/advisory panel illuminates, and both MASTER CAUTIONflashers begin flashing. The flashing MASTER CAUTIONlights can be extinguished by pressing the face of either ofthe flashing lights to reset the circuit. Subsequently, whenany caution annunciator illuminates, the MASTE RCAUTION flashers will be activated again. Normally, anilluminated caution annunciator on the caution/advisoryannunciator panel will remain on until the fault conditionis corrected, at which time it will extinguish. The MASTE RCAUTION flashers will continue flashing until one ofthem is depressed.

If the fault indicated by an illuminated caution annunciatoris not corrected, the pilot can still extinguish the annunciatorby momentarily moving the spring-loaded CAUTIONtoggle switch down to the OFF position, then releasing it

October, 1979


to the center position. This action will extinguish all illuminated caution annunciators, and will illuminate the greenCAUT LGND OFF advisory annunciator in the caution/advisory annunciator panel, to remind the pilot that anuncorrected fault condition exists, but that the cautionlegends have all been extinguished. The annunciatorts)previously extinguished with the CAUTION switch can bere-illuminated anytime by momentarily moving the switchup to the ON position. This action will also extinguish thegreen CAUT LGND OFF annunciator. If an additionalfault covered by the caution annunciators occurs afterthe caution legends have been extinguished with theCAUTION switch, the appropriate caution annunciator forthe new fault will illuminate, and all previously extinguishedannunciators will re-illurninate.

The caution/advisory annunciator panel also contains thegreen advisory annunciators. There are no master flashersassociated with these annunciators, since they are onlyadvisory in nature, indicating functional situations whichdo not demand the immediate attention or reaction of thepilot. An advisory annunciator can be extinguished only bycorrecting the condition indicated on the illuminated lens.

The warning annunciators, caution annunciators, advisoryannunciators and yellow MASTE R CAUTION flashersfeature both a "bright" and a "dim" mode of illuminationintensity. The "dim" mode will be selected automaticallywhenever all of the following conditions are met: ageneratoris on the line; the OVERHEAD FLOOD LIGHTS are OFF;the PILOT FLIGHT LIGHTS are ON; and the ambientlight level in the cockpit (as sensed by a photoelectric celllocated in the overhead light control panel) is below a presetvalue. Unless all these conditions are met, the "bright"mode will be selected automatically. The MASTER WARNING flasher does not have a "dim" mode.

The lamps in the annunciator system should be testedbefore every fl ight, and anytime the integrity of a lamp isin question. Depressing the PRESS TO TEST button,located to the right of the warning annunciator panel inthe glareshield, illuminates all the annunciator lights,MASTER WARNING flashers, and MASTER CAUTIONflashers. Any lamp that fails to illuminate when testedshould be replaced (refer to LAMP REPLACEMENTGUIDE in Section VIII, HANDLING, SERVICING ANDMAINTENANCE).

7-13



ANNUNCIATOR PANELS

NOMENCLATURE COLOR CAUSE FOR ILLUMINATION

FIRE L ENG

ALTWARN

FIRER ENG

L FUEL PRESS

INST INV

RFUEL PRESS

L BL AIR FAIL

R BL AIR FAIL

L CHIP DETECT

R CHIP DETECT

Red

Red

Red

Red

Red

Red

Red

Red

Red

Red

WARNING ANNUNCIA TOR

Fire in left engine compartment

Cabin altitude exceeds 12,500 feet

Fire in right engine compartment

Fuel pressure teilure on left side

The inverter selected is inoperative

Fuel pressure failure on right side

Melted or failed plastic left bleed air failurewarning line

Melted or failed plastic right bleed air failurewarning line

Contamination in left engine oil is detected

Contamination in right engine oil is detected

CAUTION/ADVISORY ANNUNGIA TOR

L DCGEN

L ICE VANE

RVS NOT READY

R ICE VANE

ROGGEN

IGABINDOOR

PROP SYNC ON

EXTPWR

BATTERYCHG

DUCT OVERTEMP

Yellow

Yellow

Yellow

Yellow

Yellow

Yellow

Yellow

Yellow

Yellow

Yellow

Left generator off the line

Left ice vane malfunction. Ice vane has not attainedproper position.

Propeller levers are not in the high rpm, lowpitch position with landing gear extended

Right ice vane malfunction. Ice vane has not attainedproper position.

Right generator off the line

Cabin/cargo door open or not secure

Synchrophaser turned on with landing gear extended

External power connector is plugged in

Excessive charge rate on battery

Duct air too hot

February, 1979


L AUTOFEATHER

ELEC TRIM OFF

Green

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Aototeetner armed with power levers advanced above90% Nt

Electric trim ae-enerqized by a trim disconnect switchon the control wheel with the system power switch onthe pedestal turned on.

FUEL CROSSFEED Green

AIR COND N 1 LOW Green

R AUTOFEATHER Green

L ICE VANE EXT Green

BRAKE DEICE ON Green

LANDING LIGHT Green

PASS OXYGEN ON Green

R ICE VANE EXT Green

L IGNITION ON Green

Crossfeed valve is open

Right engine rpm too low for air condWoning load

Autoteetber armed with power levers advanced above90% Nt

Ice vane extended

Brake deice system in operation

Landing lights on with landing gear up

Passenger oxygen system charged

Ice vane extended

Left starter/ignition switch is in the engine/ignitionmode or left auto ignition system is armed and leftengine torque is below 400 It Ibs

L BL AIR OFF

GAUT LGND OFF

R BL AIR OFF

R IGNITION ON

October, 1978

Green

Green

Green

Green

Left environmental bleed air valve closed

Caution annunciator is turned off

Right environmental bleed air valve is closed

Right starter/ignition switch is in the engine/ignitionmode or right auto ignition system is armed and rightengine torque is below 400 ft Ibs

7-15




October. 1978

BEECHCRAFTSuper Kin9 Air 200

GROUND CONTROL

Direct linkage from the rudder pedals allows for nose wheelsteering. When the rudder control is augmented by a mainwheel brake, the nose wheel deflection can be considerablyincreased.

The minimum wing-tip turning radius, using partial brakingaction and differential engine power, is 39 feet 10 inches.

FLAPS

Two flaps are installed on each wing. Power is delivered froman electric motor to a gearbox mounted on the forward side ofthe rear spar. The gearbox drives four flexible driveshaftswhich are connected to jackscrews, one of which operateseach flap. The motor incorporates a dynamic braking system,through the use of two sets of motor windings. This featurehelps prevent overtravel of the flaps. A safety mechanism isprovided to disconnect power to the electric flap motor in theevent of a malfunction which would cause any flap to be threeto six degrees out of phase with the other flaps.

The flaps are operated by a sliding switch handle on thepedestal just below the condition levers. Flap travel, from 0%(full up) to 100% (full down) is registered on an electricindicator on top of the pedestal. A side detent provides forquick selection of the APPROACH position (40% flaps).From the UP position to the APPROACH position, the flapscannot be stopped in an intermediate position. BetweenAPPROACH and DOWN, the flaps can be stopped anywhereby moving the handle to the DOWN position until the flapsreach the desired position, then moving the flap-switchhandle back to APPROACH. The flaps can be raised to anyposition between DOWN and APPROACH by raising thehandle to UP until the desired setting is reached, thenreturning the handle to APPROACH. Selecting theAPPROACH position will stop flap travel anytime the flapsare deflected more than 40%.

The flap-motor power circuit is protected by a 20-ampereflap-motor circuit breaker placarded FLAP MOTOR, locatedon the left circuit breaker panel below the fuel control panel.A 5-ampere circuit breaker for the control circuit (placardedFLAP CONTROL) is also located on this panel.

Lowering the flaps will produce these results:

Attitude - Nose UpAirspeed - ReducedStall Speed - LoweredTrim - Nose-Down Adjustment Required

to Maintain Attitude

LANDING GEAR

A 28-volt split-field motor, located on the forward side of thecenter-section main spar, extends and retracts the landing

October, 1978


gear. The landing gear motor is controlled by the switchplacarded LOG GEAR CaNT - UP - ON on the pilot's rightsubpanel. The switch handle must be pulled out of a detentbefore it can be moved from either the UP or the ON position.The motor incorporates a dynamic braking system, throughthe use of two motor windings, which helps preventovertravel of the gear-

Torque shafts drive the main gear actuators, and duplexchains drive the nose gear actuator. A spring-loaded frictiontype overload clutch incorporated into the gearbox preventsdamage to the structure and to the torque shafts in the eventof mechanical malfunction. An overload protection circuit,located on the landing gear panel forward of the main sparunder the center floorboard, protects the system fromelectrical overload.

The Beech air-oil type shock struts are filled with compressedair and hydraulic fluid. Spring-loaded linkage from the rudderpedals permits nose wheel steering. When the rudder controlis augmented by a main wheel brake, the nose wheeldeflection can be considerably increased. As the nose wheelretracts after lift-off. it is automatically centered and thesteering linkage becomes inoperative.

A safety switch on the right main gear torque knee opens thecontrol circuit when the strut is. compressed. The safetyswitch also actuates a solenoid-operated down-lock hook onthe landing gear control switch located on the pilot's rightsubpanel. This mechanism prevents the landing gear handlefrom being raised when the airplane is on the ground. Thehook automatically unlocks when the airplane leaves theground and can be manually overridden by pressing down onthe red down-lock release button just left of the landing gearcontrol switch handle, in the event of a malfunction of thedown-lock solenoid. The landing gear control switch handleshould never be moved out of the ON detent while theairplane is on the ground; if it is, the landing gear warninghorn will sound intermittently and the red gear-in-transit lightsin the landing gear control switch handle will illuminate(provided the MASTER SWITCH is ON), warning the pilot toreturn the handle to the DN position.

Visual indication of landing gear position is provided byindividual green GEAR DOWN indicator lights arranged in atriangle on the pilot's right subpanel. Two red, parallel-Wiredindicator lights located in the control handle illuminate toshow that the gear is in transit or not locked. They alsoilluminate Nhen the landing gear warning horn is actuated.

LANDING GEAR WARNING SYSTEM

The landing gear warning system is provided to warn the pilotthat the landing gear is not down and locked during specificflight regimes. Various warning modes result, dependingupon the position of the flaps.

7-17


SERIALS PRIOR TO 88-324

With the FLAPS UP and either or both power levers retardedbelow a certain power level, the warning horn will soundintermittently and the landing gear switch handle lights willilluminate. The horn can be silenced by pressing the WARNing HORN SILENCE button adjacent to the landing gearswitch handle; the lights in the landing gear switch handlecannot be cancelled. The landing gear warning system willbe rearmed if the power lever( s) are advanced sufficiently.

With the FLAPS in APPROACH position and either or bothpower levers retarded below a certain power level, thewarning horn and landing gear switch handle lights will beactivated and neither can be cancelled.

With the FLAPS BEYOND APPROACH position, the hornand landing gear switch handle lights will be activatedregardless of the power settings, and neither can becancelled.

SERIALS 88-324 THRU 88-452

With the FLAPS UP and either or both power levers retardedbelow a certain power level, the landing gear switch handlelights will illuminate; and if the airspeed is below 140 knots,the warning horn will sound intermittently. The horn can besilenced by pressing the WARNing HORN SILENCE buttonadjacent to the landing gear switch handle; the lights in thelanding gear switch handle cannot be cancelled. The landinggear warning system will be rearmed if the power lever(s) areadvanced sufficiently.

With the FLAPS in APPROACH or BEYOND APPROACHposition, the landing gear switch handle lights will illuminate;and if the airspeed is below 140 knots, the warning horn willsound intermittently. Neither the horn nor the lights can becancelled.

ISERIALS 88-453 AND AFTER, 8L-1 AND AFTER

With the FLAPS in UP or APPROACH position and either orboth power levers retarded below a certain power level, thewarning horn will sound intermittently and the landing gearswitch handle lights will illuminate. The horn can be silencedby pressing the WARNing HORN SILENCE button adjacentto the landing gear switch handle; the lights in the landinggear switch handle cannot be cancelled. The landing gearwarning system will be rearmed if the power lever(s) areadvanced sufficiently.

With the FLAPS BEYOND APPROACH position, the warninghorn and landing gear switch handle lights will be activatedregardless of the power settings, and neither can be cancelled.

MANUAL LANDING GEAR EXTENSION

Manual landing gear extension is provided through aseparate, manually powered, chain-drive system. Pull the

7-18


LANDING GEAR RELAY circuit breaker (on the pilot's rightsubpanel) and make certain that the landing gear switchhandle is in the down position before manually extending thegear. Pulling up on the emergency engage handle (locatedon the floor) and turning it clockwise will lock it in thatposition. When the emergency engage handle is pulled up,the motor is electrically disconnected from the system, andthe emergency drive system is locked to the gear box andmotor. When the emergency drive is locked in, the chain isdriven by a continuous-action ratchet, which is activated bypumping the ratchet handle adjacent to the emergencyengage handle.

CAUTION

Do not continue pumping the ratchet handle afterthe GEAR DOWN lights illuminate. Excessivepumping may damage the gear drive mechanismand bind the clutch so that the handle will notrelease it.

WARNING

After an emergency landing gear extension hasbeen made, do not stow pump handle or moveany landing gear controls or reset any switchesor circuit breakers until the airplane is on jacks.These precautions are necessary becau;;e thefailure may have been in the gear-up circuit, inwhich case the gear might retract on the ground.The gear can not be retracted manually.

After a practice manual extension, the landing gear may beretracted electrically. Rota'e the emergency engage handlecounterclockwise and push it down. Stow the extensionlever, push in the landing gear relay circuit breaker on thepilot's subpanel, and retract the gear in the normal mannerwith the landing gear switch handle.

BRAKE SYSTEM

The dual hydraulic brakes are operated by depressing thetoe portion of either the pilot's or copilot's rudder pedals.Shuttle valves permit braking by either pilot or copilot.

Dual parking-brake valves are installed adjacent to therudder pedals between the master cylinders of the pilot'srudder pedals and the wheel brakes. A control for the valves,placarded PARKING BRAKE, is located below the pilot's leftsubpanel. After the pilot's brake pedals have been depressedto build up pressure in the brake lines, both valves can beclosed simultaneously by pulling out the parking brakehandle. This retains the pressure in the brake lines. Theparking brake is released by depressing the pedals briefly toequalize the pressure on both sides of the valve, thenpushing in the parking brake handle to open the valve.

February, 1979


TIRES

The airplane is normally equipped with dual 18x5.5, 8-plyrated, tubeless, rim-inflation tires on each main gear. Forincreased service life, 1o-piy-rateo tires of the same size maybe installed.

Optionally, the airplane may be equipped with dual 22x6.7510, 8-ply-rated, tubeless tires on each main gear. These tirespro v ide higher flotation, and permit ope ration atapproximately 2/3 the inflation pressure required for thestandard 18x5.5 tires.

The nose gear is equipped with a single 22x6.75-10, 8-plyrated, tubeless tire.

NOTE

Prior fa serial 88-165, airplanes equipped with18x5.5 main-gear tires were delivered with a6.50x10, 6-ply-rated, tubeless tire installed onthe nose gear. These earlier nose-gear tires maybe replaced with 22x6. 75-10, a-ply-rated tires, ifdesired.

BAGGAGE COMPARTMENT

The entire aft-cabin area, which is aft of the foyer, may beutilized as a baggage compartment. A nylon web is providedfor the restraining of loose items.

CAUTION

Baggage and other objects should be secured bywebs, in order to prevent shifting in turbulent air.

Items stowed in the aft-cabin area are easily accessible inflight. The aft-cabin area can be closed off from the foyer bypulling the optional baggage compartment curtain across theopening and securing it with the snap fasteners provided.Alternately, a latching compartment door may be installed.The door is unlatched by rotating the latch handle clockwise,and latched by rotating the handle counterclockwise.

