Super King Air 200/B200 Pilot Training Manualaviaco-va.es/WP/BE20_Technical_Manual.pdf · INTRODUCTION This pilot training manual covers all systems on the Super King Air 200 and

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SUPER KING AIR 200/B200PILOT TRAINING MANUAL

Raytheon Learning CenterFlightSafety International9720 East Central AvenueWichita, KS 67206-2595(316) 612-5300(800) 488-3747

Long Beach Learning CenterFlightSafety InternationalLong Beach Municipal Airport4330 Donald Douglas DriveLong Beach, CA 90808(562) 938- 0100(800) 487-7670

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

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Houston Learning CenterFlightSafety InternationalWilliam P. Hobby Airport7525 Fauna at Airport Blvd.Houston, TX 77061(713) 393-8100(800) 927-1521

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Copyright © 2002 by FlightSafety International, Inc.All rights reserved.

Printed in the United States of America.

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

v

CONTENTS

SYLLABUS

Chapter 1 AIRCRAFT GENERAL

Chapter 2 ELECTRICAL POWER SYSTEMS

Chapter 3 LIGHTING

Chapter 4 MASTER WARNING SYSTEM

Chapter 5 FUEL SYSTEM

Chapter 6 AUXILIARY POWER UNIT

Chapter 7 POWERPLANT

Chapter 8 FIRE PROTECTION

Chapter 9 PNEUMATICS

Chapter 10 ICE AND RAIN PROTECTION

Chapter 11 AIR CONDITIONING

Chapter 12 PRESSURIZATION

Chapter 13 HYDRAULIC POWER SYSTEMS

Chapter 14 LANDING GEAR AND BRAKES

Chapter 15 FLIGHT CONTROLS

Chapter 16 AVIONICS

Chapter 17 MISCELLANEOUS SYSTEMS

Chapter 18 WEIGHT AND BALANCE/PERFORMANCE

GENERAL PILOT INFORMATION

APPENDIX

WALKAROUND

ANNUNCIATOR PANEL

INSTRUMENT PANEL POSTER

FOR TRAINING PURPOSES ONLY


NOTICE

iii

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

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

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

1-i

CHAPTER 1AIRCRAFT GENERAL

CONTENTS

Page

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

AIRPLANE STRUCTURES................................................................................................. 1-16

General........................................................................................................................... 1-16

Fuselage ......................................................................................................................... 1-19

Doors.............................................................................................................................. 1-20


SUPER KING AIR 200/B200 PILOT TRAINING MANUAL


Cabin Windows.............................................................................................................. 1-22

Control Locks ................................................................................................................ 1-23

1-ii FOR TRAINING PURPOSES ONLY



1-iii

ILLUSTRATIONS

Figure Title Page

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

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

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

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

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

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

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

1-8 Engine Control Levers.............................................................................................. 1-7

1-9 Bleed-Air Valve Control........................................................................................... 1-8

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

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

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

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

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

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

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

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

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

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

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

1-21 Pilot’s Static Air Source Valve Handle .................................................................. 1-15

1-22 Cockpit Oxygen Handles ....................................................................................... 1-16

1-23 Airplane Dimensions (BB-1439, 1444 and After) ................................................. 1-17




1-24 Airplane Dimensions (Prior to BB-1444, except 1439)......................................... 1-18

1-25 Fuselage Stations and Compartments .................................................................... 1-19

1-26 Cockpit Layout (Typical) ....................................................................................... 1-20

1-27 Cabin Door ............................................................................................................. 1-20

1-28 Door Handles ......................................................................................................... 1-21

1-29 Placard and Inspection Port.................................................................................... 1-21

1-30 Latch Bolt............................................................................................................... 1-22

1-31 Emergency Exit Release Handles .......................................................................... 1-23

1-32 Control Locks......................................................................................................... 1-24

TABLES

Table Title Page

1-1 Cabin Altitudes....................................................................................................... 1-10

1-iv FOR TRAINING PURPOSES ONLY



INTRODUCTIONThis pilot training manual covers all systems on the Super King Air 200 and B200. Chapter1 provides a general overview of the systems and the structural makeup of the airplane.Throughout this manual there are boxed warnings, cautions, and notes. As indicated inthe Aircraft Flight Manual, they are defined as follows: Warnings—Operating proce-dures, techniques, etc., which could result in personal injury or loss of life if not care-fully followed; Cautions—Operating procedures, techniques, etc., which could resultin damage to equipment if not carefully followed; Note—An operating procedure, tech-nique, etc., which is considered essential to emphasize.

GENERALThe Super King Air 200 and B200 are all metalairplanes employing a fully cantilevered, low-wing design. There are twin Pratt and Whitneyturboprop engines, and a T-tail empennage.

Both airplanes are certificated for flight asNormal Category Aircraft. By carrying requiredoperational equipment, they may be used dur-ing VFR, IFR, and in known icing conditions.

CHAPTER 1AIRCRAFT GENERAL

1-1FOR TRAINING PURPOSES ONLY



AIRPLANE SYSTEMS

ELECTRICAL POWER SYSTEM

GeneralThe airplane electrical system is a 28-VDCsystem, which receives power from a 24-volt,42-ampere hour lead acid gel cell battery(34/36-ampere hour nickel-cadmium batteryprior to BB-1632), two 250-ampere starter-generators, or through an external powersocket.

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

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

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

1. Hot Battery Bus

2. Main Battery Bus

3. Left Generator Bus

4. Right Generator Bus

5. Isolation Bus

1-2 FOR TRAINING PURPOSES ONLY



DUAL FED SUB-BUS #1

DUAL FED SUB-BUS #2

DUAL FED SUB-BUS #3

DUAL FED SUB-BUS #4

STARTRELAY

G C U

VOLT / LOADMETER

R/HGEN LINECONTACTOR

R/H STARTER/GENERATOR

325A

50A 70A 60A

325A

60A 70A 50A

L/

H

GEN

BUS

VOLT / LOADMETER

G C U

L/HGEN LINE

CONTACTOR

L/H STARTER/GENERATOR

STARTRELAY

ISOLATION BUS

MAIN BATT BUSAVIONICS

#1

INVERTER

R/

H

GEN

BUS

AVIONICS

#2

INVERTER

HOT BUS

SHUNT

BATTRELAY

BATTERY

OFF

BATTSWITCH

ON

Figure 1-1. Simplified Electrical System

6. No. 1 Dual Fed Bus




10. The avionics buses

A hot battery bus is powered by the battery,regardless of the position of the BAT switch.This bus supplies the engine fire extinguish-ers, firewall shutoff valves, entry and cargol i gh t s , c l ocks , mod i f i c a t i ons , g roundCOMMunications, RNAV memory to olderavionics, and standby boost pumps prior to BB-1096. It also powers the battery relay which,in turn, allows power through to the main bat-tery bus, provided that the battery switch is ON(Figure 1-2).

The generators are controlled by GEN 1 andGEN 2 switches, located under the same gangbar as the BAT switch. Early King Air air-planes do not have the GEN RESET position.Some airplanes have the reset function, butthey are not placarded. When reset is incor-porated (BB-88 and after), the switch mustbe held in GEN RESET for a minimum of onesecond, and then switched to ON.

The generator buses are interconnected by two325-ampere current limiters on either side ofthe isolation bus. As long as the two isolationlimiters are intact the entire bus system is sup-plied by the battery and the two generators.

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

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

External PowerAn external power socket is located on theunderside of the right wing, outboard of theengine nacelle (Figure 1-3). The airplane willaccept DC power from a ground power unit(GPU) provided the polarity is correct, and theGPU voltage is below 32 volts. The BATswitch must be positioned to ON in airplanesBB-364 and subsequent. Prior to BB-364, theGPU can energize the airplane without thebattery switch on and there is no overvoltageprotection (i.e., more than 32 volts).




Figure 1-2. Electrical Panel

Figure 1-3. External Power Socket

LIGHTING

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

Cabin lighting is controlled by an interior lightswitch on the copilot’s subpanel, labeledBRIGHT–DIM–OFF. (Prior to BB-1444, except1439, it is labeled START/BRIGHT–DIM–OFF)(Figure 1-5). This switch controls the cabin over-head fluorescent lights. Also, individual readinglights at each passenger station can be turned onor off by individual switches adjacent to the lights.

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

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

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




33N3

FO

R0

3060

90120 150 180 210

24027

030

033

0

COMPASS CORRECTIONCALIBRATE WITH

RADIO ON

ST

EE

MAX GEAR EXTENSIONMAX GEAR RETRACTMAX GEAR EXTENDEDMAX APPROACH FLAPMAX FULL DOWN FLAPMAX MANEUVERING

181 KNOTS163 KNOTS181 KNOTS 200 KNOTS 157 KNOTS 181 KNOTS

AIRSPEEDS (IAS)

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

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

CAUTION

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

DO NOT OPERATEON DRY GLASS

WINDSHIELD WIPERSOFF

PARK SLOW

FAST

OFF

MASTERPANELLIGHTS

ON

OVERHEADFLOODLIGHTS

OFFBRT

INSTRUMENTINDIRECTLIGHTS

OFFBRT

AVIONICSPANELLIGHTS

OFFBRT

ENGINEINSTRUMENT

LIGHTS

OFFBRT

PILOTFLIGHTLIGHTS

OFFBRT

OVERHEADSUB PANEL& CONSOLE

LIGHTS

OFFBRT

SIDEPANELLIGHTS

OFFBRT

COPILOT GYROINSTRUMENT

LIGHTS

OFFBRT

COPILOTFLIGHTLIGHTS

OFFBRT

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTSAC VOLTS

FREQ

Beechcraft

0 8020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

-60

0

+60BATT AMPS

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

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

old light, an aisle light, understep lighting, andthe exterior entry light. These three lights turnoff automatically when the airstair door isclosed and the handle is in the LOCK position.