SEATS, SEATBELTS, AND SHOULDERHARNESSES

SEATS

COCKPIT

The pilot and copilot seats are adjustable fore and aft, as wellas vertically. When the release lever under the front inboardcorner of the seat is lifted, the seat can be moved forward oraft as required. When the release lever under the frontoutboard corner of the seat is lifted and no weight is on the

October, 1978


seat, the seat will rise in half-inch increments to its highestposition. When weight is on the seat and the lever is lifted,the seat will slowly move downward in half-inch incrementsuntil the lever is released, or until the seat reaches its lowestpoint of vertical travel. The armrests pivot at the aft end andcan be raised to facilitate entry to and egress from the seats.

CABIN

Various configurations of passenger chairs and 2- or 4-placecouches may be installed on the continuous tracks which aremounted on the cabin floor. All passenger chairs areplacarded either FRONT FACING ONLY or FRONT OR AFTFACING on the horizontal leg cross brace. Only chairsplacarded FRONT OR AFT FACING may be installed facingaft. All aft-facing chairs (and all forward-facing chairs that areequipped with shoulder harnesses) are equipped withadjustable headrests.

WARNING

Before takeoff and landing, the headrest shouldbe adjusted as required to provide support forthe head and neck when the passenger leansagainst the seatback.

Some passenger chairs can be moved fore and aft, to suitlegroom requirements of different passengers, by lifting ahorizontal release lever that extends laterally under the frontof adjustable seats. ("Front" is the direction opposite theseatback, regardless of whether the chair faces fore or att.)

The seatbacks can be adjusted to any angle from fully uprightto fully reclining, by depressing the release lever located onthe side of the seat at the front inboard corner. When thelever is depressed and the passenger leans against theseatback, the seatback will slowly recline until the lever isreleased, or until the fully reclining position is attained. Whenno weight is placed against the seatback and the lever isdepressed, the seatback will rise until the lever is released, oruntil the fully upright position is reached. The seatbacks of alloccupied seats must be upright for takeoff and landing.

The passenger-chair seatback can also be folded flat overthe seat cushion, after releasing the lock lever located on theside of the seat at the back inboard corner.

The optional lateral-tracking passenger chairs incorporate aflat, rectangular release lever underneath the front inboardcorner of the seats. When this lever is lifted, the chairs can beadjusted fore and aft, as well as laterally. The seatbackadjustments are the same as those on the standardpassenger chairs. When occupied, these seats must be inthe outboard position (i.e., against the cabin wall) for takeoffand landing.

Inboard armrests on passenger chairs - and both armrests oncouches and lateral-tracking chairs - can be folded flush with

7·19


the top of the seat cushions to facilitate entry to and egressfrom the seat. The armrests can be lowered by lifting the flat,rectangular release plate located under the front end of thearmrest. then moving the armrest toward the front of the seatand downward. The armrest can be raised by pulling thearmrest upward and toward the seatback until it locks intoplace.

The couches are not adjustable.

FOYER

Hinged seat-cushion halves mounted on top of the toilet forman extra passenger seat when the toilet is not in use.

AFT-CABIN AREA

One or two optional folding seats may be installed in theaft-cabin area. They are mounted on the cabin sidewall andswing inboard when unfolded. A latch mechanism on the leglocks the seats in place when they are unfolded. When thisseating is not needed, the seat(s) may be folded against thecabin sidewall and held in place with retaining straps.

SEATBELTS

Every seat in the airplane is equipped with a seatbelt. Theseatbelt can be lengthened by turning the male half of thebuckle at a right angle to the belt, then pulling the male half inthe direction away from the anchored end of the belt. Thebuckle is locked by sliding the male half into the female half ofthe buckle. The belt is then tightened by pulling the short endof the belt through the male half of the buckle until a snug fit isobtained. The buckle is released by lifting the large. hingedrelease lever on the female buckle half and pulling the malehalf of the buckle free. All occupants must wear seat beltsduring takeoff and landing.

SHOULDER HARNESSES

COCKPIT

The shoulder harness installations for the pilot and copnotseats consist of two straps each. Each strap is routed fromthe lower aft area of the seat, up the seatback, and through aretaining loop on top of the seatback. One strap is worn overeach shoulder. Each strap terminates in a slotted bayonetblade fastener which is aligned with one edge of the strap.When the two bayonet blades are placed together. theshoulder harness straps can be secured by sliding the malehalf of the seatbelt buckle through the bayonet slots and intothe female half of the seatbelt buckle.

The shoulder harness straps proceed from inertia reels builtinto the crew chairs. Spring loading at the inertia reels keepsthe shoulder harnesses snug, but allows the pilot and copilot

7-20


all the freedom of movemenet normally required in flightHowever, the inertia reels incorporate a locking device thatwill secure the harness straps in the event of sudden forwardmovement.

CABIN

The shoulder harness on passenger chairs consists of asingle strap. It is routed through the top of the seatback andterminates in a triangular metal fastener. The strap is worndiagonally. It runs from the outboard shoulder to the inboardhip area, where it is secured by hooking the metal fasteneraround the securing stud on the male half of the seatbeltbuckle.

The shoulder harness strap proceeds from an inertia reelbuilt into the passenger chair. Spring loading at the inertiareel keeps the shoulder harness strap snug, but allowsconsiderable freedom of movement. However, the inertia reelincorporates a locking device that will secure the harnessstrap in the event of sudden forward movement. If the seat isequipped with a shoulder harness. it must be worn duringtakeoff and landing.

WARNING

Ensure that the seatback is in the fully uprightposition and that the headrest is properlyadjusted whenever the shoulder harness isused.

AFT-CABIN AREA

The shoulder harness for aft-cabin-area fold-up chairs is of adouble-strap configuration. The middle portion of the strap issecured by a metal slip ring which is anchored to the aftpressure bulkhead. The two ends (which actually function astwo separate straps) extend downward toward the seatbeltbuckle area. One end of the shoulder harness strapterminates in a slotted bayonet-blade fastener. The other endis attached to the upper edge of the shoulder harnessadjuster. A short adjusting strap, which is also equipped witha slotted bayonet-blade fastener, extends upward from thearea of the seatbelt buckle and slides through the lowerportion of the shoulder harness adjuster. A small, flexibleadjusting tab is also attached to the lower edge of theadjuster.

One end of the shoulder harness strap is worn over eachshoulder. When the two bayonet blades are placed together,the shoulder harness straps can be secured by sliding themale half of the seatbelt buckle through the bayonet slots andinto the female half of the seatbelt buckle. The shoulderharness strap can be lengthened by grasping the tab on theadjuster and pulling upward. The strap can be tightened bygrasping the loose end of the adjusting strap and pulling itthrough the adjuster until the shoulder harness is snug.

October, 1978


DOORS, WINDOWS, AND EXITS

IAIRSTAIR ENTRANCE DOOR (200)

The cabin door IS hinged at the bottom. It swings out anddown when opened. A stairway built onto the inboard side ofthe door facilitates entry to and egress from the airplane. Twoof the stairsteps automatically fold flat against the door whenthe door is closed. A hydraulic damper ensures that the doorwill swing down slowly when it opens. While the door is open,it is supported by a plastic-encased cable, which also servesas a handrail. Additionally, this cable is utilized when closingthe door from inside the airplane. The door closes against aninflatable rubber seal which is installed around the opening inthe door frame. When weight is off the landing gear, enginebleed air supplies pressure to inflate the door seal, whichprovides a positive pressure-vessel seal around the door.The outside door handle can be locked with a key, forsecurity of the airplane on the ground.

CAUTION

Only one person at a time should be on the doorstairway.

The door locking mechanism is operated by rotating eitherthe outside or the inside door handle, both of which movesimultaneously. Two latch bolts at each side of the door, andtwo latch hooks at the top of the door, lock into the door frameto secure the airstair door.

Whether unlocking the door from the outside or the inside,the release button adjacent to the door handle must be helddepressed before the handle can be rotated(counterclockwise from inside the airplane, clockwise fromoutside) to unlock the door. Consequently, unlocking thedoor is a two-hand operation requiring deliberate action. Therelease button acts as a safety device to help preventaccidental opening of the door. As an additional safetymeasure, a differential-pressure-sensitive diaphragm isincorporated into the release-button mechanism. Theoutboard side of the diaphragm is open to ambient airpressure, the inboard side to cabin air pressure. As thecabin-to-ambient air pressure differential increases, itbecomes increasingly difficult to depress the release button,because the diaphragm moves inboard when either theoutside or inside release button is depressed. Never attemptto unlock or even check the security of the door in flight. If theCABIN DOOR caution annunciator illuminates in flight, or ifthe pilot has any reason whatever to suspect that the doormay not be securely locked, the cabin should bedepressurized (after first considering altitude), and alloccupants instructed to remain seated with their seatbeltsfastened. After the airplane has made a full-stop landing andthe cabin has been depressurized, a crew member shouldcheck the security of the cabin door.

February, 1979


To close the door from outside the airplane, lift up the freeend of the airstair door and push it up against the door frameas far as possible. Then grasp the handle with one hand androtate it clockwise as far as it will go. The door will then moveinto the closed position. Then rotate the handlecounterclockwise as far as it will go. The release buttonshuuld pop out, and the handle should be pointing aft. Checkthe security of the door by attempting to rotate the handleclockwise without depressing the release button; the handleshould not move.

To close the door from inside the airplane, grasp the handrailcable and pull the airstair door up against the door frame.Then grasp the handle with one hand and rotate itcounterclockwise as far as it will go. continuing to pull inwardon the door. The door will then move into the closed position.Then turn the handle clockwise as far as it will go. Therelease button should pop out, and the handle should bepointing down. Check the security of the door by attemptingto rotate the handle counterclockwise without depressing therelease button; the handle should not move. Next, lift thefolded stairstep that is just below the door handle. A placardadjacent to the round observation window advises theobserver that the safety lock arm should be in positionaround the diaphragm shaft (plunger) when the handle is inthe locked position. The placard also presents a diagramshowing how the arm and shaft should be positioned. A redpush-button switch near the window turns on a lamp insidethe door, which illuminates the area observable through thewindow. If the arm is properly positioned around the shaft,proceed to check the indication in each of the visualinspection ports, one of which is located near each corner ofthe door. The green stripe painted on the latch bolt should bealigned with the black pointer in the visual inspection port. Ifany condition specified in this door-locking procedure is notmet, do not take off.

AIRSTAIR ENTRANCE DOOR (200C)

A swing-down door, hinged at the bottom, provides aconvenient stairway for entry and exit. Two of the four stepsare movable and automatically fold flat against the door inthe closed position. A self-storing platform automaticallyfolds down over the door sill when the door opens to providea stepping platform for door seal protection. A plasticencased cable provides support for the door in the openposition, a handhold for passengers, and a means of closingthe door from inside the airplane. A rubber seal around thedoor positively seals the pressure vessel while the airplaneis in flight. A hydraulic dampener permits the door to lowergradually during opening.

The door locking mechanism is operated by either of thetwo vertically staggered handles, one inside and the otheroutside the door. The inside and outside handles aremechanically interconnected. When either handle is rotatedper placard instructions, three rotating-earn-type latches oneither side of the door capture posts mounted on the cargodoor side of the opening. When in the closed position theairstair door becomes an integral part of the cargo door.

7-21


A button adjacent to the door handle. whether inside oroutside the cabin, must be depressed before the handle canbe rotated to open the door. When the airplane cabin ispressurized. cabin internal pressure inflates a bellowsbehind the button to prevent accidental opening of the door.

An annunciator light in the cockpit illuminates if the door isnot closed and all latches fully locked. The handle andlatches can all be visually checked for security. For securityof the airplane on the ground, the door can be locked with akey.

CARGO DOOR (200C)

A swing-up door. hinged at the top, provides cabin accessfor loading large or bulky items. After initial opening force isapplied. gas springs will open the cargo door automatically.The door is counterbalanced and will remain in the openposition. The support rod is used to hold the door in theopen position and to pulthe door down into the closedposition. Once closed, the gas springs apply a closing forceto assist in latching the door. A rubber seal around the doorseals the pressure vessel while the airplane is in flight.

The door locking mechanism is operated by two handles,one in the bottom forward portion of the door and the otherin the upper aft portion of the door. When the upper afthandle is operated per placard instructions, two rotatingcam-type latches on the forward side of the door and two onthe aft side rotate, capturing posts mounted on the fuselageside of the door opening. The bottom handle, whenoperated per placard instructions, actuates four pin-luglatches across the bottom of the door.

A button on the upper aft handle must be pressed beforethe handle can be operated to open or latch the door. Alatching lever on the bottom handle must be lifted to releasethe handle before the lower latches can be opened. Theseact as additional aids in preventing accidental opening ofthe door.

The cabin and cargo doors are equipped with dual sensingcircuits to provide the aircrew remote indication of cabindoor/cargo door security. An annunciator light in the cockpitilluminates if the cabin door or cargo door is open and thebattery switch is turned ON. If the battery switch is turnedOFF the annunciator will illuminate only if the cabin door isclosed but not securely latched. These circuits shall bechecked prior to the first flight of each day.

EMERGENCY EXIT

The emergency exit door, placarded EXIT-PULL, is locatedon the right side of the fuselage at the forward end of thepassenger compartment. From the inside. the door isreleased with a pull-down handle. From the outside, the dooris released with a flush-mounted, pull-out handle. The nonhinged, plug-type door removes completely from the frameinto the cabin when the latches are released. The door canbe locked with a key from the inside, to prevent opening from

7-22


the outside. The inside handle will unlatch the door, whetheror not it is locked, by overriding the locking mechanism. Thekey lock should be unlocked prior to flight, to allow removal ofthe door from the outside in the event of an emergency. Thekeyhole is in the horizontal position when the door is locked.The key cannot be removed in this position.

A wiper-type disconnect for the air duct that supplies air tothe eyeball outlet in the emergency exit door is located on theupper-aft edge of the door. As the door is removed, the ductis disconnected, since it is an integral part of the door.

Located on the lower-forward edge of the door is an electricaldisconnect for the wiring that goes to the reading light and thefluorescent light in the emergency exit door. It will unplug asthe door is being removed. Upon reinstalling the door, theelectrical disconnect should be reconnected before movingthe door into the closed position.

INTERIOR DOORS

Sliding doors are provided between the cockpit and cabin,and between the cabin and foyer. These doors provideprivacy, and preve nt the spi IIi ng of light from onecompartment into another. The doors are closed by slidingthe two partition-type door panels to the center of the aisle,where they are held together by a magnetic strip in the edgeof each door.

CABIN WINDOWS

Each cabin window pane, which is composed of a sheet ofpolyvinyl butyral (PVB) laminated between two sheets ofclear acrylic plastic, is stressed to withstand the cabin-toambient air pressure differential. It is then sealed into awindow opening in the fuselage, and forms an integral part ofthe pressure vessel.

POLARIZED TYPE

Two dust panes are mounted inboard of the cabin windowpane in each window frame. Each of these dust panes iscomposed of a film of polarizing material laminated betweentwo sheets of acrylic plastic. The inboard dust pane rotatesfreely in the window frame and has a protruding thumb knobnear the edge. Rotating the pane through an arc of 90°permits complete light regulation as desired. Rotationchanges the relative alignment between the polarizing films,thus providing any degree of light transmission from fullintensity to almost none.

WARNING

Do not look directly at the sun, even throughpolarized windows, because eye damage couldresult.

february. 1979

eeeCHCRAFTSuper King Air 200

CAUTION

When the airplane is to be parked in areasexposed to intensive sunlight, the polarizedwindows should be rotated to the clear positionto prevent deterioration of the polarized material.Sufficient ultraviolet protection is provided toprevent fading of the upholstery.

SHADE TYPE

A dust pane, which is a single sheet of tinted acrylic plastic, ismounted inboard of the cabin window pane in each windowframe. An adjustable window shade is provided to control theamount of light admitted. The shade is adjusted by squeezingthe two latch handles located on the lower center of theshade. and then positioning the shade as desired. Detents inthe shade tracks provide positive latching action at variouspositions.

CONTROL LOCKS

The control locks consist of a U-shaped clamp and two pins.all connected by a chain. The pins lock the primary flightcontrols; the U-shaped clamp fits around the engine controllevers, serving to warn the pilot not to start the engines withthe control locks installed. It is important that all the Jocks beinstalled and removed together. to preclude the possibility ofattempting to taxi or fly the airplane with the engine controllevers released. but with the pins still installed in the flightcontrols.

Install the control locks in the following sequence:

1. Position the U-c1amp around the engine control levers.2. Move the control column as necessary to align the

holes, then insert the small pin.

NOTE


200-150-.

3. Insert the L-shaped pin through the hole provided in thefloor aft of the rudder pedals. The rudder pedals must becentered to align the hole in the rudder bellcrank withthe hole in the floor. The pin is then inserted until theflange is resting against the floor. This will prevent anyrudder movement.

WARNING

Before starting engines, remove the locks.reversing the above procedure.

I On serials 88-82 and after, 8L-1 and after, theholes are aligned when the control wheel is fullforward and rotated approximately 15° to theleft.

On serials prior to 88-82, the holes are alignedwhen the control wheel is full forward and level.

February, 1979

CAUTION

Remove the control locks before towing theairplane. If towed with a tug while the rudder lockis installed. serious damage to the steeringlinkage can result.

7-23




October, 1978


ENGINES

The Cileechcraft Super King Air 200 is powered by two Pratt &Whitney Aircraft of Canada Ltd. PT6A~41 turbopropellerengines, each rated at 850 SHP. Each engine has a threestage axial-How. single-stage centrifugal-flow compressor,which is driven by a single-stage reaction turbine. The powerturbine - a two-stage reaction turbine counter - rotating withthe compressor turbine - drives the output shaft. Both thecompressor turbine and the power turbine are located in theapproximate center of the engine, with their shafts extendingin opposite directions. Being a reverse flow engine, the ramair supply enters the lower portion of the nacelle and is drawnin through the aft protective screens. The air is then routedinto the compressor. After it is compressed, it is forced intothe annular combustion chamber, and mixed with fuel that issprayed in through 14 nozzles mounted around the gasgenerator case. A capacitance discharge ignition unit andtwo spark igniter plugs are used to start combustion. Aftercombustion, the exhaust passes through the compressorturbine and two stages of power turbine and is routed throughtwo exhaust ports near the front of the engine. A pneumaticfuel control system schedules fuel flow to maintain the powerset by the gas generator power lever. Propeller speed withinthe governing range remains constant at any selectedpropeller control lever position through the action of apropeller governor, except in the beta range where themaximum propeller speed is controlled by the pneumaticsection of the propeller governor.

The accessory drive at the aft end of the engine providespower to drive the fuel pumps, fuel control, the oil pumps, the


refrigerant compressor (right engine), the starter/generator,and the tachometer transmitter. At this point, the speed of thedrive (N 1) is the true speed of the compressor side of theengine, 37,500 rpm (which corresponds to 100% N 1).Maximum continuous speed of the engine is 38,100 rpm,which equals 101.5% N1. with a transient overspeed of38,500 rpm, which equals 102.6% N1'

The reduction gearbox forward of the power turbine providesgearing for the propeller and drives the propeller tachometertransmitter, the propeller overspeed governor, and thepropeller governor. Prior to gear reduction, the turbine speedon the power side of the engine is 30,000 rpm at 2000propeller rpm.

Propeller torque value is measured by a hydro-mechanicaldevice located inside the first stage reduction gear housing toprovide an accurate indication of engine power output. Themechanism consists of a torquemeter cylinder, a piston,valve plunger and spring. Rotation of the first stage ring gearin the reduction gearbox is resisted by the helical splineswhich impart an axial movement to the ring gear andtherefore to the torquemeter piston. A torquemeter valveregulates the input of engine oil into the torque cylinder tostabilize the piston position. The pressure created in thetorque cylinder is plumbed to the torquemeter transmitter togive a relative reading of torque.

Deceleration on the ground is achieved by bringing thepropeller blades through the Beta range into a reversing pitchby utrlizinq the pitch change mechanism. The power leversmust be retarded below the IDLE position by raising them

COMBUSTION CHAMBER

POWER TURBINES COMPRESSOR TURBINE COMPRESSOR

ENGINE INLET

200·2.1·2

ENGINE CUTAWAY

October, 1978


over a detent. Reversing power is available in directproportion to the retarding of the levers in the reversingrange.

PROPULSION SYSTEM CONTROLS

The propulsion system is operated by three sets of controls;the power levers, propeller levers, and condition levers. Thepower levers serve to control engine power. The conditionlevers control the flow of fuel at the fuel control outlet andselect fuel cut off, low idle and high idle functions. Thepropeller levers are operated conventionally and control theconstant speed propellers through the primary governor.

POWER LEVERS

The power levers provide control of engine power from idlethrough take-off power by operation of the gas generator(N 1) governor in the fuel control unit. Increasing N1 rpmresults in increased engine power.

PROPELLER LEVERS

Each propeller lever operates a speeder spring inside theprimary governor to reposition the pilot valve, which results inan increase or decrease of propeller rpm. For propellerfeathering, each propeller lever manually lifts the pilot valveto a position which causes complete dumping of highpressure oil. Detents at the rear of lever travel preventinadvertent movement into the feathering range. Operatingrange is 1600 to 2000 rpm.

CONDITION LEVERS

The condition levers have three positions; FUEL CUT-OFF,LOW IDLE and HIGH IDLE. Each lever controls the idle cutoff function of the fuel control unit and limits idle speed at52% N1 for low idle, and 70% N1 for high idle.

PROPELLER REVERSING

When the power levers are lifted over the IDLE detent, theycontrol engine power through the Beta and reverse ranges.

CAUTION

Propeller reversing on unimproved surfacesshould be accomplished carefully to preventpropeller erosion from reversed airflow and, industy conditions, to prevent obscuring theoperator's vision.

Condition levers, when set at HIGH IDLE, keep the enginesoperating at 70% N1 high idle speed for maximum reversingperformance.

7-26


CAUTION

Power levers should not be moved into thereversing position when the engines are notrunning because the reversing system will bedamaged.

FRICTION LOCKS

Four friction locks are located on the power quadrant of thepedestal. When they are rotated counterclockwise, thepropulsion system control levers can be moved freely. As thefriction locks are rotated clockwise, the control leversprogressively become more resistant to movement, so thatthey will not creep out of position.

ENGINE INSTRUMENTATION

Engine instruments, located on the left of the center portionof the instrument panel, are grouped according to theirfunction. At the top, the ITT (Interstage Turbine Temperature)indicators and torquemeters are used to set take-off power.Climb and cruise power are established with thetorquemeters and propeller tachometers while observing ITTlimits. Gas generator (N 1) operation is monitored by the gasgenerator tachometers. The lower grouping consists of thefuel flow indicators and the oil pressure temperatureindicators.

The In indicator gives an instantaneous reading of enginegas temperature between the compressor turbine and thepower turbines.

The torquemeters give an indication in foot-pounds of thetorque being applied to the propeller.

The propeller tachometer is read directly in revolutions perminute. The N1 or gas generator tachometer is read inpercent of rpm, based on a figure of 37,500 rpm at 100%.Maximum continuous gas generator speed is limited to38,100 rpm or 101.5% N1.

Proper observation and interpretation of these instruments'provide an indication of engine performance and condition.

A propeller synchroscope, located to the left of the oilpressure/temperature indicators, operates to give anindication of synchronization of the propellers. If the rightpropeller is operating at a higher rpm than the left, the face ofthe synchroscope, a black and white cross pattern, spins in aclockwise rotation. Left, or counterclockwise, rotationindicates a higher rpm of the left propeller. This instrumentaids the pilot in obtaining complete synchronization ofpropellers.

October, 1978


PROPELLER SYNCHROPHASER

The propeller synchrophaser automatically matches the rpmof the right propeller (slave propeller) to that of the leftpropeller (master propeller) and maintains the blades of onepropeller at a predetermined relative position with the bladesof the other propeller. To prevent the right propeller fromlosing excessive rpm if the left propeller is feathered while thesynchrophaser is on, the synchrophaser has a limited rangeof authority from the manual governor setting. Normalgovernor operation is unchanged but the synchrophaser willcontinuously monitor propeller rpm and reset the governor asrequired. A magnetic pickup mounted in each propelleroverspeed governor and adjacent to each propeller deicebrush block transmits electric pulses to a transistorizedcontrol box installed forward of the pedestal.

The control box converts any pulse rate differences intocorrection commands, which are transmitted to a steppingtype actuator motor mounted on the right engine cowlforward support ring. The motor then trims the right propellergovernor through a flexible shaft and trimmer assembly toexactly match the left propeller. The trimmer, installedbetween the governor control arm and the control cable,screws in or out to adjust the governor while leaving thecontrol lever setting constant. A toggle switch installedadjacent to the synchroscope turns the system on. With theswitch off, the actuator automatically runs to the center of itsrange of travel before stopping to assure normal functionwhen used again. To operate the system, synchronize thepropeller in the normal manner and turn the synchrophaseron. The system is designed for in-flight operations and isplacarded to be off for take-off and landing. Therefore. withthe system on and the landing gear extended, the cautionflashers and a yellow light on the caution/advisoryannunciator panel, PROP SYNC ON, will illuminate.

The right propeller rpm and phase will automatically beadjusted to correspond to the left. To change rpm, adjustboth propeller controls at the same time. This will keep theright governor setting within the limiting range of the leftpropeller. If the synchrophaser is on but is unable to adjust tothe right propeller to match the left, the actuator has reachedthe end of its travel. To re-center, turn the switch off,synchronize the propellers manually,' and turn the switchback on.

ENGINE LUBRICATION SYSTEM

Engine oil, contained in an integral tank between the engineair intake and the accessory case, cools as well as lubricatesthe engine. An oil radiator located inside the lower nacelle,keeps the engine oil temperature within the operating limits.A thermal element is used to regulate a bypass door whichcontrols the volume of cooling air through the radiator.Engine oil also operates the propeller pitch changemechanism and the engine torquemeter system.

October, 1978


The lubrication system capacity per engine is 3.5 U.S.gallons. The oil tank capacity is 2.3 gallons with 5 quartsmeasured on the dipstick for adding purposes.Recommended oils and oil changing procedures are listed inthe SERVICING section.

MAGNETIC CHIP DETECTOR

A magnetic chip detector is installed in the bottom of eachengine nose gearbox. This detector will activate a red light onthe annunciator panel, L CHIP DETECT or A'CHIP DETECT,to alert the pilot of oil contamination indicating possible orpending engine failure.

STARTING AND IGNITION SYSTEM

Each engine is started by a three-position switch located onthe left subpanel placarded, IGNITION AND ENGINE START- ON - OFF - STARTER ONLY. Each switch may be moveddownward to the STARTER ONLY position to motor theengine for the purpose of clearing it of fuel without the ignitioncircuit on. The switch is spring loaded and will return to thecenter position when released. Moving the switch upward tothe ON position activates both the starter and ignition, andthe appropriate IGNITION ON light on the annunciator panelwill illuminate. When engine speed has accelerated through50% N1 or above on starting, the starter drive action isstopped by placing the switch in the center OFF position.

AUTO IGNITION

The auto ignition system provides automatic ignition toprevent engine loss due to combustion failure. This system isprovided to ensure ignition during takeoff, landing,turbu lence, and penetration of icing or precipitationconditions. Arming the system prior to takeoff and turning thesystem off after landing is required to assure the systembeing armed in the required conditions. To arm the system,move the required ENG AUTO IGNITION switches, locatedon the pilot's subpanel, from OFF to ARM. If for any reasonthe engine torque falls below 400 foot-pounds, the igniter willautomatically energize and the IGNITION ON light on thecaution/advisory annunciator panel will illuminate. Forextended ground operation, the system should be turned offto prolong the life of the igniter units.

INDUCTION AIR SYSTEM

The PT6A~41 is a reverse-airflow engine. The compressorwheels draw ambient air into the engine through theinduction air inlet at the lower front of the engine nacelle. Asairspeed increases, ram air pressure rises, compressing theair inside the induction air duct. The air then flows into anannular inlet-air chamber located at the aft end of the engine

7-27



compartment. It then passes through a protective screen andinto the primary compressor impeller. where it is furthercompressed. Then the air is forced through a stator ring andsuccessively through the second and third axial-flowcompressor stages. It is finally compressed in the centrifugalflow compressor stage, then discharged into the turbineplenum assembly. Air from the plenum enters the annularcombustion chamber through a series of holes in the aft endof the combustion chamber, and mixes with fuel that issprayed into the combustion chamber through 14 nozzlesmounted around the gas generator case. The air-fuel mixtureburns inside the combustion chamber, then the hot gasesexpand forward out of the chamber and pass through thecompressor turbine stage, both stages of the power turbine,and out to the atmosphere through two exhaust ports locatedat the side of each nacelle, near the front.

ICE PROTECTION

ENGINE AIR INLET

ENGINE INLET SCREEN

EXHAUST

EXHAUST GAS ~SCREEN ~

ROUTE OF AIR ~

ENGINE ICE PROTECTION

8AFFlE

Engine exhaust heat is utilized for heating the engine air inletlips. Hot exhaust is picked up by a scoop inside each engineexhaust stack and plumbed downward to connect into eachend of the inlet lip. Exhaust flows through the inside of the lipdownward to the bottom where it is plumbed out through thebottom of the nacelle. No shut-off or temperature indicator isnecessary for this system.

ICE VANES (INERTiAL SEPARATOR SYSTEM)

An inertial separation system is built into each' engine air inletto prevent moisture particles from entering the engine inletplenum under icing conditions. A movable vane and a bypass door are lowered into the airstream when operating invisible moisture at + 5°C or colder, by energizing electricalactuators with the switches, placarded ICE VANE - EXTEND- RETRACT, located in the lower left subpane!. The vanedeflects the ram airstream slightly downward to introduce asudden turn in the airstream to the engine, causing themoisture particles to continue on undeflected, because oftheir greater momentum, and to be discharged overboard.

While in the icing flight mode, the extended position of thevane and by-pass door is indicated by green annunciatorlights, L ICE VANE EXT and R ICE VANE EXT.

In the non-ice-protection mode, the vane and by-pass doorare retracted out of the airstream by placing the ice vaneswitches in the RETRACT position. The green annunciatorlights will extingUish. Retraction should be accomplished at+ 15°C and above to assure adequate oil cooling. The vanesshould be either extended or retracted; there are nointermediate positions.

If for any reason the vane does not attain the selectedposition within 15 seconds, a yellow L ICE VANE or RICEVANE light illuminates on the caution/advisory panel. In this

7-28

event, a mechanical backup system is provided, and isactuated by pulling the T-handles just below the pilot'ssubpanel placarded ICE VANE MANUAL· PULL - LEFTENG • RIGHT ENG. Decrease airspeed to 160 knots or lessto reduce forces for manual extension. Normal airspeed maythen be resumed.

CAUTION

Once the manual override system has beenengaged (i.e., anytime the manual ice vane Thandle has been pulled out), do not attempt toretract or extend the ice vanes electrically, evenif the T-handle has been pushed back in, until theoverride linkage in the engine compartment hasbeen properly reset on the ground. (See themaintenance manual for resetting procedure.)

When the vane is successfully positioned with the manualsystem, the yellow annunciator lights will extinguish. Thevane may also be retracted with the manual system. Duringmanual system use, the electric motor switch position mustmatch the manual handle position for a correct annunciatorreadout.