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

MASTER WARNING SYSTEM

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

Annunciator SystemThe warning annunciator panel is located in thecenter glareshield. It contains red indicators,each of which represents a fault requiring thepilot’s immediate attention and action. At thesame time, red MASTER WARNING flasherson the glareshield directly in front of eachpilot begin flashing. The MASTER WARNINGflashers can be extinguished by depressing ei-ther of the lights. The red lights on the warn-ing annunciator panel remain illuminated untilaction is taken to correct the fault.

A caution/advisory annunciator panel is lo-cated on the center subpanel (amber indicatorsfor cautions and green for advisory). An ambercaution illumination requires the pilot’s im-mediate attention to a fault but does not requireimmediate reaction. There are also two amberMASTER CAUTION f l a she r s on t heglareshield, just inboard of the red MASTERWARNING flashers. These operate the sameway as the MASTER WARNING flasher.

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

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

FUEL SYSTEM

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

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

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

When the auxiliary tanks contain fuel, thisfuel is used first and is automatically trans-ferred into the nacelle tank.




Figure 1-6. Exterior Lights ControlSwitches

Each engine drives a high-pressure fuel pumpand a low-pressure boost pump. In addition,an electrically-driven low-pressure standbyboost pump is in the bottom of each nacelletank. The standby boost pump serves threefunctions:

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

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

3. To p u m p f u e l d u r i n g c r o s s f e e do p e r a t i o n .

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

If aviation gasoline is used, a limitation of150 hours of operation per engine before over-hauls must be observed.

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

The fuel quantity is measured by a capaci-tance system, which reads out in pounds on theleft and right fuel gages (Figure 1-7). A switch

between the gages allows the pilot to monitorMAIN or AUXILIARY fuel levels.

POWERPLANTS

GeneralThe Super King Air is powered by two Prattand Whitney turbopropeller PT6A engines,each rated at 850 SHP. They each have a three-stage, axial-flow, single-stage centrifugal flowcompressor (rpm indicated as N1) which isdriven by a single-stage reaction turbine. Thepower turbine is a two-stage reaction turbinecounter rotating with the compressor turbine.A pneumatic fuel control schedules fuel flow.Propeller speed remains constant within thegoverning range for any given propeller con-trol lever position.

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

Engine instruments are grouped at the leftcenter of the instrument panel.




BB-1484, 1486 AND AFTER PRIOR TO BB-1486, EXCEPT BB-1484

Figure 1-7. Fuel Control Panels

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

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

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

3. Condition levers have three positions:

• FUEL CUTOFF

• LOW IDLE

• HIGH IDLE

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

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

Propeller reversing on unimprovedsurfaces should be accomplishedcarefully to prevent propeller ero-sion from reversed airflow and industy or snowy conditions to preventobscured vision.

Condition levers, when set to HIGH IDLE,keep the engine operating at a minimum of70% N1 for quicker reversing response due toless spool up time.

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

CAUTION

CAUTION




BB-1439, 1444 AND AFTER PRIOR TO BB-1444, EXCEPT 1439

Figure 1-8. Engine Control Levers

FIRE PROTECTIONThere are two fire-detection systems. On BB-1439, 1444 and subsequent the system consistsof a temperature sensing cable for each engine.Prior to BB-1444, except 1439, the systemuses three detectors incorporated into eachengine nacelle. Each system has red warningannunciator readouts and a test function. Theoptional engine fire-extinguisher system addsan extinguisher cylinder within each engine na-celle. When the system is installed, glareshieldcontrol switches and additional positions onthe test switch are added (one for each extin-guisher cartridge). There are two portable fireextinguishers installed: one under the copilot’sseat, and the other near the entrance door.

BLEED-AIR SYSTEM

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

The bleed air used for the pneumatic systemis tapped off prior to the flow control unit andis routed through a shutoff valve to a regula-tor. This pneumatic air is then used for surfacedeice, rudder boost, door seal, bleed-air warn-ing system, the flight hour meter, brake deice(if installed) and the landing gear hydraulicreservoir ( if installed). Through the use of aventuri, vacuum suction is developed for flightinstruments, pressurization controller opera-tion and deice boots. One engine can supplysufficient bleed air for all associated systems.

Bleed-Air WarningThe pressurization and pneumatic bleed-airlines have follower plastic tubing containing“regulated (18 psi)” bleed-air. If a bleed-airline ruptures, the released heat will melt theaccompanying plastic tube and the loss ofpressure will cause the respective red L or R

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

ICE AND RAIN PROTECTION

Ice ProtectionIce protection is accomplished either pneu-matically or electrically. Pneumatic ice pro-tection uses engine bleed air for surfacedeicing of wing and horizontal stabilizer lead-ing edges , and ho t b rakes , i f ins ta l l ed .Electrical heating elements are used for wind-shield heating, fuel vent heat, propeller deic-ing, pitot mast heat, and stall warning vane heat(Figure 1-10).

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




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

Figure 1-9. Bleed-Air Valve Control

are used which are moved into the airstream.These cause a slight deflection in the enteringairflow, introducing a turn in the airstream. Theaccelerated moisture particles continue on to thedischarge port, rather than entering the engine.On BB-1439, 1444 and subsequent a secondelectric actuator is employed as a backup. Priorto BB-1444 except 1439, if the electric ice vanecontrols do not work, mechanical extensionhandles may be used. Operation of the vanes aredisplayed either by green L or R ENG ANTI-ICE advisory lights (normal operation) or byamber L or R ENG ICE FAIL caution lights, in-dicating a possible malfunction. (Prior to BB-1444 except 1439, these annunciators are labeledL or R ICE VANE EXT and L or R ICE VANE,respectively.)

An optional brake deice system allows a flowof hot bleed air to the brakes. If installed, op-eration is controlled by a switch on the ICEpanel (Figure 1-10) and indicated by a greenBRAKE DEICE ON advisory annunciatorlight.

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

AIR CONDITIONING ANDHEATING

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

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

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

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






PARK SLOW

FAST

Figure 1-11. Windshield Wiper ControlSwitch

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

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

Electric Heating (BB-1439, 1444and Subsequent)An optional electric heating system is avail-able for ground operation only. A ground powerunit must be used prior to engine starting orgenerator power after engine starting in orderto use electric heating system. It is for groundoperation only and is used in conjunction witheither manual heat or automatic temperaturecontrol mode. A green advisory light on the an-nunciator panel is provided to indicate poweris being supplied to the unit. Both the ventblower and aft blower must be operating whenusing the electric heater.

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

PRESSURIZATION

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

Bleed air from the engine compressor sectionis used to supply airplane pressurization.Engine bleed is mixed with ambient air toform a suitable mixture. The flow control unitand BLEED AIR VALVE switches, as seen inFigure 1-9, control the mixture. If this switchis positioned to ENVIRonmental OFF orINSTrument and ENVIRonmental OFF, thebleed-air valve will be closed. When posi-tioned to OPEN, air is routed through a heatexchanger and then into a mixing plenum. Itmixes with recirculated air, is routed to the out-let ducts, and is introduced into the cabin.




FLIGHT ALTITUDE CABIN ALTITUDE

ALTITUDES ARE IN FEET

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

20,000 3,900 2,800

31,000 9,900 8,600

35,000 11,700 10,400

Table 1-1. CABIN ALTITUDES

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

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

1. To provide positive pressure relief ifthe outflow valve malfunctions.

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

3. To maintain an unpressurized state whileon the ground with the left landing gearsafety switch compressed.

4. To prevent negative differential.

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

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

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

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

LANDING GEAR AND BRAKES

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

Individual gear position is indicated by threegreen lights adjacent to the handle. The gear han-dle contains two red lights, which illuminatewhen the gear is in transit or not properly locked.Two versions of the control panel are found inFigure 1-13. On airplanes with the hydrauli-cally-actuated gear, the square light assemblyhas green NOSE, L, and R indicator segments.




Figure 1-12. Cabin PressurizationController

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

Manual Extension (ElectricGear)The landing gear can be manually extended bypulling the LANDING GEAR RELAY circuitbreaker and placing the landing gear switchhandle in the DN position. Pulling up andturning the emergency engage handle (Figure1-14) positions an emergency drive gear tothe gearbox. A continuous-action ratchet isthen pumped to lower the gear. The system maybe reverted to electrical operation by reposi-




HYDRAULIC GEAR ELECTRIC GEAR

Figure 1-14. Manual Extension Controls

PRIOR TO BB-453 (SUBSEQUENT MODELS HAD THE GEAR DOWN INDICATOR LIGHTS IN A CUBE ARRANGEMENT)

BB-1439, 1444 AND AFTER

Figure 1-13. Landing Gear Control Panel

tioning both handles on the floor and resettingthe circuit breaker.

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

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

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

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

FLIGHT CONTROLS

GeneralThe airplane uses conventional ailerons andrudders. There is a T-tail horizontal stabilizerand elevator mounted at the top of the verti-

cal stabilizer. Interconnected conventionalcontrol columns within the cockpit controlthe ailerons and elevators. Rudder pedals arealso connected so that either the pilot or copi-lot can operate the rudder. There are dual flapson each wing. Rudder, elevator, and ailerontrim are adjustable with controls mounted onthe center pedestal. The flight control sur-faces are illustrated in Figure 1-16.