FUEL CONTROL

The basic engine fuel system consists of an engine drivenfuel pump, a fuel control unit, a fuel manifold dump valve, adual fuel manifold and fourteen fuel nozzles. The automaticfuel drain valves are provided to clear residual fuel afterengine shutdown. The engine fuel control unit works with atemperature compensating unit to supply information for theengine fuel control system. This fuel control unit is ahydromechanical computing and metering device whichdetermines the proper fuel schedule for the engine to provide

October, 1978


the power required. as established by the position of thepower levers. This is accomplished by controlling the speedof the compressor turbine. The temperature compensatoralters the acceleration fuel schedule of the fuel control unit tocompensate for variations in compressor inlet airtemperature. Engine characteristics vary with changes ininlet temperature and the acceleration fuel schedule must inturn be altered to prevent compressor stall and/or excessiveturbine temperature.

FIRE DETECTION SYSTEM

The fire detection system is designed to provide immediatewarning in the event of fire in either engine compartment. Thesystem consists of the following: three photoconductive cellsfor each engine; a control amplifier for each engine; two redwarning lights on the warning annunciator panel, oneplacarded FIRE L ENG, the other FIRE R ENG; a test switchon the copilot's left subpanel; and a circuit breaker placardedFIRE DET on the right side panel. The six photoconductivecell flame detectors are sensitive to infrared radiation. They

section VIISystems Descriptions

are positioned in each engine compartment so as to receiveboth direct and reflected rays. thus monitoring the entirecompartment with only three photocells. Heat level and rateof heat rise are not factors in the sensing method.

Conductivity through the photocell varies in direct proportionto the intensity of the infrared radiation striking the cell. Asconductivity increases. the amount of current from theelectrical system flowing through the flame detecto rincreases proportionally. To prevent stray light rays fromsignaling a false alarm, a relay in the control amplifier closesonly when the signal strength reaches a preset alarm level.When the relay closes. the appropriate left or right warningannunciators illuminate. When the tire has beenextinguished, the cell output voltage drops below the alarmlevel and the relay in the control amplifier opens. No manualresetting is required to reactivate the fire detection system.

The test switch on the copilot's left subpanel, placardedTEST SWITCH - FIRE DET & FIRE EXT, has six positions:OFF - RIGHT EXT - LEFT EXT - 3 . 2 - 1. (If the optionalengine-fire-extinguisher system is not installed, the RIGHT

CONTROLAMPLIfiERS

~Off

liST SWIICHF'al DElICIION

200-482-3

October, 1978

FIRE DETECTION SYSTEM SCHEMATIC

7-29


EXT and LEFT EXT positions on the left side of the testswitch will not be installed.) The three test positions for thefire detector system are located on the right side of the switch(3 - 2 - 1)_ When the switch is rotated from OFF (down) to anyone of these three positions, the output voltage of acorresponding flame detector in each engine compartment isincreased to a level sufficient to signal the amplifier that a fireis present. The following should illuminate: the red pilot andcopilot MASTER WARNING flashers; and, if the optionalengine-fire-extinguisher system is installed, the red lensesplacarded L ENG FIRE - PUSH TO EXT and R ENG FIRE PUSH TO EXT on the fire-extinguisher activation switches.The system may be tested anytime, either on the ground or inflight. The TEST SWITCH should be placed in all threepositions, in order to verify that the circuitry for all six firedetectors is functional. Illumination failure of all the firedetection system annunciators when the TEST SWITCH is inanyone of the three f1ame-detector-test positions indicates amalfunction in one or both of the two detector circuits (one ineach engine) being tested by that particular position of theTEST SWITCH.

FIRE EXTINGUISHER SYSTEM

The optional engine-fire-extinguisher system incorporates apyrotechnic cartridge inside the nacelle of each engine.

LEFT

f T

, )RIGHT .2

OFFTESTSWITCH

FIRE OET • FIRE U:T

BEECHCAAFTSuper King Air 200

When the activation valve is opened, the pressurizedextinguishing agent is discharged through a plumbingnetwork which terminates in strategically located spraynozzles.

The fire extinguisher control switches used to activate thesystem are located on the glareshield at each end of thewarning annunciator panel. Their power is derived from thehot battery bus. Each push-to-actuate switch incorporatesthree indicator lenses. The red lens, placarded L (or) R ENGFIRE ~ PUSH TO EXT, warns of the presence of fire in theengine. The amber lens, placarded D, indicates that thesystem has been discharged and the supply cylinder isempty. The green lens, placarded OK, is provided only for thetest function. To discharge the cartridge, raise the safetywired clear plastic cover and press the face of the lens. Thisis a one-shot system and will be completely expended uponactivation. The amber 0 light will illuminate and remainilluminated, regardless of battery switch position. until thepyrotechnic cartridge has been replaced.

The fire-extinguisher-system test functions incorporated inthe TEST SWITCH - FIRE DET & FIRE EXT test the circuitryof the fire extinguisher pyrotechnic cartridges. Duringpreflight, the pilot should rotate the TEST SWITCH to each ofthe two positions (RIGHT EXT and LEFT EXT) and verify theillumination of the amber 0 light and the green.OK light on

200·482-2

7-30

FIRE EXTINGUISHER SYSTEM SCHEMATIC

October. 1978


each fire-extinguisher-activation switch on the glareshield.

A gage, calibrated in psi, is provided on each supply cylinderfor determining the level of charge. The gages should bechecked during preflight.

PROPELLER SYSTEM

DESCRIPTION

Each engine is equipped with a conventional three-blade,full-feathering, constant-speed, counter-weighted, reversing,variable-pitch propeller mounted on the output shaft of thereduction gearbox. The propeller pitch and speed arecontrolled by engine oil pressure, through single-action,engine-driven propeller governors. Centrifugalcounterweights, assisted by a feathering spring, move theblades toward the low rpm (high pitch) position and into thefeathered position. Governor boosted engine oil pressuremoves the propeller to the high rpm (low pitch) hydraulic stopand reverse position. The propellers have no low rpm (highpitch) stops; this allows the blades to feather after engineshutdown.

Propeller tie-down boots are provided for use on the mooredairplane to prevent windmilling at zero oil pressure.

PRIMARY LOW PITCH STOP

Low pitch propeller position is determined by the primary lowpitch stop which is a mechanically actuated, hydraulic stop.Beta and reverse blade angles are controlled by the powerlevers in the Beta and reverse range.

PROPELLER GOVERNORS

Two governors, a constant speed governor, and anoverspeed governor, control the propeller rpm. The constantspeed governor, mounted on top of the gear reductionhousing, controls the propeller through its entire range. Thepropeller control lever operates the propeller by means of thisgovernor. If the constant speed governor should malfunctionand request more than 2000 rpm, the overspeed governorcuts in at 2080 rpm and dumps oil from the propeller to keepthe rpm from exceeding approximately 2080 rpm. A solenoid,actuated by the PROP GOV TEST & LOW PITCH STOPTEST switch located on the pilot's subpanel, is provided forresetting the overspeed governor to approximately 1830 to1910 rpm for test purposes.

If the propeller sticks or moves too slowly during a transientcondition causing the propeller governor to act too slowly toprevent an overspeed condition, the power turbine governor.

October, 1978

Section VIISlstems Descriptions

contained within the constant speed governor housing, actsas a fuel topping governor. When the propeller reaches 2120rpmI the fuel topping governor limits the fuel flow to the gasgenerator, reducing N1 rpm, which in turn prevents thepropeller rpm from exceeding approximately 2200 rpm.Durinq operation in the reverse range, the fuel toppinggovernor is reset to approximately 95% propeller rpm beforethe propeller reaches a negative pitch angle. This ensuresthat the engine power is limited to maintain a propeller rpmsomewhat less than that of the constant speed governorsetting. The constant speed governor therefore will alwayssense an underspeed condition and direct oil pressure to thepropeller servo piston to permit propeller operation in Betaand reverse ranges.

AUTOFEATHER SYSTEM

The automatic feathering system provides a means ofimmediately dumping oil from the propeller servo to enablethe feathering spring and counterweights to start thefeathering action of the blades in the event of an enginefailure. Although the system is armed by a switch on thesubpanel, placarded AUTOFEATHER - ARM - OFF - TEST,the completion of the arming phase occurs when both powerlevers are advanced above 90% N1 at which time both theright and left indicator lights on the caution/advisoryannunciator panel indicate a fully armed system. Theannunciator panel lights are green, placarded LAUTOFEATHER and R AUTOFEATHER. The system willremain inoperative as long as either power lever is retardedbelow 90% N1 position. The system is designed for use onlyduring take-off and landing and should be turned off whenestablishing cruise climb. During take-off or landing, iftorquemeter oil pressure on either engine drops below aprescribed setting, the oil is dumped from the servo, thefeathering spring starts the blades toward feather, and theautofeather system of the other engine is disarmed.Disarming of the autofeather portion of the operative engineis further indicated when the annunciator indicator light forthat engine extinguishes.

FUEL SYSTEM

The fuel system consists of two separate systems connectedby a valve-controlled crossfeed line. The separate fuelsystem for each engine is further divided into a main andauxiliary fuel system. The main system consists of a nacelletank, two wing leading edge tanks, two box section bladdertanks, and an integral (wet cell) tank, all interconnected toflow into the nacelle tank by gravity. This system of tanks isfilled from the filler located near the wing tip.

The auxiliary fuel system consists of a center section tankwith its own filler opening, and an automatic fuel transfersystem to transfer the fuel into the main fuel system.

When the auxiliary tanks are filled. they will be used first.During transfer of auxiliary fuel, which is automatically

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controlled, the nacelle tanks are maintained full. A swingcheck valve in the gravity feed line from the outboard wingprevents reverse fuel flow. Upon exhaustion of the auxiliaryfuel, normal gravity transfer of the main wing fuel into thenacelle tanks will begin.

An anti-siphon valve is installed at each filler port whichprevents loss of fuel or collapse of a fuel cell bladder in theevent of improper securing or loss of the filler cap.

The two systems are vented through a recessed ram ventcoupled to a protruding heated ram vent on the underside ofthe wing adjacent to the nacelle. One vent is recessed toprevent icing and the protruding vent is added as a backupand is heated to prevent icing.

All fuel is filtered with a firewall-mounted 20 micron filter.These filters incorporate an internal bypass which opens topermit uninterrupted fuel supply to the engine in the event offilter icing or blockage. In addition, a screen strainer islocated at each tank outlet before the fuel reaches the boostand transfer pumps. The main engine driven fuel pump hasan integral strainer to protect the pump.

A fuel drain collector system is provided in the aftcompartment of each engine. The system allows the fuel thatis left in the fuel nozzle manifolds at engine shutdown togravity drain into a collector tank, and eventually be returnedto the nacelle tank.

FUEL PUMPS

The engine driven fuel pump (high pressure) is mounted onthe accessory case in conjunction with the fuel control unit.Failure of this pump results in an immediate flameout. Theprimary boost pump (low pressure) is also engine driven andis mounted on a drive pad on the aft accessory section of theengine. This pump operates when the gas generator (N 1) isturning and provides sufficient fuel for start, take-off, all flightconditions except operation with hot aviation gasoline above20,000 feet altitude, and operation with crossfeed.

An electrically driven standby boost pump (low pressure)located in the bottom of each nacelle tank performs threefunctions; it is a backup pump for use in the event of aprimary fuel boost pump failure, it is for use with hot aviationgasoline above 20,000 feet, and it is used during crossfeedoperations. In the event of an inoperative standby pump,crossfeed can only be accomplished from the side of theoperative pump.

Electrical power to operate the standby boost pumps is. controlled by lever lock toggle switches, placardedSTANDBY PUMP - ON - OFF, located on the fuel controlpanel and is supplied power from two independent sources.One source of power for either the right or the left standbypump is provided through the number 3 or number 4 feederbuses and is protected by a 10-ampere circuit breakerlocated on the fuel control panel. This power is only available

October, 1978


when the master switch is turned on. Another source ofpower comes directly from the battery through the hot batterybus and is protected by dualS-ampere fuses located in theright wing center section. The fuse panel may be servicedthrough an access door on the bottom side of the wingoutboard of the battery. This power source makes poweravailable for the pumps at all times. regardless of the batterymaster switch position. These circuits are protected bydiodes to prevent the failure of one circuit from disabling theother circuit. During shutdown, make certain both standbypump switches are off to prevent battery discharge.

In the event of a primary boost pump failure, the respectivered FUEL PRESS light in the annunciator panel willilluminate. This light illuminates when pressure decreasesbelow 9 to 11 psi. The light will be extinguished by switchingon the standby fuel pump on that side, thus increasingpressure above 9 to 11 psi.

CAUTION

Engine operation with the fuel pressure light onis limited to 10 hours between overhaul, orreplacement, of the engine driven fuel pump.

When using aviation gasoline, during climbs above 20,000feet, the first indication of insufficient fuel pressure will be anintermittent flicker of the FUEL PRESS lights. A widefluctuation of the fuel flow indicator may also be noted. Theseconditions can be eliminated by turning on a standby pump.

AUXILIARY FUEL TRANSFER SYSTEM

The auxiliary tank fuel transfer system automaticallytransters the fuel from the auxiliary tank to the nacelle tankwithout pilot action. Motive flow to a jet pump mounted in theauxiliary tank sump is obtained from the engine fuel plumbingsystem downstream from the engine driven boost pump androuted through the transfer control motive flow valve. Themotive flow valve is energized to the open position by thecontrol system to transfer auxiliary fuel to the nacelle tank tobe consumed by the engine during the initial portion of theflight. When an engine is started, pressure at the enginedriven boost pump closes a pressure switch which, after a 30to 50 second time delay to avoid depletion of fuel pressureduring starting, energizes the motive flow valve. When theauxiliary fuel is depleted, a low level float switch de-energizesthe motive flow valve after a 30 to 60 second time delayprovided to prevent cycling of the motive flow valve due tosloshing fuel.

In the event of a failure of the motive flow valve or theassociated control circuitry, the loss of motive flow pressurewhen there is still fuel remaining in the auxiliary fuel tank issensed by a pressure switch and float switch, respectively,which illuminates a light placarded NO TRANSFER on thefuel control panel. During engine start, the pilot should note

7-33


that the NO TRANSFER lights extinguish 30 to 50 secondsafter engine start. A manual override is incorporated as abackup for the automatic transfer system. This is initiated byplacing the AUX TRANSFER switch, located in the fuelcontrol panel to the OVERRIDE position.

USE OF AVIATION GASOLINE

If you find you must top off the fuel tanks with aviationgasoline as an alternate fuel, you will need to determine howmany hours the airplane is operated on gasoline. Since thegasoline is being mixed with the regular fuel. it is expedient torecord the number of gallons of gasoline taken aboard. Agood rule to follow for determining the number of hours ofoperation on aviation gasoline is:

Each engine will consume approximately 50gallons of fuel per hour. Divide the number ofgallons of gasoline pumped into each side by 50to get,.. the number of hours of operation ongasoline. Example: If 150 gallons of gasoline arepumped into one side, divide by 50 and the totalis 3 hours. This means that the engine shouldhave 3 hours charged against it toward themaximum of 150 hours between-overhaul limit.Maintain a record of hours charged against eachengine.

CROSSFEED

During emergency single engine operation, it may becomenecessary to supply fuel to the operative engine from the fuelsystem on the opposite side. The simplified crossfeedsystem is placarded for fuel selection with a diagram on theupper fuel control panel. Place the standby pump switches inthe OFF position when crossfeeding. A lever lock switch,placarded CROSSFEED FLOW, is moved from the centerOFF position to the left or to the right. depending on directionof fuel flow. This opens the crossfeed valve, energizing thestandby pump on the side from which crossfeed is desired.and de-energizes the motive flow valve in the fuel system onthe side being fed. When the crossfeed mode is energized, agreen FUEL CROSSFEED light on the caution/advisorypanel will illuminate.

FIREWALL SHUTOFF

The system incorporates two firewall shutoff valvescontrolled by two switches, one on each side of the fuelsystem circuit breaker panel. located on the fuel controlpanel. These switches, respectively LEFT and RIGHT, areplacarded FUEL FIREWALL SHUTOFF VALVE - OPEN CLOSED. A red guard over each switch is an aid inpreventing inadvertant operation. Like the boost pumps, thefirewall shutoff valves receive electrical power from the mainbuses and also from the hot battery bus which is connecteddirectly to the battery.