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

Rudder, elevator, and aileron trims are ad-justable with controls on the center pedestal.Elevator trim is manual or optionally electri-cal. There is a position indicator on eachpedestal tab control (Figure 1-17).

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

PITOT AND STATIC SYSTEMS

GeneralCertain flight instruments operate from impact(pitot) and static pressures.




Figure 1-15. Parking Brake Handle




Figure 1-17. Trim Tab Controls andIndicators

Figure 1-18. Flap Control Lever

ELEVATORS

TRIM TABS

RUDDER

TRIM TAB

AILERON

TRIM TAB

FLAPS

FLAPS

GROUND ADJUSTABLE TAB

AILERON

Figure 1-16. Flight Control Surfaces

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

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

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

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

OXYGEN SYSTEM

GeneralThe airplane’s oxygen system is based on an ad-equate flow for the altitude to which the airplaneis certificated: 31,000 feet or 35,000 feet.

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

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

WARNING




Figure 1-19. Pitot Tubes

Figure 1-20. Static Ports

Figure 1-21. Pilot’s Static Air Source Valve Handle

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

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

Oxygen supply is controlled by a push-pullhandle, placarded PULL ON-SYStem READYand is located on the left side of the pedestal(Figure 1-22). (Prior to BB-1444, except 1439they are overhead in the cockpit — Figure 1-22). When pushed in, no oxygen is availableanywhere in the airplane. It should be pulledout prior to engine start to ensure availableoxygen when needed. The primary oxygensystem delivers oxygen to the two crew masks,to the first-aid outlet in the toilet area, and tothe passenger oxygen system shutoff valve.

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

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

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

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

AIRPLANESTRUCTURESGENERALThe Super King Air is 43 feet 9 inches longfrom the nose to the aft most point of the hor-izontal stabilizer (Figures 1-23 and 1-24). Theairplane sections consist of the:

• Fuselage

• Wings

• Empennage




Figure 1-22. Cockpit Oxygen Handles





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

15' 0"

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

17.25" (1), (3)16.79" (2), (3)16.75" (1), (4)16.29" (2), (4)

93.0" DIA (3)94.0" DIA (4)

32.1" (3)31.6" (4)

CONFIGURATION:(1) STANDARD LANDING GEAR(2) HIGH FLOTATION LANDING GEAR(3) HARTZELL PROPELLER(4) McCAULEY PROPELLER

18' 5"

WING AREA: 303.0 SQUARE FEET

54' 6"

17' 2"

Figure 1-23. Airplane Dimensions (BB-1439, 1444 and After)




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

29.60" (3)29.85" (4)

WING AREA303.0 SQUARE FEET

18' 5"

54' 6"

17' 2"

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

CONFIGURATIONS:(1) STANDARD LANDING GEAR(2) HIGH FLOTATION LANDING GEAR(3) HARTZELL PROPELLER(4) MCCAULEY PROPELLER

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

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

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

Figure 1-24. Airplane Dimensions (Prior to BB-1444, except 1439)

The fuselage is composed of the:

• Nose section

• Cockpit

• Cabin

• Foyer and aft cabin

• Aft fuselage

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

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

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

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

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

equipped with shoulder harnesses have ad-justable headrests.

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

Couches, if installed, are not adjustable.

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

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

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

CAUTION

CAUTION




CREW

ROW 1 ROW 2 ROW 3

AFTCABIN

FOYERCABINCOCKPIT

Figure 1-25. Fuselage Stations and Compartments

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

DOORS

Cabin DoorThe cabin door is located on the left side ofthe fuselage, in the foyer area. The cabin dooris hinged at the bottom, and swings out anddown when opened (Figure 1-27). A hydraulicdamper ensures a slow opening.

A stairway is built onto the inboard side forentry and egress. Two of the steps fold flatagainst the door when it is closed. When thedoor is fully extended, it is supported by aplastic-encased cable, which also serves as ahandrail. A second handrail (optional prior to

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




Figure 1-26. Cockpit Layout (Typical)

Figure 1-27. Cabin Door

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

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

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

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

cabin has been depressurized, a crewmembercan then check the door security.

When closing the door from inside the air-plane, pull up on the handrail until the airstairdoor reaches the door frame. Rotate the doorhandle up as far as possible, pulling inward onthe door. The door should seal; then rotate thehandle down to lock the door (Figure 1-28).Positive locking may be checked by attempt-ing to rotate the handle without depressingthe release button. It should not move. A plac-ard is located beneath the folded step justbelow the door handle. The placard showshow to check the locks in the inspection portwindows near each corner of the door (Figure1-29). A green stripe painted on each of thefour latch bolts should be aligned with its re-spective black pointer (Figure 1-30).

CAUTION




INSIDE DOOR HANDLEOUTSIDE DOOR HANDLE

Figure 1-28. Door Handles

Figure 1-29. Placard and Inspection Port

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

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

Once the latches are retracted, initial pres-sure must be exerted outward to start the open-ing action. After the sequence begins, gassprings will open the door the rest of the way.The door is counterbalanced, and will stayopen. The gas springs will resist the effort toclose the door, and that pressure must be over-

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

The door closes against a rubber seal, to main-tain the pressure vessel integrity. The seal isnot inflated by pneumatic bleed air, but ratherallows cabin-pressurized air to seep into holeson the inside. This allows for greater sealingwhen there is a high pressure differential.

Emergency ExitThe emergency exit window, placarded EXIT-PULL (Figure 1-31) is located at the forwardright side of the passenger compartment. Itcan be released from the inside by using apull-down handle, or from the exterior (if itis unlocked) by a flush-mounted, pull-outhandle (Figure 1-31). It is a plug-type exit,which is removed completely from the frameand taken into the cabin. The exit can belocked from the inside, but can be openedfrom the inside even when it is locked. ForBB-415 and after, the locking mechanism isactivated by pulling out a handle below thedoor release handle (Figure 1-31). Prior air-craft and BL-1 and after have a key next tothe door release handle that can lock/unlockthe door. This key cannot be removed whenthe door is locked.

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

CABIN WINDOWSEach cabin windowpane is composed of asheet of polyvinyl butyral between two trans-parent sheets of acrylic plastic. It is stressedto withstand the cabin pressure differential.There are two types of windowpanes available:polarized and shade type.




Figure 1-30. Latch Bolt

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

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

When the airplane is to be parked inareas exposed to intensive sunlight,the polarized windows should be ro-tated to the clear position to preventdeterioration of the polarized mate-rial. Sufficient ultraviolet protectionis provided to prevent fading of theupholstery.

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

CONTROL LOCKSThe flight and engine controls are mechani-cally locked by a U-shaped clamp and two pinswithin the cockpit, as seen in Figure 1-32. Thepins lock the primary flight controls and the U-shaped clamp fits around the engine controllevers. A pin is inserted through the controlcolumn to lock the ailerons and elevator. A sec-ond pin is inserted through a hole in the floor,which locks the rudder bellcrank. All locksmust be installed and removed together to pre-clude taxiing or flying with the engine controllevers released but the flight controls locked.

CAUTION

WARNING




Figure 1-31. Emergency Exit Release Handles

Before starting engines, remove thelocks.

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

CAUTION

WARNING




Figure 1-32. Control Locks

2-i

CHAPTER 2ELECTRICAL POWER SYSTEMS

CONTENTS

Page

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Inverters ......................................................................................................................... 2-22

Circuit Breakers ............................................................................................................. 2-22

QUESTIONS......................................................................................................................... 2-29




2-iii

ILLUSTRATIONSFigure Title Page

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

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

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

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

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

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

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

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

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

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

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

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

2-13 Electrical Distribution .............................................................................................. 2-9

2-14 Circuit-Breaker Panels—Pilot’s ............................................................................. 2-10

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

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

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

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

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

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

2-21 Electrical System—Super King Air B200 (BB-1484, 1486 and Subsequent;BW-1 and Subsequent)........................................................................................... 2-17

2-22 Electrical System—Super King Air B200 (BB-1449, 1458-1462, 1464-1485,Except 1484; BL-139, 140).................................................................................... 2-18

2-23 Electrical System—Super King Air B200 (BB-1439, 1444-1448, 1450-1457) .... 2-19




2-24 Electrical System—Super King Air B200 (BB-734, 793, 829, 854-870,874-891, 894, 896-911, 913-1438, 1440-1443, BL 37-138).................................. 2-20

2-25 Electrical System—Super King Air 200 (B-2, 6-733, 735-792, 794-828, 830-853871-873, 892, 893, 895, 912, BL-1-36) ................................................................. 2-21

TABLES

Table Title Page

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

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

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

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




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

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

GENERALThe DC power is supplied by a 24-volt batteryand by two 30-volt, regulated to 28.25 ± .25volts, 250-ampere starter-generators. Eitherone of two inverters supplies AC power for en-gine instruments and for avionics (Figure 2-1).

Each component of the electrical power systemis capable of supplying power to all systems thatare necessary for normal operation of the air-plane; however, the battery, if it is the onlysource of power, does have a limited life.

#1 S

ERVO

SYSTEM

BATT HOT

BAT OFF

AC

GEN

#1 D

C

GEN

#1 E

NG

OIL PL

CHAPTER 2ELECTRICAL POWER SYSTEMS




DC POWER

BATTERY

For BB-1632 and subsequent, a single, 24-volt,42 ampere-hour sealed lead acid gel cell batteryis located in the right wing center section for-ward of the main spar. Prior to BB-1632, a sin-gle 24-volt, 34/36 ampere-hour nickel-cadmium(NiCad) battery is installed. This NiCad batteryrequires air cooling through a thermostaticallycontrolled valve installed in the ram air tube ad-jacent to the battery drain (Figure 2-2).