FUEL ROUTING IN ENGINE COMPARTMENT

Just forward of the firewall shutoff valve is/the primary boostpump. From the primary boost pump. the fuel is routed to themain fuel filter, the fuel flow indicator transmitter. through afuel heater that utilizes heat from the engine oil to warm thefuel, through the engine driven fuel pump, then to the fuelcontrol unit. From there it is directed through the dual fuelmanifold to the fuel outlet nozzles and into the annularcombustion chamber. Fuel is al so taken fro m ju s tdownstream of the main fuel filter to supply the jet transferpump motive flow.

FUEL DR~\INS

During each preflight, the fuel sumps on the tanks. pumpsand filters should be bled to check for fuel contamination.There are five sump drains and one filter drain in each wing.They are located as follows:

DRAINS LOCATION

Leading edge tank Outboard of nacelleunderside of wing

Integral tank Underside of wingforward of aileron

Fil ewall fuel filter Underside of cowlingforward of firewall

Sump strainer Bottom center of nacelleforward of wheel well

Gravity feed line Aft of wheel well

Auxiliary tank At wing root justforward of the flap

FUEL DRAIN COLLECTOR SYSTEM

After engine shutdown, a small amount of fuel present in thefuel nozzle manifolds drains into a small collector tank. Thetank is mounted to one of the lower fire shields in the aftengine compartment. An electric float switch senses the tankfuel level and activates an electric pump which then transfersthe fuel back to the nacelle tank. When the couector tank isemptied, the float switch turns off the pump. The entireoperation is automatic and requires no input or additionalduties from the crew.

FUEL GAGING SYSTEM

Fuel quantity in either the main or auxiliary fuel system ismonitored by a capacitance fuel gaging system. Quantity is

October, 1978


read directly in pounds. A maximum 3% error may beencountered in the system. However, the system iscompensated for density changes due to temperatureexcursions. A graph is provided in the CRUISE CONTROLsection to allow more accurate readings for all the approvedjet fuels and for aviation gasoline. A selector switch on thefuel control 'panel, placarded FUEL QUANTITY - MAIN AUXILIARY, allows monitoring of the main or auxiliarysystem fuel. There are two gages, one for .each side.

ELECTRICAL SYSTEM

The airplane electrical system is a 28·VDC (nominal) systemwith the negative lead of each power source grounded to themain airplane structure. DC electrical power is provided byone 34-ampere-hour, air-cooled. 20-cell, nickel-cadmiumbattery, and two 250-ampere starter/generators connected inparallel. The system is capable of supplying power to allsubsystems that are necessary for normal operation of theairplane. A hot battery bus is provided for emergencyoperation of certain essential equipment and the cabin entrythreshold light circuit. Power to the main bus from the batteryis routed through the battery relay which is controlled by aswitch placarded BATT - ON - OFF, located on the leftsubpanel. Power to the bus system from the generators isrouted through reverse-current-protection circuitry. Reversecurrent protection prevents the generators from absorbingpower from the bus when the generator voltage is less thanthe bus voltage. The generators are controlled by switches.placarded GEN 1 and GEN 2, located on the left subpanel.

NOTE

I On Serials 88-88 and After, BL-1 and After: Inorder to turn the generator ON. the generatorcontrol switch must first be held upward in thespring-loaded RESET position for a minimum ofone second, then released to the ON position.

Starter power to each individual starter/generator is providedfrom the main bus through a starter relay. The start cycle iscontrolled by a three-position switch for each engine,placarded IGNITION AND ENGINE START, on the leftsubpanel. The starter/generator drives the compressorsection of the engine through the accessory gearing. Thestarter/generator Initially draws approximately 1100,amperes, then drops rapidly to about 300 amperes as theengine reaches 20% of the gas generator speed.

Power is supplied from three sources: the battery. the rightgenerator, and the left generator. The generator busses areinterconnected by two 325-ampere current limiters. Theentire bus system operates as a single bus, with power beingsupplied by the battery and both generators. There are fourdual-fed sub-busses. Each sub-bus is supplied power fromeither generator main bus through a 60-amp limiter, a70-amp diode, and a 50-amp circuit breaker. All electricalloads are divided among these busses except as noted on

February, 1979


the accompanying Power Distribution Schematic. The equipment on the busses is arranged so that all items withduplicate functions (such as right and left landing lights) areconnected to different busses. Among the loads on thegenerator busses are the number 1 and number 2 inverters.Through relay circuitry, the INVERTER selector switchactivates the selected inverter, which provides 400-hertz,115-volt, alternating current to the avionics equipment, and400~hertz, 26 VAC to the torquemeters (and 26 VAC to thefuel flow indicators on serials prior to BB-225). Thevolt/frequency meter indicates the voltage and frequency ofthe alternating current being supplied to the avionicsequipment.

The generators are controlled by individual voltage regulators which allow a constant voltage to be presented tothe busses during variations in engine speed and electricalload requirements. The generators are manually connectedto the voltage regulating circuits by means of contrbl switcheslocated on the pilot's left subpanel. The voltage regulatingcircuit will automatically disable or enable a generator'scapabilities on the bus. The load on each generator isindicated by the respective left and right volt/loadmeterlocated in the overhead panel.

Overheating of the nickel-cadmium battery will cause thebattery charge current to increase. Therefore, a yellowBATTERY CHG caution annunciator light is provided in thecaution/advisory annunciator panel to alert the pilot of thepossibility of battery overheating. Airplane serials B8-2 thruB8-36 (except those serials modified per Service ,nstructionsNo. 0701-356) are equipped with a Battery Charge CurrentSensor, which will cause illumination of the yellow BATIERYCHG annunciator whenever an increase in the batterycharging current occurs. This module is self-testing in thatthe BATTERY CHG annunciator will illuminate during anengine start, but will extinguish after a few seconds.However, the yellow MASTER CAUTION flasher willcontinue to flash until it is reset by being depressed. just as isthe case when any other caution annunciator has illuminated.

Airplane serials 8B-37 and after and BL-1 and after (and allearlier serials modified per Service Instructions No. 0701356) are equipped with a Battery Charge Current Detector,which will cause illumination of the yellow BATIERY CHGannunciator whenever the battery charge current is abovenormal. Thus the BATTERY CHG annunciator mayoccasionally illuminate for short intervals when heavy loadsswitch off. Following a battery-powered engine start, thebattery recharge current is very high and causes illuminationof the BA lTERY CHG annunciator, thus providing anautomatic self-test of the detector and the battery. As thebattery approaches a full charge and the charge currentdecreases to a satisfactory level, the annunciator willextinguish. This will normally occur within a few minutesafter an engine start, but may require a longer time, if thebattery has a low state of charge, low charge voltage percell (20-cell battery). or low battery temperature. Thissystem is designed for continuous monitoring of the batterycondition.

7-35


'""zQ>.

INV.NO.2

7-36

POWER DISTRIBUTION SCHEMATIC

~~!RIGHT

STARTRELAY

ISOLATION LIMITER


October, 1978


With either system, illumination of the BATTERY CHGannunciator in flight cautions the pilot that conditions mayexist that may eventually damage the battery. The operatorshould check the battery charge current with the loadmeter.This is accomplished by turning off one generator and notingthe load on the remaining generator. Turn off the battery andnote the loadmeter change. If the change is greater than.025, the battery should be left off the bus and should beinspected after landing. If the annunciator remains on afterthe battery switch is moved to the OFF position, amalfunction is indicated in either the battery system or chargecurrent detector, in which case the airplane should be landedas soon as practicable. The battery switch should be turnedON for landing in order to avoid electrical transients causedby power fluctuations.

EXTERNAL POWER

For ground operation, an external power socket, locatedunder the right wing outboard of the nacelle, is provided forconnecting an auxiliary power unit. A relay in the externalpower circuit will close only if the external source polarity iscorrect. The Battery Master Switch should be ON beforeapplying external power, in order to absorb voltage transientswhen operating avionics equipment and during engine starts.Otherwise, the transients might damage the many solid statecomponents in the airplane. (On serials BB-364 and after

land BL-1 and after, the Battery Master Switch must be onbefore the external power relay will close and allow externalpower to enter the airplane electrical system.)

For starting, an external power source capable of supplyingup to 1000 amperes (400 amperes maximum continuous)should be used. A caution light on the caution/advisoryannunciator panel, EXT PWR, is provided to alert theoperator when an external DC power plug is connected to theairplane.

LIGHTING SYSTEMS

COCKPIT

An overhead light control panel, easily accessible to bothpilot and copilot, incorporates a functional arrangement of alllighting systems in the cockpit. Each light group has its ownrheostat switch placarded BRT - OFF. The MASTER PANELLIGHTS switch controls the overhead light control panellights, fuel control panel lights, engine instrument lights, radiopanel lights, subpanel and console lights, pilot and copilotinstrument lights, and gyro instrument lights. The instrumentindirect lights in the glareshieJd and overhead map lights areindividually controlled by separate rheostat switches. Thepush-button FREE AIR TEMP switch in the overhead lightcontrol panel turns on and off the light in the outside airtemperature gage, located in the ceiling near the oxygencontrol.

FebruaryI 1979


CABIN

A three-position switch on the copilot's subpanel, placardedINTR LIGHTS . START BRT - DIM - OFF, controls thefluorescent cabin lights. The switch to the right of the interiorlight switch activates the cabin NO SMOKING/FASTENSEAT BELT signs and accompanying chimes. This threeposition switch is placarded CABIN SIGN - BOTH - OFF FSB.

The two baggage-area lights are controlled by a two-positionswitch just inside the airstair door aft of the door frame.

A threshold light is located forward of the airstair door at floorlevel, and an aisle light is located at floor level aft of the sparcover. A switch adjacent to the threshold light turns boththese lights on and off. This switch also turns the exteriorentry light on and off if a separate switch for the entry light isnot installed adjacent to the threshold light switch. When theairstair door is closed, all the lights controlled by thethreshold light switch will extinguish.

When the master switch is on, the individual reading lightsalong the top of the cabin may be turned on or off by thepassengers with a push-button switch adjacent to each light.

EXTERIOR

Switches for the landing lights, taxi lights, wing ice ·Iights,navigation lights, recognition lights, rotating beacons, andwing-tip and tail strobe lights are located on the pilot'ssubpanel. They are appropriately placarded as to theirfunction.

Tail floodlights, if installed, are incorporated into thehorizontal stabilizers and are designed to illuminate bothsides of the vertical stabilizer. A switch for these lights,placarded TAIL FLOODLIGHT, is located on the overheadlight control panel.

fl. flush-mounted floodlight forward of the flaps in the bottomof the left wing may be installed. This entry light providesillumination of the area around the airstair door, to providepassenger convenience at night. On earlier airplane serials,there are two switches for this light: one just inside the dooron the forward door frame, and one on the pilot's overheadlight control panel, placarded ENTRY LIGHT. On laterserials, there are no separate switches for this light; it iscontrolled by the threshold light switch just inside the door onthe forward door frame, and will extinguish automaticallywhenever. the cabin door is closed.

ENVIRONMENTAL SYSTEM

The environmental system consists of the bleed airpressurization, heating and cooling systems, and theirassociated controls.

7-37

Section VIISystems Descr iptions


CEIUNG OUTlET

CEOUNG= .1---- ---

F\.OOAoensr

CEILINGoonsr

MlXlNG PLENUM 7F""'~~:~~1\=~""';-SH AIR.VALVE.... (CLOSED 'Nl-IEN~Ol

Wll'.:OSo-!:E:..:> oeFRO$TE.~ F\YOEVAPORATOR(ON ~SHlEL::l AIR FlLTEA

FWD PAESSURE FWD EVAPOAA1"OR c=====~BU..Kl-lEAO RE1'\JRN AIR VALVECREW HE.'.T OUCT RET'URN AOR F1l.TER

INSTRUMENT P....'IlE!. ccencr-s VENT A1Rc:oc=<X.

CA!l'N A<Rc:oc=<X.CEILING DUCTrFLOORDUCTOMOER

E:'IN!FlC)N.v.E....'TALBl..EEOAlR.SHUT OFF VALVE

;:>NEUMA.TICBl..EEOA<R

$ENSOA ~:.._~~~~~CABtN AlRCONTROL VALVE

REFRIGERANT UNESAlA INlET

MNOSHlEl.O DEFROSTERCON''''''

ENVt~~NTAL SLEEOA<RF1.OW c:oc=<X.Uo\,l1T INC MOOUlATINGANDSHUTOFF VALVE

PIl.ors CEUNG OUTlET

ENVlRONMENTALBl..EEOA<R$I-I1JTOFF VALVE

A<Rll'UT scooP

CA&N. MEATCONTROL VALVE

Io4QT ENGINE Bl..EEO AIR_

PRESSURE VESSEl.c:l

ENVlRONY.ENTAL. BLEED A1RDREClRCU.ATEO CA&N A1R(AlR-CONOmONEO WH£,'Il

EVAPORATOR IS Qtl.1

PNEUMATICBl.EeOAlRSHUTOF!'VALVE

FIREWAll. \

AlMOo"'AHEAT EXCHANGER

+-- -+=-=-

ENVIRONMENTAL SYSTEM SCHEMATIC

7-38 October. 1978


PRESSURIZATION SYSTEM

The pressurization system is designed to provide a normalworking pressure differential of 6.0:±:.1 psi, which willprovide cabin pressure altitudes of approximately: 3900 feetat an airplane altitude of 20,000 feet: 9900 feet at 31,000feet; and 11,700 feet at 35.000 feet.

Bleed air from the compressor section of each engine isutilized to pressurize the pressure vessel. A flow control unitin the nacelle of each engine controls the pressure of thebleed air and mixes ambient air with it, in order to provide anair mixture suitable for the pressurization function. Themixture flows to the environmental bleed air shutoff valve,which is controlled by a switch placarded BLEED AIR VALVE- LEFT (or) RIGHT - OPEN - ENVIR OFF - INSTR & ENVIROFF in the ENVIRONMENTAL controls group on thecopilot's subpanet. When this switch is in either the ENVIR(onmental air) OFF or the INSTR(ument air) &ENVIR(onmental air) OFF position, the valve is closed. Whenit is in the OPEN position, the air mixture flows through thevalve and to the air-to-air heat exchanger. Depending uponthe position of the bypass valves, a greater or lesser volumeof the air mixture will be routed through or around the heatexchanger. The temperature of the air flowing through theheat exchanger is lowered as heat is transferred to coolingfins, which are in turn cooled by ram airflow through the finsof the heat exchanger. The air leaving both (left and right)bypass valves, is then ducted into a single muffler, locatedunder the right floorboard forward of the main spar, whichhelps ensure quiet operation of the environmental bleed airsystem. The air mixture is then ducted from the muffler intothe mixing plenum, located under the copilot's floorboard.

A partition divides the mixing plenum into two sections. Onesection supplies the floor-outlet duct, and the other suppliesthe ceiling-outlet duct. Both sections receive recirculatedcabin air from the forward vent blower. This air passesthrough the forward evaporator. so it will be cooled if the airconditioner is operating. Even in the event that the forwardvent blower becomes inoperative, some air will still becirculated, due to a special nozzle in the discharge side of themixing plenum.

The environmental bleed air duct is routed into the floor-ductsection of the mixing plenum, then curves back to dischargethe environmental bleed air toward the aft end of the floorduct section of the mixing plenum. Forward of the dischargeend of the environmental bleed air duct, warm air is tappedoff and ducted up through the top of the mixing plenum andinto the crew heat duct, which also receives recirculatedcabin air from the mixing plenum. A valve on the forward sideof the crew heat duct allows air to be tapped off for delivery tothe windshield defroster when the DEFROST AIR knob onthe pilot's left subpanel is pulled out.