A hot battery bus (Figure 2-3) is provided foroperation of essential equipment and the cabinthreshold light circuit when the battery and

generators are not on. Power to the main busfrom the battery is routed via the battery relay,which is controlled by the BAT ON–OFFswitch on the pilot’s left subpanel.

For aircraft BB-1632 and subsequent, the bat-tery ammeter (Figure 2-4) provides a directreading of the charge or discharge rate of thebattery (–60 amps to +60 amps). The chargerate should be 0 to +10 amperes for take-off.

On aircraft prior to BB-1632 with a NiCadbattery, a battery charge current detector isinstalled. This senses an increase in normalcurrent flow and causes an amber BATTERYCHG caut ion annuncia tor to i l luminate(Figure 2-5), alerting the flight crew that thebat tery charge current is above normal .




STARTER–GENERATOR

INVERTER

INVERTER

BATTERY

EXTERNALPOWER

CONNECTORSTARTER–

GENERATOR

PRINTEDCIRCUIT BOARDS

Figure 2-1. Electrical Component Location




Figure 2-2. Battery Cooling (Nickel Cadium)

ISOLATION

BUS

MAIN

BATTERY

BUS

BATTERYRELAY

BATTERYSWITCH

HOT

BATTERY

BUS

BATTERY

S

H

U

N

T

TOBATTERYCHARGESENSOR

Figure 2-3. Battery Control Circuit

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

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

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

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

Starter power to each starter-generator is pro-vided from the main battery bus through astarter relay. The start cycle is controlled bya three-position switch for each engine la-beled IGNITION AND ENGINE START.When placed to the ON (up) position, theswitch becomes mechanically locked and mustbe pulled out to reposition. When held to thedown position, labeled STARTER ONLY, theassociated engine will motor, but ignition willnot occur. When released, the spring-loadedswitch will move to the center position, whichis labeled OFF.




Figure 2-5. BATTERY CHG AnnunciatorFigure 2-6. Generator

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTSAC VOLTS

FREQ

Beechcraft

-60

0

+60BATT AMPS

Figure 2-4. Volt-Loadmeters-Battery Ammeter




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

The generator operation is controlled by indi-vidual generator switches located on the pilot’sleft subpanel under the MASTER SWITCHgang bar with the BAT switch. As shown inFigure 2-7, the switches are labeled GEN 1and GEN 2. In order to turn the generator on,the control switch must be held upward in theGEN RESET position (Figure 2-7) for a min-imum of one second, then released to the ONposi t ion. (Prior to BB-88, the generatorswitches do not have the reset position.)

Figure 2-8 shows that power to the bus systemfrom the generators is protected by GeneratorControl Units (GCU). For BB-88 and after, theGCU operates a line contactor relay to protectthe generator. Prior to BB-88, reverse-currentprotection is provided by a unit in line with thegenerator output.The generators are controlled by individualgenerator control units, which maintain aconstant voltage during variations in enginespeed and electrical load requirements. Thevoltage regulating circuit will automaticallyconnect or disconnect a generator’s output tothe bus. The load on each generator is indi-cated by the respective left and right volt-

loadmeter (Figure 2-8) on the overhead panelwhich reads in percent of the generator ’smaximum continuous capacity. Normally,this value is 250 amps; therefore, a loadme-ter reading of .5, or 50%, is equal to 125amps of generator output.

NOTEThe generators will drop off the lineif underexcitation, overexcitation,overvoltage, or undervoltage condi-tions exist.

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

1. The ground power source polarity iscorrect.

2. The BAT SWITCH is on.

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

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

For starting, an external power source capa-ble of supplying up to 1,000 amperes (300amperes maximum continuous) should beused. A caution light on the caution advisoryannunciator panel labeled EXT PWR is pro-vided to alert the operator when a groundpower plug is connected to the airplane. Someearlier airplanes used a switch to sense powerplug connection, and later airplanes incorpo-rated an electronic circuit utilizing the smallpin of the plug (Figure 2-10).

Figure 2-7. Generator Switches




ISO

LAT

ION

BU

S

MA

IN B

AT

TE

RY

BU

S

VOLTLOAD

METER

SHUNT

ISOLATION LIMITER

ISOLATIONLIMITER

RIGHTSTARTRELAY

BATTERYSWITCH

L GEN LINECONTACTOR

HOTBATTERY

BUS

LEFT GEN CONTROL

LEFTSTARTER

GEN

RIGHT GEN BUS

BB-88 AND AFTER

VOLTLOAD

METER

SHUNT

RIGHT GEN CONTROL

RIGHTSTARTER

GEN

R GEN LINECONTACTOR

OFF

BATTERY RELAY

BATTERYRELAY

SHUNT

BATTERYCHARGEMONITOR

BATTERY

LEFTSTARTRELAY

ISO

LAT

ION

BU

S

MA

IN B

AT

TE

RY

BU

S

BATTERY

LEFT GEN CONTROL

VOLTLOAD

METER

SHUNT

ISOLATION LIMITER

LEFTSTARTER

GEN

VOLTLOAD

METER

SHUNT

ISOLATION LIMITER

RIGHTSTARTRELAY


BATTERYSWITCH

LEFTSTARTRELAY

HOTBATTERY

BUS

BATTERYRELAY

SH

UN

T

LEFT GEN BUS

REVERSECURRENT

PROTECTION

RIGHT GEN CONTROL

RIGHTSTARTER

GEN

REVERSECURRENT

PROTECTION

RIGHT GEN BUS

PRIOR TO BB-88

LEFT GEN BUS

BATTERY RELAY

Figure 2-8. Generator Control Circuit

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

must be recharged, or replaced witha battery indicating at least 20 volts,before connecting ground power.

Observe the following precautions when usinga ground power source:

1. Use only a ground power source that isnegatively grounded. If polarity of thepower source is unknown, determinethe polarity with a voltmeter before con-necting the unit to the airplane.

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

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

CAUTIONCAUTION




Figure 2-9. Ground Power Connector

ISOLATIONLIMITER

EXT POWERCONNECTOR

HOT BATTERYBUS

BATTERYRELAY

EXTERNALPOWERM

AIN

BA

TT

ER

Y B

US

ISO

LAT

ION

BU

S

EXT POWERRELAY

BATTERY

SHUNT


BATTERYSWITCH

OFF

ON

BATTERYRELAY

BB-364 AND AFTERBB-88 TO BB-363PRIOR TO BB-88

ISOLATIONLIMITER

EXT POWERCONNECTOR

HOT BATTERYBUS

EXT POWERRELAY

MA

IN B

AT

TE

RY

BU

S

ISO

LAT

ION

BU

S

BATTERYRELAY

BATTERYSWITCH

BATTERYRELAY

BATTERY


TO ANNUNCIATORADVISORY LIGHT

EXTERNAL POWERPLUG ENGAGED

SENSOR

SH

UN

T

EXT POWERSENSEISOLATION

LIMITER

EXT POWERCONNECTOR

HOT BATTERYBUS

EXT POWERRELAY

MA

IN B

AT

TE

RY

BU

S

ISO

LAT

ION

BU

S

BATTERYRELAY

BATTERYSWITCH

BATTERYRELAY

BATTERY


EXTERNAL POWERPLUG ENGAGED

SENSOR

SH

UN

T

Figure 2-10. External Power Circuit

first applying battery voltage. If thebattery is removed from the airplaneor if the battery switch is to be placedin the OFF position, turn each indi-v idual radio and o ther avionicsequipment off.

3. After the external power plug is con-nected and power is applied, leave thebattery on during the entire groundpower operation to protect transistor-ized equipment against transient volt-age spikes.

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

Only use a ground power source fitted with anAN2552-type plug. If uncertain of the polar-ity, check it with a voltmeter to ensure that itis a negative-ground plug. Connect the posi-tive lead to the larger center post of the re-ceptacle, and connect the negative-groundlead to the remaining large post. The small postis the polarizing pin; it must have a positivevoltage applied to it in order for the externalpower relay to close.

CONTROLS AND INDICATORSElectrical control switches are convenientlylocated on the pilot’s left subpanel (Figure 2-11). The battery switch and the two generatorswitches are positioned under a hinged flap la-beled MASTER SWITCH, commonly referredto as the gang bar. When this flap is depressed,the battery and both generators are switched off.

Electrical component indication is throughlights on the annunciator panel or meters onthe overhead panel (Figure 2-12).

When a generator is off the line, the respec-tive amber L or R DC GEN caution annunci-ator illuminates. There are also optional redGEN OVHT warning lights to warn of a gen-erator overheat condition on B200 airplanes.

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

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

DISTRIBUTIONThe battery is connected to a hot battery bus(Figure 2-13) which powers threshold lights,the fire extinguishing system, firewall shut-off valves, the battery relay, ground com-munications, auxiliary DC bus (if installed),external power light (BB-88 and after), RNAVmemory, stereo, and prior to BB-1098 (ex-cluding BB-1096), standby boost pumps.With the battery switch on, power is fed tothe main battery bus, which is connectedthrough the start relays to both starter-gen-erators. The main battery bus feeds the iso-lat ion bus and, through two 325-ampere

CAUTION




Figure 2-11. MASTER SWITCHES

L DC GEN R DC GEN

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

Figure 2-12. Lights and Meters

isolation limiters (current limiters), connectsthe left and right generator buses together.