The air from the environmental bleed air duct is mixed withrecirculated cabin air (which mayor may not be airconditioned) in the mixing plenum, then routed into the f1ooroutlet duct. This pressurized air is then introduced into the

October, 1978


cabin through the floor registers. Finally. the air flows out ofthe pressure vessel through the outflow valve, located on theaft pressure bulkhead. A silencer on the outflow andsafety/dump valves ensures quiet operation.

The mixture from both flow control units is delivered to thepressure vessel at a rate which can vary from about 8 to 16pounds per minute, depending upon ambient temperatureand pressure altitude. Pressure within the cabin and the rateof cabin-pressure changes are regulated by pneumaticmodulation of the outflow valve, which controls the rate atwhich air can escape from the pressure vessel.

A vacuum-operated safety valve is mounted adjacent to theoutflow valve on the aft pressure bulkhead. It is designed toserve three functions: to provide pressure relief in the eventof malfunction of the normal outflow valve; to allowdepressurization of the pressure vessel whenever the cabinpressure switch is moved into the DUMP position; and tokeep the pressure vessel unpressurized while the airplane ison the ground with the left landing-gear safety switchcompressed. A negative-pressure relief function is alsoincorporated into both the outflow and the safety valves. Thisprevents outside atmospheric pressure's exceeding cabinpressure by more than 0.1 psi during rapid descents. even ifbleed air inflow ceases.

When the BLEED AIR VALVE switches on the copilot'ssubpanel are OPEN (up), the air mixture from the flow controlunits enters the pressure vessel. While the airplane is on theground, a left-landing-gear-safety-switch-actuated solenoidvalve in each flow control unit keeps the ambient-air intakeport closed: allowing only bleed air to be delivered into thepressure vessel. At lift-off, the safety valve closes and theambient air shutoff solenoid valve in the left flow control unitopens; approximately 6 seconds later, the solenoid in theright flow control unit opens. Consequently. by increasing thevolume of airflow into the pressure vessel in stages,excessive pressure bumps during takeoff are avoided.

An adjustable cabin pressurization controller is mounted inthe pedestal. It commands modulation of the outflow valve. Adual-scale indicator dial is mounted in the center of thepressurization controller. The outer scale (CABIN ALT)indicates the cabin pressure altitude which the pressurizationcontroller is set to maintain. The inner scale (ACFT ALT)indicates the maximum ambient pressure altitude at whichthe airplane can fly without causing the cabin pressurealtitude to climb above the value selected on the outer scale(CABIN ALT) of the dial. The indicated value on each scale isread opposite the index mark at the forward (top) position ofthe dial. Both scales rotate together when the cabin altitudeselector knob, placarded CABIN ALT is turned. Themaximum cabin pressure altitude is selected by turning thecabin altitude selector knob until the desired setting on theCABIN ALT dial is aligned with the index mark. Themaximum cabin altitude selected may be anywhere from-1000 to + 10,000 feet MSl. The rate control selector knobis placarded RATE· MIN - MAX. The rate at which the cabinpressure altitude changes from the current value to the


selected value is controlled by rotating the rate controlselector knob. The rate of change selected may be fromapproximately 200 to approximately 2000 feet per minute.

The actual cabin pressure altitude is continuously indicatedby the cabin altimeter, which is mounted in the right side ofthe panel that is located between the caution/advisoryannunciator panel and the pedestal. Immediately to the left ofthe cabin altimeter is the cabin vertical speed (CABINCLIMB) indicator, which continuously indicates the rate atwhich the cabin pressure altitude is changing.

The cabin pressure switch, located to the left of thepressurization controller on the pedestal, is placarded CABINPRESS - DUMP - PRESS - TEST. When this switch is in theDUMP (forward) position, the safety valve is held open, sothat the cabin will depressurize and/or remain unpressurized.When it is in the PRESS (center) position, the safety valve isnormally closed in flight, and the outflow valve is controlledby the pressurization controller, so that the cabin willpressurize. When the switch is held in the spring-loadedTEST (aft) position, the safety valve is held closed,bypassing the landing-gear safety switch, to facilitate testingof the pressurization system on the ground.

Prior to takeoff, the cabin altitude selec'or knob should beadjusted so that the ACFT ALT scale on the indicator dialindicates an altitude approximately 1000 feet above theplanned cruise pressure altitude, and the CII.BIN ALT scaleindicates an altitude at least 500 feel above the take-off fieldpressure altitude. The rate control selector knob should beadjusted as desired; setting the index mark between the 9and tz-o'clock positions will provide the most comfortablecabin rate of climb. The cabin pressure switch should bechecked, to ensure that it is in the PRESS position. As theairplane climbs, the cabin pressure altitude climbs at theselected rate of change until the cabin reaches the selectedpressure altitude. The system then maintains cabin pressurealtitude at the selected value. If the airplane climbs to analtitude higher than the value indexed on the ACFT ALTscale of the dial on the face of the controller, the cabin-toambient pressure differential will reach the pressure reliefsetting of the outflow valve and safety valve. Either or bothvalves will then override the cabin pressurization controller inorder to limit the cabin-to-ambient pressure differential to 6.0:t .1 psi. If the pressure altitude should reach a value of12.500 feet, a pressure-sensing switch mounted on theforward pressure bulkhead will close. This causes the ALTWARN annunciator light to illuminate, warning the pilot ofoperation requiring the use of oxygen. If the auto-deploymentoxygen system is installed. a pressure-sensing switchmounted on the cabin sidewall forward of the emergency exitalso closes, signaling the passenger oxygen masks to dropout. During cruise operation. if the flight plan calls for analtitude change of 1000 feet or more. reselect the newaltitude plus 1000 feet on the CABIN ALT dial.

During descent and in preparation for landing, the cabinaltitude selector should be set to indicate a cabin altitude ofapproximately 500 feet above the landing field pressure

7·40


altitude, and the rate control selector should be adjusted asrequired to provide a comfortable cabin-altitude rate ofdescent. The airplane rate of descent should be controlled sothat the airplane altitude does not catch up with the cabinpressure altitude until the cabin pressure altitude reaches theselected value and stabilizes. Then, as the airplanedescends to and reaches the cabin pressure altitude, thenegative-pressure relief function modulates the outflow andsafety valve poppets toward the fully open position. therebyequalizing the pressure inside and outside the pressurevessel. lI.s the airplane continues to descend below the preselected cabin pressure altitude, the cabin will beunpressurized and will follow the airplane rate of descent totouchdown.

FLOW CONTROL UNIT

Each flow control unit consists of an ejector and an integralbleed air modulating valve, firewall shutoff valve, ambient airmodulating valve, and a check valve that prevents the bleedair from escaping through the ambient air intake. The flow ofbleed air through the flow control unit is controlled as afunction of atmospheric pressure and temperature. Ambientair flow is controlled as a function of temperature only. Whenthe BLEED AIR VII,LVE switches on the copilot's subpanelare OPEN, an electric solenoid valve on each flow controlunit opens to allow the bleed air into the unit. As the bleed airenters the flow control unit, it passes through a filter beforegoing to the reference pressure regulator. The regulator willreduce the pressure to a constant value (18 to 20 psi). Thisreference pressure is then directed to the variou scomponents within the flow control unit that regulate theoutput to the cabin. One reference pressure line is routed tothe firewall Shutoff valve located downstream of the ejector.An orifice is placed in the line immediately before the shutoffvalve to provide a controlled opening rate. At the same time,the reference pressure is directed to the ambient airmodulating valve located upstream of the ejector. Apneumatic thermostat with a variable orifice is connected to

.NIU..O""'Ir ... NIUMAlIC.1HUMOS.IAY)

~~lG S~Af(T't' SW~lCH

::.. ::":::J I;

tcll~~O~D I~c:::::::.·.~...:::::-'.J

BLEED AIR FLOW CONTROL UNIT

October, 1978


the modulating valve. The pneumatic thermostat is locatedon the lower aft side of the fireseal forward of the firewall. Thebimetalic sensing discs of the thermostat are inserted into thecowling intake. These discs sense ambient temperature andregulate the size of the thermostat orificies. Warm air willopen the orifice; cold air will restrict it until. at - 30°F, theorifice will completely close. When the variable orifice isclosed. the pressure buildup will cause the modulating valveto close off the ambient air source. An electric solenoid valvelocated in the line to the pneumatic thermostat is wired to theleft landing gear safety switch. When the airplane is on theground, the solenoid valve is closed, thereby directing thepressure to the modulating valve, causing it to shut off theambient air source. The exclusion of ambient air allows fastercabin warmup during cold weather operation. An electriccircuit containing a time-delay relay is wired to the abovementioned solenoid valves to allow the left valve to operateapproximately 6 seconds before the right valve. Thisprecludes the simultaneous openmg of the shutoff valves,which would result in a sudden pressure surge into the cabin.A check valve. located downstream from the modulatingvalve, prevents the loss of bleed air through the ambient airintake. At the same time the above operations have beentaking place in the control unit. reference pressure is directedto the ejector flow control actuator. This actuator isconnected to another variable orifice of the pneumaticthermostat and a variable orifice controlled by an isobaricaneroid. The thermostat orifice is restricted by decreasingambient temperature, and the isobaric aneroid orifice isrestricted by decreasing ambient pressure. The restriction ofeither orifice will cause a pressure buildup on the ejector flowcontrol actuator, permitting more bleed air to enter theejector.

UNPRESSURIZED VENTILATION

Fresh-air ventilation is provided by two sources. One source.which is available during both the pressurized and theunpressurized mode. is the bleed air heating system. This airmixes with recirculated cabin air and enters the cabin throughthe floor registers. The volume of air from the floor registers isregulated by moving a sliding handle at the side of eachinboard-facing register.

The second source of fresh air, which is available during theunpressurized mode only, is ambient air obtained (through acheck valve) from the condenser section in the nose of theairplane. During pressurized operation, cabin pressureforces the check valve closed. During the unpressurizedmode. a spring holds the check valve open, so that theforward blower can draw this air into the cabin. The ambientair then mixes with recirculated cabin air, goes through theforward blower, through the forward evaporator (if it isoperating, the air will be cooled), into the mixing plenum. intoboth the ceihnq-outlet and the floor-outlet duct, and into thecabin through all the ceiling and floor outlets. Air ducted toeach individual ceiling eyeball outlet can be directionallycontrolled by moving the eyeball in the socket. Volume isregulated by twisting the outlet to open or close the damper.

October, 1978


HEATING

When air is compressed, its temperature is increased.Therefore. the bleed air extracted from the compressorsection of each engine for pressurization purposes is hot.This heat is utilized to warm the cabin.

When the left landing gear safety switch is in the on-theground position, the ambient air valve in each flow controlunit is closed. Consequently, only bleed air is delivered to theenvironmental bleed air duct when the airplane is on theground. In flight, the ambient air valve is open. and ambientair is mixed with the engine bleed air in the flow control unit.This environmental bleed air mixture is then routed into thecabin.

If the environmental bleed air mixture is too warm for cabincomfort, the bypass valve routes some or all of it through theair-to-air heat exchanger. located in the wing center section.The position of the damper in the cabin-heat control valve isdetermined by positioning of the controls in theENVIRONMENTAL group on the copilot's subpanel. An airintake on the leading edge of the inboard wing brings ram airinto the heat exchanger to cool the bleed air. After leaving theheat exchanger, the ram air is ducted overboard throughlouvers on the underside of the wing.

After the bleed air passes through or around the air-to-airheat exchanger, it is ducted to the mixing plenum. Some ofthis environmental bleed air is tapped off and delivered to thepilot/copilot heat duct. which is located below the instrumentpanel. An outlet at each end of this duct is provided to deliverwa rm air to the pilot and copilot. P. mechanically controlleddamper in each outlet permits the volume of airflow to beregulated. The pilot's damper is controlled by the PILOT AIRknob. located on the pilot's subpanel just below and outboardof the control column. The copilot's damper is controlled bythe CO-PILOT AIR knob. located on the copilot's subpaneljust below and outboard of the control column. TheDEFROST AIR control knob is located on the puots subpaneljust below and inboard of the control column. This knobcontrols a valve at the forward side of the pilot/copilot heatduct which admits air to two ducts that deliver the warm air tothe defroster, located just below the windshields in the top ofthe glareshield.

The remainder of the air in the environmental bleed air duct isdischarged into the floor-outlet-duct section of the mixingplenum and mixed with recirculated cabin air. This air mixtureis then ducted aft through the floor-outlet duct. If the airtemperature inside this duct becomes excessive. a sensorinside the duct causes the yellow DUCT OVERTEMP cautionannunciator to illuminate. Refer to the ILLUMINATION OF"DUCT OVERTEMP" ANNUNCIATOR procedure in theEMERGENCY PROCEDURES Section of this manual forcorrective action.

After passing the temperature sensor, the air passes throughthe cabin air control valve. This valve is controlled by theCABIN AIR control knob on the copilot's subpanel, just below

7-41


and inboard of the control column. When this knob is pulledout to the stop, only a minimum amount of warm air will bepermitted to pass through the valve, thereby increasing theamount of warm air available to the pilot and copilot heatoutlets, and to the defroster. When this knob is pushed fullyin, the valve is open and the air in the duct will be directed tothe tloor-outlet registers in the cabin.

RADlt~NT HEATING

A supplemental electric radiant heating system is availablefor cabin comfort. It is turned on and off by a switch in theENVIHONMENTAL group on the copilot's subpanelplacarded RADIANT HEAT. This system can be used inconjunction with an auxiliary power unit for warming the cabinprior to starting the engines, and it can be used assupplemental heat in flight. However, it should be used inconjunction with the manual temp control mode only.

IA radiant heater element, installed in the cargo door, iscontrolled by the Cabin Temperature Mode switch andoperates in all heating modes. This unit providessupplemental heat to the cabin for additional passengercomfort.

AIR CONDITIONING SYSTEM

Cabin air conditioning is provided by a refrigerant-gas-vaporcycle refrigeration system consisting of: a belt-driven I

engine-mounted compressor, installed on the right-engineaccessory pad; refrigerant plumbing; and N1 speed switch;high- and low-pressure-protection switches; a condensercoil; a condenser blower; an evaporator; a receiver-dryer; anexpansion valve; and a bypass valve. The plumbing from thecompressor is routed through the right·wing inboard leadingedge to the fuselage. It is then routed forward to thecondenser coil, receiver-dryer, expansion valve, bypassvalve, and evaporator, which are all located in the nose of theairplane.

The high- and low-pressure-limit switches and the N1 enginespeed switch are provided to prevent compressor operationoutside of established limitation parameters. The N1 speedswitch will prevent the flow of electric current to thecompressor clutch when the engine speed is below 60% N1.When the N1 speed swuch is open and there is a demand forair conditioning. the green AIR CND N 1 LOW advisoryannunciator will illuminate. If either the high- or lowcompressor-pressure limit is exceeded. the correspondinghigh- or low-pressure switch, located in the right wing centersection leading edge, will cause blowing of the 7.5-amperecompressor-pressure-Iimit fuse which is inline between thepower source and the electrically activated compressorclutch. The resulting interruption of the power will cause thecompressor clutch to disengage. The fuse is accessiblethrough an access door on the underside of the right wing,just outboard of the battery _The system should be thoroughlychecked before replacing a blown fuse. The compressorclutch circuit breaker is located in the DC power distributionpanel inside the lower forward equipment bay.

7-42

BEECHCP.AFTSuper King Air 200

The forward evaporator utilizes a solenoid-operated hot -qasbypass valve to prevent Icing. A 33° F thermal switch on theforward evaporator controls the valve solenoid.

The forward vent blower blows recirculated cabin air (plusoutside ambient air if the cabin is unpressurized) through theforward evaporator, into the mixing plenum, and into both thefloor-outlet and ceiling-outlet ducts. If the cooling mode isoperating, refrigerant will be circulating through theevaporator and the air leaving it will be cool. All the airentering the ceiling-outlet duct will be cool. This air isdischarged through "eyeball" outlet nozzles in the cockpitand cabin. Each nozzle is movable, so that the airstream canbe directed as desired. When the nozzle is twisted, a damperopens or closes to regulate airflow volume.

Cool air will enter the floor-outlet duct, but in order to providecabin pressurization. warm environmental bleed air will,alsoenter the floor-outlet duct anytime either BLEED AIR valve isOPEN. Therefore, pressurized air discharged from the floorregisters will always be warmer than that discharged at theceiling outlets, no matter what temperature mode is in use. Alever on each floor-outlet register (except the forward-facingregister in the baggage compartment) can be movedvertically to regulate the airflow volume.

A vane-axial blower in the hose section' draws ambient airthrough the condenser when the cooling mode is operating.On serials 8B-345 and after and BL-1 and after (and any.earlier serials that have complied with 'i?<:<:chuafr ServiceInstructions No. 0968 by installation of Kit Number 101·5035-1 S or 101-5035-3 S) this blower shuts off when theairplane is airborne. The current limiter for the blower islocated in the DC power distribution panel in the lowerforward equipment bay.

The receiver-drier and sight gage are located high in thecondenser compartment. They can be viewed by removingthe upper-compartment access panel, located on top of thenose section just left of the centerline.

An optional aft evaporator and blower may be installed in thefuselage center aisle equipment bay behind the rear sparRefrigerant will flow through the aft evaporator anytime itflows through the forward evaporator, However. it will provideadditional cooling only when the aft blower is operating..recirculating cabin air through the aft evaporator anddelivering it to the aft floor and ceiling outlets. See theBLOWER CONTROL description for details concerningoperation of the aft blower.

ENVIRONMENTAL CONTROLS

The ENVIRONMENTAL control section on the copilot'ssubpanel provides for automatic or manual control of thesystem. This section contains all the major controls of theenvironmental function: bleed air valve switches; a forwardvent blower control switch; an aft evaporator on/off switch; amanual temperature switch for control of the cabintemperature control valves in the air-to-air heat exchangers;

February, 1979


a cabin-temperature-Ievel control; and the cabin temp modeselector switch, for selecting automatic heating or cooling,manual heating or cooling, or off. Four additional manualcontrols on the main instrument subpanels may be utilized forpartial regulation of cockpit comfort when the cockpit partitiondoor is closed and the cabin comfort level is satisfactory.They are: pilot's air, defroster air, cabin air, and copilot's aircontrol knobs. The fully out position of all these controls willprovide the maximum heating to the cockpit, and the fully inposition will provide minimum heating to the cockpit.

For warm flights, such as short, low-altitude flights insummer, all the cabin floor registers and ceiling outletsshould be fully open for maximum cooling. For cold flights,such as high-altitude flights, night flights, and flights in coldweather, the ceiling outlets should all be closed and the flooroutlets fully open for maximum heating in the cabin.

If the cabin temperature is comfortable but the cockpittemperature is not, the following procedures are suggested:

HEA TING MODE

If the cockpit is too cold:

1. PILOT AIR, CO-PILOT AIR. and DEFROST AIR Knobs- PULLED FULLY OUT, or as required.

2. CABIN AIR Knob - PULLED OUT IN SMALLINCREMENTS (Allow 3 to 5 minutes after eachadjustment for system to stabilize.)

If the cockpit is too hot:

1. PILOT AIR, CO-PILOT AIR, DEFROST AIR, andCABIN AIR Knobs - PUSHED FULLY IN, or as required.

COOLING MODE

If the cockpit is too cold:

1. PILOT AIR, CO-PILOT AIR, and DEFROST AIR Knobs- PUSHED FULLY IN, or as required.

2. Cockpit Overhead Eyeball Outlets - CLOSED, or asrequired.

If the cockpit is too hot:

1. PILOT AIR and CO-PILOT AIR Knobs - PULLEDFULLY OUT, or as required.

2. CABIN AIR Knob - PUSHED IN IN SMALLINCREMENTS (Allow 3 to 5 minutes after eachadjustment for system to stabilize.)

NOTE

If the CABIN AIR knob is fully in before obtainingsatisfactory cockpit temperature, it may be

October, 1978


necessary to place the aft vent blower switch inthe ON position, so that cabin air will recirculatethrough the aft evaporator to provide additionalcooling.

AUTOMA TIC MODE CONTROL

When the CABIN TEMP MODE selector switch on thecopilot's subpanel is in the AUTO position, the heating andair conditioning systems operate automatically. The systemsare connected to a control box by means of a balancedbridge circuit. When the temperature in the cabin hasreached the selected setting, the automatic temperaturecontrol modulates the bypass valves to allow heated air tobypass the air-to-air heat exchangers in the wing centersections. The warm bleed air is mixed with recirculated cabinair (which mayor may not be air-conditioned) in the forwardmixing plenum.

When the automatic control drives the environmental systemfrom a heating mode to a cooling mode, the cabin-heatcontrol valves close. When the left valve reaches the fullyclosed position, the refrigeration system will begin cooling,provided the right engine speed is above 60% N1. When thebypass valve is opened to approximately the 30° position, therefrigeration system will turn off .:

The CABIN TEMP - INCR control provides regulation of thetemperature level in the automatic mode. A temperaturesensing unit in the cabin, in conjunction with the controlsetting, initiates a heat or cool command to the temperaturecontroller, requesting the desired pressure-vesselenvironment. A duct anticipator temperature probe (duct stat)allows the system to anticipate changes in temperature ofinlet air, thereby providing more even temperature control.

MANUAL MODE CONTROL

When the CABIN TEMP MODE selector is in the MAN HEATor MAN COOL position, regulation of the cabin temperatureis accomplished manually by momentarily holding theMANUAL TEMP switch to either the INCR or DECR positionas desired. When released. this switch will return to thecenter (no change) position. Moving this switch to the INCRor DECR position results in modulation of the cabin-heatcontrol valves in the bleed air lines. Allow approximately 30seconds per valve (1 minute total time) for the valves to moveto the fully open or fully closed position. Only one valve at atime moves. Movement of these valves varies the amount ofbleed air routed through the air-to-air heat exchanger.Consequently, the temperature of the incoming bleed air willvary. This bleed air mixes with recirculated cabin air (whichwill be air-conditioned if the refrigeration system is operating)in the mixing plenum, and is then ducted to the floor registers.As a result. the cabin temperature will vary according to theposition of the cabin-heat control valves, whether or not theair conditioner is operating.

7-43


When the CABIN TEMP MODE selector is in the MAN COOLposition, the air conditioner system will operate, provided thespeed of the right engine is above 60% N1.

NOTE

I On serials 88-345 and after and 8L-1 and after(and any earlier serials that have complied with'ileechcrah Service Instructions 0968 byinstallation of Kit Number 101-5035-1 S or 1015035-3 S), the air conditioner compressor willnot operate unless the cabin-heat control valvesare closed. To ensure that the valves areclosed, hold the MANUAL TEMPerature switchin the DECRease position for one minute.

BLEED AIR CONTROL

Bleed air entering the cabin is controlled by the switchesplacarded BLEED AIR VALVE - OPEN - ENVIR OFF - INSTR& ENVIR OFF. When the switch is in the OPEN position, theenvironmental flow control unit and the pneumatic instrumentair valve are open. When the switch is in the ENVIR OFFposition, the environmental flow control unit is closed and thepneumatic instrument air valve is open; in the INSTR &ENVIR OFF position, both are closed. For maximum coolingon the ground. turn the bleed air valve switches to the ENVIROFF position.

VENT BLOWER CONTROL

The forward vent blower is controlled by a switch in theENVIRONMENTAL group placarded VENT BLOWER - HI LO - AUTO. When this switch is in the AUTO position, theforward vent blower will operate at low speed if the CABINTEMP MODE selector switch is in any position other thanOFF (i.e., MANual COOL, MANual HEAT, or AUTOmatic).

When the VENT BLOWER switch is in the AUTO positionand the CABIN TEMP MODE selector switch is in the OFFposition. the blower will not operate. Anytime the VENTBLOWER switch is in the LO position, the forward ventblower will operate at low speed, even if the CABIN TEMPMODE selector switch is OFF. Anytime the VENT BLOWERswitch is in the HI position, the forward vent blower willoperate at high speed, regardless of the position of theCABIN TEMP MODE selector switch (i.e., MAN COOL, M.A,NHEAT, OFF, or AUTO).

If the optional aft evaporator unit is installed in the airconditioning system, an aft blower is also installed under thefloor.

The aft blower draws in cabin air, blows it across the aftevaporator, and to the aft floor and ceiling outlets. Thisblower operates at high speed only. It is turned on and off bythe AFT BLOWER switch in the ENVIRONMENTAL group onthe pilot's subpanel, and is independent of any other control.

7-44


The aft blower is intended for use only when maximum cabincooling (air conditioning) is desired. If the blower should beturned on during a heating mode of operation, the doorbetween the aft-blower duct and the warm-air (floor-outlet)duct will open. This will stop the flow of heated air to the aftfloor registers, and deliver recirculated cabin air (which is notcooled, since refrigerant is not flowing through the aftevaporator) to the aft floor registers and ceiling outlets.

NOTE

On serials prior to 88-39, some airplanes weredelivered with a two-speed aft blower which didnot have a separate AFT BLOWER switch, butwas controlled by the forward VENT BLOWERswitch and a special temperature sensor. In suchan installation, operation of the aft blower isentirely automatic and cannot be controlled bythe pilot.

Both blower circuit breakers are located in the DC powerdistribution panel in the lower forward equipment bay.

OXYGEN SYSTEM

The Super King Air 200 has several oxygen systemsavailable and may utilize a combination of the systems.These systems are based on an adequate flow for an altitudeof 31,000 feet. The masks and Oxygen Duration Chart(NORMAL PROCEDURES Section) are based on 3.7 SLPM(Standard Uters Per Minute). The only exception is thediluter·demand crew mask when used in the 100% mode.For oxygen duration computation, each diluter-demand maskbeing used in the 100% mode is counted as two masks at 3.7SLPM.

MANUAL PLUG-IN SYSTEM

The manual plug·in system is of the constant-flow type. Eachmask plug is equipped with its own regulating orifice. Thepilot and copilot oxygen masks are kept under the crewseats. Oxygen outlets are located on the forward cockpitsidewalls. Passenger masks are kept in seatback pocketsexcept in the couch installation, in which case they are storedunder the couch. The cabin outlets are located on the cabinheadliner at the top center and at both the forward and aftends of the cabin. When not in use, the cabin outlets areprotected by access doors. All masks are easily plugged inby pushing the orifice in firmly and turning clockwiseapproximately one-quarter turn. Unplugging is easilyaccomplished by reversing the motion.

AUT()..DEPLOYMENT SYSTEM

A push/pull handle (PULL ON - SYStem READY) located aftof the overhead light control panel is used in conjunction with

February, 1979


the automatically deployed passenger oxygen system. Thishandle operates a cable which opens and closes the shut-offvalve located at the oxygen supply bottle in the aft,unpressurized area of the fuselage. When this handle ispushed in, no oxygen supply is available anywhere in theairplane. It should be pulled out prior to engine starting, toensure that oxygen will be immediately available anytime it isneeded. When this handle is pulled out, the primary oxygensupply line is charged with oxygen, provided the oxygensupply bottle is not empty. The primary oxygen supply linedelivers oxygen to the two crew oxygen outlets in the cockpit,to the first aid oxygen outlet in the toilet area, and to thepassenger oxygen system shut-off valve.

When the auto-deployment passenger oxygen system isinstalled, the crew is normally provided with diluter-demand,quick-donning oxygen masks. These masks hang on the aftcockpit partition behind and outboard of the pilot and copilotseats. They are held in the armed position by spring-tensionclips, and can be donned immediately with one hand. Thediluter-demand crew masks deliver oxygen to the user onlyupon inhalation. Consequently, there is no loss of oxygenwhen the masks are plugged in and the PULL ON - SYStemREADY handle is pulled out, even though oxygen isimmediately available upon demand.

A small lever on each diluter·demand oxygen mask permitsthe selection of two modes of operation: NORMAL and100%. In the NORMAL position, air from the cockpit is mixedwith the oxygen supplied through the mask. This reduces therate of depletion of the oxygen supply, and it is morecomfortable to use than 100% aviator's breathing oxygen.However, in the event of smoke or fumes in the cockpit, the100% position should be used to prevent the breathing ofcontaminated air. For this reason, the selector levers shouldbe left in the 100% position when the masks are not in use.

Anytime the primary oxygen supply line is charged, oxygencan be obtained from the first aid oxygen mask, located in thetoilet area, by manually opening the overhead access door(placarded FIRST AID OXYGEN - PULL) and opening theon/off valve inside the box. A placard (NOTE: CREWSYStem MUST BE ON) reminds the user that the PULL ON •SYStem READY handle in the cockpit must be pulled outbefore oxygen will flow from the first aid oxygen mask.

The auto-deployment passenger oxygen system is of theconstant-flow type. Anytime the cabin pressure altitudeexceeds approximately 12,500 feet, a barometric-pressureswitch automatically energizes a solenoid which opens thepassenger oxygen system shut-off valve. The pilot can openthe valve manually anytime by pulling out the PASSENGERMANUAL OverRIDE handle, located aft of the overhead lightcontrol panel. Once the passenger oxygen system shut-offvalve has been opened (either automatically or manually),oxygen will flow into the passenger oxygen supply line, if theprimary oxygen system line has been charged (i.e., if theoxygen supply bottle contains oxygen and the PULL ON •SYStem READY handle in the cockpit is pulled out). Whenoxygen flows into the passenger oxygen system supply line,

October, 1978


a pressure-sensitive switch in the line closes a circuit toilluminate the green PASS OXYGEN ON annunciator on thecautionary/advisory annunciator panel. On serials beginningwith the 1979 model year, this switch will also cause thecabin lights to illuminate in the full bright mode, regardless ofthe position of the interior lights switch (lNTR LIGHTS) on thecopilot's left subpanel.

The pressure of the oxygen in the passenger oxygen systemsupply line then automatically extends a plunger againsteach of the passenger oxygen mask dispenser doors, forcingthe doors open. The oxygen masks then drop down about 9inches below the dispensers. The lanyard valve pin at the topof the oxygen mask hose must be pulled out in order foroxygen to flow from the mask. The pin is connected to theoxygen mask via a flexible cord; when the oxygen mask ispulled down for use, the cord pulls the pin out of the lanyardvalve. The lanyard valve pin must be manually reinserted intothe valve in order to stop the flow of oxygen when the mask isno longer needed. The passenger oxygen can be shut offand the remaining oxygen isolated to the crew and first aidoutlets by pulling the OXYGEN CONTROL circuit breaker inthe ENVIRONMENTAL group on the right side panel,provided the PASSENGER MANUAL O'RIDE handle ispushed in to the off position.

PITOT AND STATIC SYSTEM

The pitot and static system provides a source of impact airand static air for operation of the flight. instruments. P.heated

pitot mast is located on each side of the lower portion of thenose. Tubing from the left pitot mast .is connected to thepilot's airspeed indicator, and tubing from the right pitot mastis connected to the copilot's airspeed indicator.

The normal static system provides two separate sources ofstatic air - one for the pilot's flight instruments, and one forthe copilot's. Each of the normal static air lines opens to theatmosphere through two static air ports - one on each sideof the aft fuselage.

An alternate static air line is also provided for the pilot's flightinstruments. In the event of a failure of the pilot's normalstatic air source (e.g.. if ice accumulations should obstructthe static air ports), the alternate source should be selectedby lifting the spring-clip retainer off the PilOT'S STATIC AIRSOURCE valve handle. located on the right side panel. andmoving the handle aft to the ALTEANATE position. This willconnect the alternate static air line to the pilot's flightinstrumerns. The alternate line obtains static air just aft of therear pressure bulkhead, from inside the unpressurized areaof the fuselage.