When the battery, generators, or GPU are pro-viding power, the isolation bus, L generatorbus, and R generator bus function as one unit,as long as both current limiters are not open.There are four subbuses fed by both the left andright generator buses. They are labeled No. 1through No. 4 DUAL FED BUS. Each subbusis fed from either side through a 60-ampere cur-

rent limiter, a 70-ampere reverse current diode,and a 50-ampere circuit breaker which is ac-cessible to the crew. There are eight of these 50-amp feeder breakers. Four are located on thecopilot’s side panel for the No. 1 and No. 2 sub-buses, and on the fuel panel circuit breaker busfor the No. 3 and No. 4 subbuses. Of those itemswith paired circuits such as the left and rightlanding lights, the distribution will be such thatthe left circuit is on the No. 1 or No. 3 dual fedbus and the right is on the No. 2 or No. 4.




DUAL FED SUB-BUS #1

DUAL FED SUB-BUS #2

DUAL FED SUB-BUS #3

DUAL FED SUB-BUS #4

STARTRELAY

G C U

VOLT / LOADMETER

R/HGEN LINECONTACTOR

R/H STARTER/GENERATOR

325A

50A 70A 60A

325A

60A 70A 50A

L/

H

GEN

BUS

VOLT / LOADMETER

G C U

L/HGEN LINE

CONTACTOR

L/H STARTER/GENERATOR

STARTRELAY

ISOLATION BUS

MAIN BATT BUSAVIONICS

#1

INVERTER

R/

H

GEN

BUS

AVIONICS

#2

INVERTER

HOT BUS

SHUNT

BATTRELAY

BATTERY

OFF

BATTSWITCH

ON

Figure 2-13. Electrical Distribution

Generally, dual fed bus No. 1 and No. 2 runin alternating rows on the copilot’s circuitbreaker panel (excluding the avionics sec-tion). Dual fed bus No. 3 and No. 4 are on thepilot’s circuit-breaker panel.

With BB-1484, 1486 and subsequent, dual-powered engine instruments are also on thepilot’s circuit breaker panel and they are pow-ered by No. 1 dual fed bus (left engine in-struments), No. 2 dual fed bus (right engineinstruments), or by the isolation bus (shouldeither of the subbuses fail).

See Figure 2-14 for the pilot’s circuit-breakerpanel distribution. See Figure 2-15 for thecopilot’s circuit-breaker panel distribution.

OPERATIONThe DC electrical system is activated by turn-ing the battery switch on, then after the enginesare stabilized, turning the generators on.Monitor the generator loadmeters and all elec-trical indicating lights throughout the flight.




OPEN

CLOSED

FIREWALLSHUTOFF

VALVE

LEFT

RIGHT

FUEL SYSTEM CLOSED

ITT PROPTACH

TURBINETACH

FUELFLOW

OILPRESS

OILTEMP

5 5 5 5 5 5 5

5 5 5 5 5 5 5

TORQUE

FIREWALLVALVE

AUXTRANS

FER

AUXTRANS

FER

QTYIND

PRESSWARN

5 10 5 5 5 5

STANDBYPUMP

STANDBYPUMP

CROSSFEED

QTYIND

FIREWALLVALVE

5 5 5 10 5

PRESSWARN

RIGHTLEFT

OPEN

ENGINE INSTRUMENTS

FIREWALLSHUTOFF

VALVE

NO 4

BUSFEEDERS

BUSFEEDERS

NO 3

50

50

RIGHT

PROPDEICE

LEFT

25

25

CONTROL

MOTORPROPDEICE

FLAP

20

5

GOV

CONTROL

PROP

5

5

RIGHT

IGNITORPOWER

LEFT

5

5

RIGHT

STARTCONTROL

LEFT

5

5

NO 4

NO 3

50

50

OPEN

CLOSED

FIREWALLSHUTOFF VALVE

FUEL SYSTEM CIRCUIT BREAKERS

CLOSED

FIREWALLVALVE

AUXTRANS

FER

AUXTRANS

FER

QTYIND

PRESSWARN

STANDBYPUMP

STANDBYPUMP

CROSSFEED

QTYIND

FIREWALLVALVE

5 10 5 5 5 5 5 5 5 10 5

PRESSWARN

RIGHTLEFT

OPEN


BUS FEEDERS

50 50

PROP DEICE

20 20

CONTROL

FLAP

205

GOV

PROP

5

IGNITOR

5 5

CONTROL

5 5

NO 4NO 3

50 50

START

PROP PROP

5

MOTOR CONTROL POWER

BB-1484, 1486 AND SUBSEQUENT

BB-2 — BB-1485, EXCEPT 1484

PARKING BRAKEOFF

COLLINS

RIGHTRIGHTLEFT RIGHTLEFT LEFT

Figure 2-14. Circuit-Breaker Panels—Pilot’s




FURNISHING

WEATHERLIGHTS

ELECTRICALENVIRONMENTAL FLIGHT

WARNINGS

ENGINES AVIONICS

5

PROP

SYNC

5

AUTO

FEATHER

5

AVIONICS

ANN

5

READING

5

OXYGEN

CONTROL

TEMP

5

LEFT

CHPDETR

5

RIGHT

10

NO SMK

FSB &CABIN

5

INSTR

INDIRECT

5

PRESS

CONTROL

5

CONTROL

5

FIRE

DET

5

AVIONICS

& ENGINSTR

SUB PNLOVHD &

CONSOLE

5

LEFT

BLEEDAIR

CONTROL

5

RIGHT

5

LEFT

STBY ENGANTI-ICE

5

RIGHT

71/2

PLT FLT

SIDE PNL

71/2

COPLT FLT

INSTR

/

ALT

ALERT

5

LEFT

MN ENGANTI-ICE

5

RIGHT

5

PILOT

TURN & SLIP

GENCONTROL

BUSFEEDERS

5

COPLT

71/2

LEFT

FUELCONTROL

HEAT

71/2

RIGHT

5

STALL

WARN

3

PILOT

ALTMAIR DATA

|

COPLT

ENCD ALTM

71/2

TCAS

5

LEFT

BLEEDAIR

WARN

5

RIGHT

5

PITCH

TRIM

5

RUDDER

BOOST

2

GPWS

5

WARN

LANDINGGEAR

5

IND

5

OUTSIDE

AIRTEMP

2

GPS

71/2

POWER

ANN

5

IND

3

FMS

5

LEFT

5

FUELVENT

RIGHT

10

LEFT

10

RIGHT

10

MASTER

POWER

2

RADIO

ALTM

71/2

LEFT

71/2

ENGINSTR

POWER

RIGHT

5

SURF

DEICE

10

WSHLD

WIPER

50

5

CIGAR

LIGHTER

5

RADAR

5

BRAKE

DEICE

50

2

VOICE

RCDR

1

AURAL

WARN

AVIONICS

NO 1

30

AVIONICS

NO 2

71/2

LEFT

71/2

MULTIFCTN

PRCSR

RIGHT

3

XPNDR

NO 1

3

XPNDR

NO 2

5

AVIONICS

MASTER

2

PILOT

AUDIO

COMM

NO 1

71/2

COMM

NO 2

5

PILOTEADI

5

PILOTEHSI

2

DME

NO 1

2

DME

NO 2

5

RADIO

PHONE

5

CABIN

AUDIO

NAV

NO 1

2

NAV

NO 2

71/2

DSPL

PRCSR

1

ELEKDSP

CONTR

71/2

NORMAL

EFISFANS

71/2

STBY

2

COPILOT

AUDIO

COMPASS

NO 1

3

30 71/2 2 3

COMPASS

NO 2

2

ADF

4

MULTIFCTN

DSPL

10

AP

SERVO

2

EHSICOPILOT

15

EFISAUX

BAT

2

RMI

NO 2

2

RMI

NO 1

2

FCS

POWER

5

HDG

PRCSR

50 50

NO 1

NO 2

71/2


FURNISHING

WEATHERLIGHTS


WARNINGS

ENGINES AVIONICS

5

NO. 1

5

NO. 2

INVCONTROL

BB-1439, 1444–1485, EXCEPT 1463 AND 1484.

5

PROP

SYNC

5

AUTO

FEATHER

5

AVIONICS

ANN

5

READING

5

OXYGEN

CONTROL

TEMP

71/2

COPLT FLT

INSTR

5

AVIONICS& ENG

INSTR

5

PRESS

CONTROL

5

CONTROL

5

FIRE

DET

5

INSTR

INDIRECT

OVHD &

CONSOLE

5

LEFT

BLEEDAIR

CONTROL

5

RIGHT

5

LEFT

MN ENGANTI-ICE

5

RIGHT

5

PLT FLT

SIDE PNL

71/2

NO SMKFSB &

CABIN

5

PILOT

TURN & SLIP

GENCONTROL

BUSFEEDERS

5

5

STALL

WARN

3

PILOT

ALTMAIR DATA

5

LEFT

BLEEDAIR

WARN

5

RIGHT

5

PITCH

TRIM

5

RUDDER

BOOST

5

WARN

LANDINGGEAR

5

IND

1

ALT

ALERT

71/2

POWER

ANN

5

IND

5

LEFT

5

FUELVENT

RIGHT

10

LEFT

10

RIGHT

10

MASTER

POWER

5

LEFT

5

FUELFLOW

RIGHT

5

SURF

DEICE

10

WSHLD

WIPER

50

5

CIGAR

LIGHTER

50

10

AURAL

AVIONICS

NO 1

30

AVIONICS

NO 2

71/2

RMI

71/2

3

3

RADAR RADIO

5

AVIONICS

MASTER

5

PILOT

COMM

NO 1

71/2

COMM

NO 2

5

DME

5

2

2

ADF

ALTM

2

CABIN

AUDIO

NAV

NO 1

2

NAV

NO 2

71/2

COMPASS

1

10

AP

SERVO

XPNDR

2

30 71/2 2 2

2

FCS

POWER

2

EHSI HDG

PRCSR

71/2

LEFT

FUELCONTROL

HEAT

71/2

RIGHT

50 50

NO 1

NO 2

71/2

5

LEFT

STBY ENGANTI-ICE

5

RIGHT

5

LEFT

CHIPDETR

5

RIGHT

COPLT

5

COPILOT

AUDIOAUDIOWARN

NO 1

XPNDR

NO 2

NO 1 NO 1NO 2

RMI DME COMPASS

NO 2 NO 2NO 1

5

5

COPLT

ENCD ALTM

OUTSIDE

AIR TEMP

5

LEFT

5

TORQUEMETER

RIGHT

5

LEFT

5

OILPRESS

RIGHT

5

LEFT

5

OILTEMP

RIGHT

Figure 2-15. Circuit-Breaker Panels—Copilot’s (1 of 2)

If the pilot suspects a NiCad battery malfunc-tion, he should refer to the Battery ConditionCheck procedures, in the Normal Proceduressection of the Aircraft Flight Manual.