7-45


WARNING

The pilot's airspeed and altimeter indicationschange when the alternate static air source is inuse. Refer to the Airspeed Calibratiorl - AlternateSystem, and the Altimeter Correction - AlternateSystem graph in the PERFORMANCE Sectionfor operation when the alternate static air sourceis in use.

When the alternate static air source is not needed, ensurethat the PILOTS STATIC AIR SOURCE valve handle is heldin the forward (NORMAL) position by the spring-clip retainer.

Three petcocks are provided to facilitate draining moisturefrom the static air lines. They are located behind the STATICAIR LINE DRAIN access cover below the circuit breakers onthe right side panel. The drain valves should be opened torelease any trapped moisture at each 100-hour inspection,and after exposure to visible moisture on the ground. Theymust be closed after draining.

ENGINE BLEED AIR PNEUMATIC SYSTEM

High-pressure bleed air from each engine compressor,routed through the firewall shutoff valves and regulated at 18psi, supplies pressure for the surface deice system andvacuum source. Vacuum for the flight instruments is derivedfrom a bleed air ejector. One engine can supply sufficientbleed air for all these systems.

During single-engine operation, a check valve in the bleed airline from each engine prevents flow back through the line onthe side of the inoperative engine. A suttion gage calibratedin inches of mercury, located on the copilot's subpanel,indicates instrument vacuum. To the right of the suction gageis a pneumatic pressure gage, calibrated in pounds persquare inch, which indicates air pressure available to thedeice distributor valve.

BLEED AIR WARNING SYSTEM

The bleed air lines from the engines to the cabin are shieldedwith insulation to protect other components from heat. Heat isalso dissipated in the air-to-air heat exchanger in the centerwing section. The bleed air lines are accompanied in closeproximity by plastic tubing from the engines to the cabin. Oneend of the tubing is plugged off; the other end is connected toa bleed air source in the cabin, to supply the line withpressure. Excessive heat on the plastic tubing caused by aruptured bleed air line will cause the tubing to fail. Uponrelease of pressure in the tubing, a normally open switch inthe line, located under the copilot's floor in the fuselage, willclose, causing a circuit to be completed to the respective BlAIR FAIL light in the warning annunciator panel. When theindication of bleed air line failure becomes evident, the bleedair for that side should be turned off by placing the respective

7-46


lever-lock BLEED AIR VALVE switch on the copdot'ssubpanel in the INSTR & ENVIR - OFF position.

AUTOMATIC DEVICES IN THE CONTROLSYSTEM

YAW DAMP

A yaw damp system is provided to aid the pilot in maintainingdirectional control, and to increase ride comfort. The systemmay be used at any altitude, and is required for flight above17,000 feet. It should be deactivated for takeoff and landing.

If the airplane is equipped with an autopilot, the yaw dampsystem will be a part of the autopilot. Operating instructionsfor this system will be contained in the appropriate AirplaneFlight Manual Supplement.

If an autopilot is not installed in the airplane, yaw damping isprovided by an independent yaw damp system. Thecomponents include a yaw sensor, amplifier, and controlvalve. Regulated air pressure from the control valve isdirected to the same pneumatic servos used for the rudderboost system. The system (on airplanes without autopilots) iscontrolled by a YAW DAMP switch adjacent to the RUDDERBOOST switch on the pedestal. In the event the YAW DAMPswitch is inadvertently left ON during takeoff or landing, thecircuit for the yaw damping system will be interrupted by theleft landing gear safety switch while the airplane is on theground, rendering it inoperative.

STALL WARNING SYSTEM

The stall warning system consists of a transducer, a liftcomputer, a warning horn, and a test switch. Angle of attackis sensed by aerodynamic pressure on the lift transducervane located on the left wing leading edge. When a stall isimminent, the output of the transducer activates a stallwarning horn.

The system has preflight test capability through the use of aswitch placarded STALL WARN - TEST - OFF on the rightsubpanel. Holding this switch in the TEST position actuatesthe warning horn.

ICE PROTECTION SYSTEMS

WINDSHIELD HEAT

Two levels of heat are provided. When the switches are in theNORMAL (up) position, heat is supplied to the major portionof the windshields. When they are in the HI (down) position, ahigher level of heat is supplied to a smaller area of thewindshields. Each switch must be lifted over a detent beforeit can be moved into the HI position. This lever-lock featureprevents inadvertent selection of the HI position when

October. 1978

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sectionVIISystems Description

moving the switches from NORMAL to the OFF (center)position.

Controllers with temperature-sensing units provide forproper heat at the windshield surfaces. Five-ampere circuitbreakers, located on a panel on the forward pressurebulkhead, protect the control circuits. The power circuit ofeach system is protected by a 40-ampere circuit breaker(50-ampere on 88-35 and after and BL-1 and after, and onearlier serials modified per Service Instructions No. 0712365) located in the power distribution panel under the floorforward of the main spar.

NOTE

Erratic operation of the magnetic compass mayoccur while windshield heat is being used.

PROPELLER ELECTRIC DEICE SYSTEM

I On airplanes 88-2, 88-6 thru 88-815, 88-817 thru 88824, 8L -1 tnru BL-29

The propeller electric deice system includes: an electricallyheated boot with two elements (inner and outer) for eachpropener blade, brush assemblies, slip rings, an ammeter,a timer for automatic operation, and a circuit for manualcontrol for backup.

A 20-ampere circuit breaker switch on the pilot's suopanel,placarded PROP-AUlO-OFF, is provided to activate theautomatic system. A deice ammeter on the right subpanelregisters the amount of current (14 to 18 amperes) passingthrough the system being used. During AUTO operation,

BEECHCRAFTSuper KingAir 200

power to the timer will be cut off if the current rises abovethe circuit breaker switch rating. Current flows from thetimer to the brush assembly and then to the slip ringsinstalled on the spinner backing plate. The slip rings carrythe current to the deice boots on the propeller blades. Heatfrom the boots reduces the grip of ice, which is thenthrown off by centrifugal force aided by the air blast overthe propeller surfaces. Power to the two heating elementson each blade, the inner and outer element, is cycled bythe timer in the following sequence: right propeller outerelements, right propeller inner elements, left propeller outerelements, and left propeller inner elements. Loss of oneheating element circuit on one side does not mean that theentire system must be turned off.

Proper operation can be checked by noting the correctlevel of current usage on the ammeter. An intermittentflicker of the needle approximately every 30 secondsindicates switching to the next group of heating elementsby the timer.

The manual prop deice system is provided as a backup tothe automatic system. A control switch located on the leftsubpanel, placarded PROP-iNNER-OUTER, control themanual override relays. When the switch is in the OUTERposition, the automatic timer is overridden and power issupplied to the outer heating elements of both propellerssimultaneously. The switch is of the momentary type andmust be held in position until the ice has been dislodgedfrom the propeller surface. After deicing with the outerelements, the switch is to be held in the INNER position toperform the same function for the inner elements of bothpropellers. The loadmeters will indicate approximately a -05increase of load per meter when manual prop deice is

DEICE BOOTS

BRUSHBLOCK

FUELCONTROL

PANElPROP DEICE

~CIRCUIT

BREAKERS

200·251·'

DEICE BOOTS

LEFT DEICEMANUAL OVERRIDERELAY AND DIODE

7·48

PROPDEICE TIMER

.d RIGHT DEICEMANUAL OVERRIDERELAY AND DIODE

•• . - ------4--PROPELLER ELECTRIC DEICE SCHEMATIC

February, 1981


operating. The prop deice ammeter will not indicate anyload in the manual mode of operation.

On airplanes 88-816, 88-825 and after, 8L-30 and after:

The propeller electric deice system includes: electricallyheated deicer boots, slip rings and brush block assemblies.a timer for automatic operation, an ammeter, three circuitbreakers located on the fuel control panel for left and rightpropeller and control circuit protection, and two switcheslocated on the pilot's subpanel for automatic or manualcontrol of the system.

A circuit breaker switch located on the pilot's subpanel,placarded PROP-AUTO-OFF, is provided to activate theautomatic system. Upon placing. the switch to the AUTOposition, the timer diverts power through the brush blockand slip ring to all heating elements on one propeller.Subsequently, the timer then diverts power to all heatingelements on the other propeller for the same length oftime. This cycle will continue as long as the switch is in theAUTO position. This system utilizes a metal foil type singleheating element energized by DC voltage. The cycling timefor this system is 90 seconds.

A manual prop deice system is provided as a backup tothe automatic system. A control switch located on the leftsubpanel, placarded PROP-MANUAL-OFF, controls themanual override relay. Upon placing the switch in theMANUAL position. the automatic timer is overridden andpower is then supplied to the heating elements of bothpropellers simultaneously. This switch is of the momentarytype and must be held in position for approximately 45seconds to dislodge ice from the propeller surface. Repeatthis procedure as required to avoid significant buildup ofice which will result in loss of performance, vibration, andimpingement upon the fuselage. The prop deice ammeterwill not indicate a load while the propeller deice system isbeing utilized in the manual mode. However, theloadmeters will indicate an approximate .05 increase ofload per meter while the manual prop deice system isoperating.

SURFACE DEICE SYSTEM

The surface deice system removes ice accumulations fromthe leading edges of the wings and horizontal stabilizers.Ice removal is accomplished by alternately inflating anddeflating the deice boots. Pressure-regulated bleed air fromthe engines supplies pressure to inflate the boots. Aventuri ejector, operated by bleed air, creates vacuum todeflate the boots and hold them down while not in use. Toassure operation of the system in the event of failure ofone engine. a check valve is incorporated in the bleed airline from each engine to prevent loss of pressure throughthe compressor of the inoperative engine. Inflation anddeflation phases are controlled by a distributor valve.

February, 1981

section VIISystems Descriptions

A three-position switch on the pilot's subpanel, placardedDEICE CYCLE w SINGLE - OFF - MANUAL, controls thedeicing operation. The switch is spring-loaded to return tothe OFF position from SINGLE or MANUAL. When theSINGLE position is selected, the distributor valve opens toinflate the wing boots. After an inflation period ofapproximately 6 seconds, an electronic timer switches thedistributor to deflate the wing boots. and a 4-secondinflation begins in the horizontal stabilizer boots. Whenthese boots have inflated and deflated, the cycle iscomplete.

When the switch is held in the MANUAL position, all theboots will inflate simultaneously and remain inflated untilthe switch is released. The switch will return to the OFFposition when released. After the cycle, the boots willremain in the vacuum hold down condition until againactuated by the switch.

For most effective deicing operation, allow at least 1/2 inchof ice to form before attempting ice removal. Very thin icemay crack and cling to the boots instead of shedding.Subsequent cyclings of the boots will then have a tendencyto build up a shell of ice outside the contour of the leadingedge, thus making ice removal efforts ineffective.

PITOT MAST

Heating elements are installed in the pitot masts located onthe nose. Each heating element is controlled by anindividual circuit breaker switch placarded PITOT - LEFT RIGHT, located on the pilot's subpane!. It is not advisableto operate the pitot heat system on the ground except fortesting or for short intervals of time to remove ice or snowfrom the mast.

STALL WARNING VANE

The lift transducer is equipped with anti-icing capability onboth the mounting plate and the vane. The heat iscontrolled by a switch located on the pilot's subpanelplacarded STALL WARN. The level of heat is minimal forground operation. but is automatically increased for flightoperation through the left landing gear safety switch.

WARNING

The heating elements protect the lift transducervane and face plate from ice. However, a build·up of ice on the wing may change or disruptthe airflow and prevent the system fromaccurately indicating an imminent stall.Remember that the stall speed increaseswhenever ice accumulates on any airplane.

7-49

Section VIISystems Description


ENGINE BLEED AIR I. fFIRESEAL~n-r

BLEED AIR FLOWCONTROL UNIT

DEICE BOOT

PRESSURE .-------.VACUUM----

PRESSURE OR VACUUM---

\<~ENGINEBLEED AIR

A FIRESEAL

BLEED AIR FLOWCONTROL UNIT

,-~"=a&.L--t--_ FIREWALL

DEICE BOOT

BRAKE DEICE VALVE

NOTE-AIRPLANES EQUIPPED WITHTHE CARGO DOOR DO NOTHAVE AN INFLATABLE DOORSEAL WITH ITS ATTENDANTVALVE AND PLUMBING

200-193-2

REAR SPAR--- -- ~

BLEED AIRWARNING

SYSTEM

DOOR SEAL SOLENOIDPNEUMATIC VALtVE VALVE CHECK VALVE

TO VACUUM ~O DOOR REGULATOR~ TO PNEUMATIC GAGE

T ~ I DEICESEAL -~ - - =- - --. _=-,MAIN SPAR DISTRIBUTOR

TO FLIGHT HOUR VALVEMETER SWITCH { "! C1~' --":.--. EJECTOR

DEICE LINE DEICE LINE

BLEED AIR LINE ~D AIR LINE~1-1f=~---- CHECK VALVE

BLEED AIRDEICE PRESSURELINE REGULATOR

SURFACE DEICE SYSTEM SCHEMATIC

7-50 February, 1981


FUEL

An oil-to-fuel heat exchanger, located on the engineaccessory case, operates continuously and automatically toheat the fuel sufficiently to prevent ice from collecting inthe fuel control unit.

Each pneumatic fuel control line is protected against ice byan electrically heated jacket. Power is supplied to each fuelcontrol air line jacket heater by two switches actuated bymoving the condition levers in the pedestal out of the fuelcutoff range. Fuel control heat is automatically turned onfor all flight operations.

COMFORT FEATURES

TOILET

The toilet is installed in the foyer and faces the airstairdoor. The foyer can be closed off from the cabin by slidingthe two partition-type door panels to the center of thefuselage, where they are held closed by magnetic strips.The toilet may be either the chemical type or theelectrically flushing type. In either case, the two hinged lidhalf-sections must be raised to gain access to the toilet. Atoilet tissue dispenser is contained in a slide-outcompartment on the forward side of the toilet cabinet.

CAUTION

If a Monogram electrically flushing toilet isinstalled, the sliding knife valve should be openat all times, except when actually servicing theunit. The cabinet below the toilet must beopened in order to gain access to the knifevalve actuator handle.

RELIEF TUBES

A relief tube is contained in a special tilt-out compartmentat the aft side of the toilet cabinet. A relief tube may alsobe installed in the cockpit, and stowed under the pilot orcopilot chair. The hose on the cockpit relief tube is ofsufficient length to permit use by both pilot and copilot.

February, 1981


A valve lever is located on the side of the relief tube horn.This valve lever must be depressed at all times while therelief tube is in use. Each tube drains into the atmospherethrough its own special drain port, which protrudes fromthe bottom of the fuselage. Each drain port is designed toatomize the discharge and keep it away from the skin ofthe airplane.

NOTE

The relief tubes are designed for use duringflight only.

CABIN FEATURES

FIRE EXTINGUISHERS

An optional portable fire extinguisher may be installed onthe 1\oor on the left side of the airplane forward of theairstair entrance door, just aft of the rearmost seat. Anotherone may also be installed underneath the copilot's seat.

WINDSHIELD WIPERS

The dual windshield wiper installation consists of a motor.arm assemblies, drive shafts. and converters, all locatedforward of the instrument panel. The system includes acontrol switch, located in the upper left corner of theoverhead panel. The system circuit breaker is located inthe right subpanel. Windshield wipers may be operated forboth flight and ground operations. Do not use them on dryglass. The control knob, placardedPARK-OFF-SLOW-FAST, controls the wipers. They havetwo speeds. one for light and one for heavy precipitation.After the control is turned to PARK to bring the wipers totheir most inboard position. spring-loading returns thecontrol to the OFF position.

CARGO RESTRAINT (200C)

Beech Aircraft Corporation offers an FAA approved cargorestraint system as Kit No. 101-5040. Any other restraintsystem used in this airplane must be approved by the FAA.Such approval. is the sole responsibility of theowner/operator of the airplane.

7-51

200/200c poh 101-590010-127a13 - beechcraft · rudder boost 7-6 instrument panel 7-6 typical...

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