AVIONICS MASTER SWITCHThe avionics power relays are normally closedand supply power to the buses. Note that therelays require DC power to open and discon-nect the avionics buses (Figure 2-16).

Typical avionics bus distribution for an EFISequipped aircraft is shown in Figure 2-17.

AC POWER INVERTERSEither one of two inverters (Figure 2-18) pro-vides the AC power. The inverters are installedin the wing center section outboard of each na-celle. Each inverter provides both 115-voltand 26-volt, 400-Hertz power to be used foravionics equipment and engine instruments.

CONTROLS AND INDICATORSIf the airplane has an AC VOLT-FREQ meter,inverter output can be monitored. The meternormally reads frequency (Figure 2-19) butwill display volts when the button is depressed.




FURNISHING

WEATHERLIGHTS


WARNINGS

ENGINES

5

PROP

SYNC

5

AUTO

FEATHER

5

AVIONICS

ANN

5

READING

5

OXYGEN

CONTROL

TEMP

71/2

LEFT

FUELCONTROL

HEAT

71/2

RIGHT

71/2

FLIGHT

INSTR

5

AVIONICS& ENG

5

PRESS

CONTROL

5

CONTROL

5

FIRE

DET

5

INSTR

INDIRECT

OVHD &

CONSOLE

5

LEFT

BLEEDAIR

CONTROL

5

RIGHT

5

LEFT

OILTEMP

5

RIGHT

5

SIDE

PANEL

71/2

NO SMKFSB &

CABIN

5

LEFT

OILPRESS

5

RIGHT

5

PILOT

TURN & SLIP GEN

CONTROLBUS

FEEDERS

5

ENCO AL TM

2

LEFT

TORQUEMETER

2

RIGHT

5

STALL

WARN

3

PILOT

ALTMAIR DATA

5

LEFT

BLEEDAIR

WARN

5

RIGHT

5

PITCH

TRIM

5

RUDDER

BOOST

5

WARN

LANDINGGEAR

5

IND

1

ALT

ALERT

71/2

POWER

ANN

5

IND

5

LEFT

5

FUELVENT

RIGHT

10

LEFT

10

RIGHT

10

MASTER

POWER

5

LEFT

5

FUELFLOW

RIGHT

5

SURF

DEICE

10

WSHLD

WIPER

50

5

CIGAR

LIGHTER

5

BRAKE

DEICE

50

50 50

NO 1

NO 2

71/2

5

LEFT

ICEVANE

CONTROL

5

RIGHT

5

LEFT

CHPDETR

5

RIGHT

5

LEFT

5

RIGHT

INVCONTROL

COPILOT

B-2 THROUGH BB-1443, EXCEPT 1439

INSTR

AVIONICS

10

AURAL

AVIONICS

NO 1

30

AVIONICS

NO 2

71/2

RMI

71/2

3

3

RADAR RADIO

5

AVIONICS

MASTER

5

PILOT

COMM

NO 1

71/2

COMM

NO 2

5

DME

5

2

2

ADF

ALTM

2

CABIN

AUDIO

NAV

NO 1

2

NAV

NO 2

71/2

COMPASS

1

10

AP

SERVO

XPNDR

2

30 71/2 2 2

2

FCS

POWER

2

EHSI HDG

PRCSR

5

COPILOT

AUDIOAUDIOWARN

NO 1

XPNDR

NO 2

NO 1 NO 1NO 2

RMI DME COMPASS

NO 2 NO 2NO 1

Figure 2-15. Circuit-Breaker Panels—Copilot’s (2 of 2)




AVIONICS MASTER POWER CB

AVIONICS MASTER POWER SWITCH

5ANUMBER 1

DUAL FED BUS

ON

OFF

RIGHTGENERATOR

BUS

40A

30A

NUMBER 2AVIONICSBUS

NUMBER 1AVIONICSBUS

30A

40A

LEFTGENERATOR

BUS

Figure 2-16. Avionic Power Distribution




AVIONICSMASTER

RIGHT MULTI FCTN PRCSR

COMM NO 1

NAVNO 1

COMPASS NO 1

ADF NO 1

RMI NO 2

XPNDR NO 1

DME NO 1

EFIS FANS NORMAL

DTU

STEREO

FMS

RADIO ALTM

AP SERVO

FCS POWERPILOT TURN & SLIP

PILOT ALTM & AIR DATA

PITCH TRIM

OUTSIDE AIR TEMP

CVR

AURAL WARN

PILOT AUDIO

ALT ALERT

LEFT MULTI FCTN PRCSR

PILOT EADI

DSPL PRCSR

COPLT TURN & SLIP

COPLT ENCD ALTM

CABIN AUDIO

COPILOT AUDIO

PILOT EHSI

ELEK DSP

COMM NO 2

NAVNO 2

COMPASS NO 2

EFIS AUX BAT

ADF NO 2

RADAR

MULTI FCTN DSPL

RMI NO 1

XPNDR NO 2

DME NO 2

EFIS FANS STBY

COPILOT EHSI

HDG PRCSR

OFF

ON

AVIONICSMASTERSWITCH

LIMITER

40A

30AAVIONICS

NO 1

AVIONICSBUS NO 1RELAY

AVIONICSBUS NO 1

AVIONICSBUS NO 2

AVONICSBUS NO 2RELAY

30AAVIONICS

NO 2

ISOLATION BUS& BATTERY BUS*

R/HGENERATORBUS*

NO 1 DUAL FEDELECTRICALBUS*

NO 2 DUAL FEDELECTRICALBUS*

40A

LIMITER

* NOTE NOT ALL OFBUS SHOWN ONLYAVIONIC ITEMS CONDITION 1. MAIN AIRCRAFT POWER ON

AVIONICS MASTER SWITCH SELECTED OFF

L/HGENERATORBUS*

AVIONICS ANN

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

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

In the event the inverter fails, a red INVERTER(INST INV on 200 models) light on the warn-ing annunciator panel will illuminate.

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

OPERATIONTurn the INVERTER select switch to either in-verter position, note that the INVERTER(INST INV on 200 models) warning light ex-tinguishes, and then monitor the VOLT-FREQmeter.




Figure 2-18. Inverters

Figure 2-19. Volt-Frequency Meter

2-16FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

No.1 INVERTER ON LINE

AC VOLT/FREQMETER

POWERRELAY

LIMITER

LIMITERPOWER

RELAY

DC GROUND

DC GROUND

DC POWER

DC POWER

115 VAC

115 VAC

26 VAC

26 VAC

AC COMMON

AC COMMON

5A

NO 1

OFF

NO 2INVERTERSELECTSWITCH

5A

115 VACSELECTRELAY

26 VACSELECTRELAY

L/H GENBUS

R/H GENBUS

INVERTER NO 1115 VACBUS

AC COMBUS

26 VACBUS

5A

5A

AC TESTJACK (BLUE)

AC POWERRETURNS

FROMSYSTEMS

1A

1A

1A

2A

1A

2A

1A

1A

1A

1A

50A

50A

1A

1A

B200 AC Power System

5AANN IND

AVIONICS JUNCTION BOX

INVERTER No 2

No. 2 INVERTERCONTROL

No. 1INVERTERCONTROL

28VDC

5A

10A

10A

5A

INVERTERSELECTRELAY

INVERTERWARNING

RELAY

VG POWER & REF TO OTHER SYSTEMS

RADAR REF FROM VG FOR STABILIZATION

AP REF FROM VG

AP YAW RATE GYRO POWER

COMPASS 2 REF TO RMI NO. 1 & MPU

ADF 1 REF SIGNAL FOR RMI NO. 1 & NO. 2 & COPILOT EHSI

COMPASS 1 REF TO RMI NO. 2

ADF 2 REF SIGNAL FOR RMI NO. 1 & NO. 2 & COPILOT EHSI

NAV 1 REF SIGNAL FOR RMI NO. 1 & NO. 2 & COPILOT EHSI

NAV 2 REF SIGNAL FOR RMI NO. 1 & NO. 2 & COPILOT EHSI

FMS FOR HDG & AUTOPILOT

COMPASS 1 REF TO DPU, MPU, & UNSIK

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

DUAL FED NO. 2 BUS

**

LEGEND

28 VDC POWER

115 VAC POWER

26 VAC POWER

GROUND

Figure 2-20. Inverters Control Circuit

2-17FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

LANDINGGEAR

MOTOR

AVIONICS BUS NO. 2

LANDINGGEARRELAY

AVIONICS NO. 2POWER RELAY

RIGHT GENERATOR BUS

TOINVERTERCONTROLSWITCH

BLUETESTJACK

INVERTERWARNRELAY

VOLTFREQ

METER

TO INVERTERSELECT SWITCH

26 VAC

115VAC

26 VAC

TO AVIONICS

OIL TEMP(ENG INSTR-RIGHT)

OIL PRESS(ENG INSTR-RIGHT)

FUEL FLOW(ENG INSTR-RIGHT)

TURBINE TACH(ENG INSTR-RIGHT)

PROP TACH(ENG INSTR-RIGHT)

TORQUE(ENG INSTR-RIGHT)

ITT(ENG INSTR-RIGHT)

FUEL CROSSFEED

LEFT FIREWALL VALVE

LEFT FUEL QUANTITY

LEFT START CONTROL

LEFT IGNITOR POWER

FLAP MOTOR

RIGHT FUELPRESSURE WARNING

RIGHT FUEL QUANTITY

RIGHT AUX FUEL QTYWARNING & TRANSFER

RIGHT STANDBY FUELPUMP

RIGHT FIREWALL VALVE

RIGHT START CONTROL

RIGHT IGNITOR POWER

PROPELLER GOVERNOR

RIGHT MANUALPROP DEICE

MANUAL PROPDEICE CONTROL

NO

. 4 D

UA

L F

ED

BU

S

TO AVIONICS

115VAC

INVERTERNO. 2

INVERTERNO. 1

TOINVERTERCONTROLSWITCH

SHUNT

RIGHTSTARTRELAY

R GEN LINECONTACTOR

R GEN CONTROL

RIGHTSTARTERGEN

V LMETER

– +

EXT POWER CONNECTOR

EXTERNALPOWER

EXT POWERRELAY

BATTERYCHARGEMONITOR ORBATTERYAMMETER

BATTERYRELAY

BATTERY

BATTERY

MA

IN B

ATTE

RY B

US

ISO

LATI

ON

BU

S

NO

. 1 D

UA

L FE

D B

US

NO

. 2 D

UA

L FE

D B

US

NO

. 3 D

UA

L FE

D B

US

BATTERYSWITCH

ON

OFF

SHUNT

+ + –

HOT BATTERYBUS

ENTRY LT. CLOCK LT.& EXT POWER SENSE

BATTERY RELAY

L ENG FIRE EXT

R ENG FIRE EXT

MOD

RIGHT FIREWALLSHUT OFF VALVE

LEFT FIREWALLSHUT OFF VALVE

LEFT GEN CONTROL

LEFTSTARTERGEN

LEFT GEN LINECONTACTOR

SHUNT

ISOLATION LIMITER

LEFTSTARTRELAY

V LMETER

ENGINEINSTRUMENT

– +


AVIONICS BUS NO. 1

ON

OFF

TO AVIONICSMASTER CONTROL CB

COPILOT'S WINDSHIELDANTI-ICE

AFT ELECTRIC HEAT

DC TEST JACK

VENT BLOWER POWER

AFT EVAP BLOWER

AIR CONDITIONERCLUTCH

SUBPANELS

TOENGINE

INSTRUMENTS CB(ISOLATION BUS)

OIL TEMP(ENG INSTR-LEFT)

OIL PRESS(ENG INSTR-LEFT)

FUEL FLOW(ENG INSTR-LEFT)

TURBINE TACH(ENG INSTR-LEFT)

PROP TACH(ENG INSTR-LEFT)

TORQUE(ENG INSTR-LEFT)

ITT(ENG INSTR-LEFT)

LEFT STANDBY FUELPUMP

LEFT AUX FUEL QTYWARNING & TRANSFER

LEFT FUELPRESSURE WARNING

FLAP CONTROLAND INDICATOR

LEFT MANUALPROP DEICE

CONDENSER BLOWERFWD ELECTRIC HEAT

AVIONICS BUS NO. 3

PILOT'S WINDSHIELDANTI-ICE

LEFT GENERATOR BUS

CABIN AUDIO(AVIONICS)

COPILOT AUDIO(AVIONICS)

MASTER POWER(FURNISHING)

CIGAR LIGHTER(FURNISHING)

RIGHT GEN CONTROL(ELECTRICAL)

RUDDER BOOST(FLIGHT)

COPILOT TURN &SLIP (ENVIRONMENTAL)

RIGHT BLEED AIRCONTROL (ENVIRONMENTAL)

CABIN TEMP CONTROL(ENVIRONMENTAL)

CABIN READINGLIGHT

COPILOT FLT INSTR(LIGHTS)

OVHD. SIDEPANEL &CONSOLE (LIGHTS)

INSTR INDIRECT(LIGHTS)

RIGHT BLEED AIRWARN (WARNING)

LANDING GEARIND (WARNING)

ANN IND(WARNING)

RIGHT FUEL VENT(WEATHER)

WSHLD WIPER(WEATHER)

RIGHT ENG INSTRPOWER (ENGINES)

RIGHT FUEL CONTROLHEAT (ENGINES)

RIGHT MAIN ENGANTI-ICE (ENGINES)

RIGHT STBY ENGANTI-ICE (ENGINES)

RIGHT CHIP DET(ENGINES)

AUTOFEATHER(ENGINES)

R LANDING LIGHT

TAXI LIGHT

ICE LIGHT

NAVIGATION LIGHT

RECOGNITION LIGHT

LANDING GEAR CONTROL

R PITOT HEAT

STALL WARNING HEAT

L LANDING LIGHT

L PITOT HEAT

PROP AUTOMATIC HEAT

TAIL FLOOD LIGHTS

STROBES LIGHTS

BEACON LIGHTS

PROP SYNC(ENGINES)

LEFT CHIP DET(ENGINES)

FIRE DET(ENGINES)

LEFT STBY ENGANTI-ICE (ENGINES)

LEFT MAIN ENGANTI-ICE (ENGINES)

LEFT FUEL CONTROLHEAT (ENGINES)

ISOLATION LIMITER

LEFT ENG INSTRPOWER (ENGINES)

BRAKE DEICE(WEATHER)

SURF DEICE(WEATHER)

LEFT FUEL VENT(WEATHER)

ANN POWER(WARNING)

LANDING GEARWARN (WARNING)

L BLEED AIRWARN (WARNING)

STALL WARN(WARNING)

PLT FLT & SIDEPNL (LIGHTS)

NO SMK FSB &CABIN (LIGHTS)

AVIONICS & ENG INSTR(LIGHTS)

AVIONICS ANN(LIGHTS)

OXYGEN CONTROL(ENVIRONMENTAL)

PRESS CONTROL(ENVIRONMENTAL)

LEFT BLEED AIRCONTROL (ENVIRONMENTAL)

ALT ALERT(FLIGHT)

PILOT TURN &SLIP (FLIGHT)

PILOT ATM AIRDATA (FLIGHT)

PITCH TRIM(FLIGHT)

OUTSIDE AIRTEMP (FLIGHT)

LEFT GEN CONTROL(ELECTRICAL)

AURAL WARN(AVIONICS)

AVIONICS MASTER(AVIONICS)

PILOT AUDIO(AVIONICS)

* (SERIALS PRIOR TO BB-1632 AND BW-30)

** (SERIALS BB-1632 AND AFTER, BL-141 AND AFTER, BW-30 AND AFTER)

**

*

Figure 2-21. Electrical System—Super King Air B200 (BB-1484, 1486 and Subsequent;BW-1 and Subsequent)




FUEL CROSSFEED


RIGHT STANDBYFUEL PUMP


RIGHT FUEL QUANTITY


NAV MEMORY

HOT BATTERYBUS


BATTERY RELAY

L ENG FIRE EXT

R ENG FIRE EXT

MOD

RIGHT FIREWALLSHUTOFF VALVE

LEFT FIREWALLSHUTOFF VALVE

AVIONICSBUS NO. 1




MA

IN B

AT

TE

RY

BU

S

ISO

LA

TIO

N B

US

LANDINGGEAR

MOTORLANDINGGEARRELAY

ISOLATION LIMITER

R GEN CONTROL

SHUNT

V LMETER

RIGHTSTARTERGEN

– +

–++

R GEN LINECONTACTOR

RIGHTSTARTRELAY

BATTERYCHARGEMONITOR BATTERY

SHUNT

EXT POWERRELAY

EXTERNALPOWER

LEFTSTARTRELAY

EXT POWER CONNECTOR

ON

OFFBATTERYSWITCH

BATTERYRELAY

AVIONICSBUS NO 2


AFT ELECTRIC HEAT

DC TEST JACK

RIGHT GENERATOR BUS

TO INVERTERCONTROLSWITCH

VENT BLOWER POWER

AFT EVAP BLOWER


MOD/OPTIONAL FURNISHINGS

BLUETESTJACK

INVERTERWARNRELAY

VOLTFREQ

METER


LEFT GENERATOR BUS

26 VAC

115VAC

26 VAC

26 V

AC

BU

S

TO AVIONICS

TO AVIONICS

115VAC

NO

. 2 D

UA

L F

ED

BU

S

NO

. 3 D

UA

L F

ED

BU

S

NO

. 4 D

UA

L F

ED

BU

S

26 V

AC

115

VAC

26 V

AC

115

VAC

NO

. 1 D

UA

L F

ED

BU

S

INVERTERNO. 2

RIGHT START CONTROL

RIGHT IGNITOR POWER

PROPELLER GOVERNOR

MANUAL PROP DEICECONTROL

RIGHT MANUAL PROPDEICE

INVERTERNO. 1

AVIONICS BUS NO. 3

FWD ELECTRIC HEAT


CONDENSER BLOWER

ON

OFF


SHUNT

ISOLATIONLIMITER


L. GEN CONTROL

V LMETER

– + LEFTSTARTERGEN

LEFT TORQUE METER

RIGHT TORQUE METER

YAW RATE

LEFT START CONTROL

RIGHT GEN CONTROL

CIGAR LIGHTER

RUDDER BOOSTCONTROL

CPILOT ILS INDICATOR

RIGHT BLEED AIRCONTROL

CABIN TEMPERATURECONTROL

CABIN READINGLIGHTS

AVIONIC & ENGINEINSTRUMENT LIGHTS

OVHD. SUBPANEL ANDCONSOLE LIGHTS

CABIN LIGHTS &ORDINANCE

R BLEED AIRWARNING

LANDING GEARPOSITION IND

ANNUNCIATORINDICATOR

R FUEL VENT HEAT

WINDSHIELD WIPER

R FUEL FLOWINDICATOR

R OIL PRESSUREINDICATOR

R OIL TEMPINDICATOR

R ENGINE FUELCONTROL HEAT

R ICE VANE CONTROL

R ICE VANE EMER

AUTOFEATHER

R CHIP DETECTOR

FURNISHINGS MASTERCONTROL

STALL WARNING HEAT

TAXI LIGHT

ICE LIGHTS

NAVIGATION LIGHT

RECOGNITION LIGHT

L PILOT HEAT

LANDING GEARCONTROL

R LANDING LIGHT

LEFT IGNITOR POWER


FLAP MOTOR

LEFT MANUAL PROPDEICE

LEFT FIREWALL VALVE

LEFT STANDBYFUEL PUMP


LEFT FUEL QUANTITY


LEFT GEN CONTROL

OUTSIDE AIR TEMP

PITCH TRIM

PILOT ILS INDICATOR

LEFT BLEED AIRCONTROL

CABIN PRESSURECONTROL

AUTOMATIC OXYGENCONTROL

CPILOT FLT INSTRLIGHTS

INSTRUMENT INDIRECTLIGHTS

PILOT FLT INSTRSIDE PANEL

OVHD FLOOD LIGHTS

L BLEED AIRWARNING

LANDING GEARWARNING HORN

ANNUNCIATOR POWER

L FUEL VENT HEAT

BRAKE DEICE

L FUEL FLOWINDICATOR

L OIL PRESSUREINDICATOR

L OIL TEMPINDICATOR

L ENGINE FUELCONTROL HEAT

L ICE VANE CONTROL

L ICE VANE EMER

FIRE DETECTION

L CHIP DETECTOR

PROP SYNCHROPHSERPROP BALANCE

BEACON LIGHTS

STROBE LIGHTS

TAIL FLOOD LIGHTS

PROP AUTOMATIC HEAT

L PILOT HEAT

L LANDING LIGHT

YAW DAMPER

AVIONICS MASTERCONTROL

PNEUMATIC SURFACEDEICE

STALL WARNINGSYSTEM

Figure 2-22. Electrical System—Super King Air B200 (BB-1449, 1458-1462, 1464-1485,Except 1484; BL-139, 140)




FUEL CROSSFEED




RIGHT FUEL QUANTITY


NAV MEMORY

HOT BATTERYBUS


BATTERY RELAY

L ENG FIRE EXT

R ENG FIRE EXT

MOD

RIGHT FIREWALLSHUTOFF VALVE

LEFT FIREWALLSHUTOFF VALVE

AVIONICSBUS NO. 1




MA

IN B

AT

TE

RY

BU

S

ISO

LA

TIO

N B

US

LANDINGGEAR

MOTORLANDINGGEARRELAY

ISOLATION LIMITER

R GEN CONTROL

SHUNT

V LMETER

RIGHTSTARTERGEN

– +

–++

R GEN LINECONTACTOR

RIGHTSTARTRELAY

BATTERYCHARGEMONITOR BATTERY

SHUNT

EXT POWERRELAY

EXTERNALPOWER

LEFTSTARTRELAY

EXT POWER CONNECTOR

ON

OFFBATTERYSWITCH

BATTERYRELAY

AVIONICSBUS NO 2


AFT ELECTRIC HEAT

DC TEST JACK

RIGHT GENERATOR BUS


VENT BLOWER POWER

AFT EVAP BLOWER


MOD/OPTIONAL FURNISHINGS

BLUETESTJACK

INVERTERWARNRELAY

VOLTFREQ

METER


LEFT GENERATOR BUS

26 VAC

115VAC

26 VAC

26 V

AC

BU

S

TO AVIONICS

TO AVIONICS

115VAC

NO

. 2 D

UA

L F

ED

BU

S

NO

. 3 D

UA

L F

ED

BU

S

NO

. 4 D

UA

L F

ED

BU

S

26 V

AC

115

VAC

26 V

AC

115

VAC

NO

. 1 D

UA

L F

ED

BU

S

INVERTERNO. 2

RIGHT START CONTROL

RIGHT IGNITOR POWER

PROPELLER GOVERNOR

MANUAL PROP DEICECONTROL

RIGHT MANUAL PROPDEICE

INVERTERNO. 1

FWD ELECTRIC HEAT


CONDENSER BLOWER

ON

OFF

AVIONICS BUS NO. 3


SHUNT

ISOLATIONLIMITER


L. GEN CONTROL

V LMETER

– + LEFTSTARTERGEN

LEFT TORQUE METER

RIGHT TORQUE METER

YAW RATE

LEFT START CONTROL

RIGHT GEN CONTROL

NO. 2 INV CONTROL

CIGAR LIGHTER

RUDDER BOOSTCONTROL

CPILOT ILS INDICATOR



CABIN READINGLIGHTS

AVIONIC & ENGINEINSTRUMENT LIGHTS

OVHD. SUBPANEL ANDCONSOLE LIGHTS

CABIN LIGHTS &ORDINANCE

R BLEED AIRWARNING



R FUEL VENT HEAT

WINDSHIELD WIPER

R FUEL FLOWINDICATOR

R OIL PRESSUREINDICATOR

R OIL TEMPINDICATOR

R ENGINE FUELCONTROL HEAT

R ICE VANE CONTROL

R ICE VANE EMER

AUTOFEATHER

R CHIP DETECTOR


STALL WARNING HEAT

TAXI LIGHT

ICE LIGHTS

NAVIGATION LIGHT

RECOGNITION LIGHT

L PILOT HEAT

LANDING GEARCONTROL

R LANDING LIGHT

LEFT IGNITOR POWER


FLAP MOTOR

LEFT MANUAL PROPDEICE

LEFT FIREWALL VALVE



LEFT FUEL QUANTITY


LEFT GEN CONTROL

NO. 1 INV CONTROL

OUTSIDE AIR TEMP

PITCH TRIM

PILOT ILS INDICATOR



AUTOMATIC OXYGENCONTROL

CPILOT FLT INSTRLIGHTS

INSTRUMENT INDIRECTLIGHTS

PILOT FLT INSTRSIDE PANEL

OVHD FLOOD LIGHTS

L BLEED AIRWARNING


ANNUNCIATOR POWER

L FUEL VENT HEAT

BRAKE DEICE

L FUEL FLOWINDICATOR

L OIL PRESSUREINDICATOR

L OIL TEMPINDICATOR

L ENGINE FUELCONTROL HEAT

L ICE VANE CONTROL

L ICE VANE EMER

FIRE DETECTION

L CHIP DETECTOR

PROP SYNCHROPHSERPROP BALANCE

BEACON LIGHTS

STROBE LIGHTS

TAIL FLOOD LIGHTS

PROP AUTOMATIC HEAT

L PILOT HEAT

L LANDING LIGHT

YAW DAMPER



STALL WARNINGSYSTEM

Figure 2-23. Electrical System—Super King Air B200 (BB-1439, 1444-1448, 1450-1457)




AVIONICSBUS NO. 1

TO AVIONICSMASTERCONTROL CB

ON

LEFT GENCONTROL

LINECONTACTOR

SHUNT

ISOLATION LIMITER

EXT POWER CONNECTOR

OFF



LEFTSTARTER

GEN

RIGHTSTARTER

GEN

LEFTSTARTRELAY

HOT BATTERYBUS

BATTERYSW

CIRCUITEDINTOEXTERNALPOWERRELAY

EXT POWER RELAY

BATTERYRELAY

MA

IN B

AT

TE

RY

BU

S

ISO

LA

TIO

N B

US

BATTERY SW

ENTRY LTS &CLOCK LT &EXT PWRSENSE

RNAV MEMORY(OPT)

STEREO (OPT)

L ENG FIRE EXT




REMOVEDFROM HOTBUS ONBB1098 ANDAFTER


R ENG FIRE EXT

THIS LINE OFFUSES CHANGETO 5A CIRCUITBREAKERS ON BB1098 ANDAFTER

BATTERYSH

UN

T

BATTERYCHARGESENSOR

VOLTLOAD

METER

LANDINGGEAR

MOTOR

VOLTLOAD

METER

SHUNT

RIGHT GENCONTROL

RIGHTSTARTRELAY

ISOLATION LIMITER

LINECONTACTOR

AVIONICSBUS NO 2

COPILOT'SWINDSHIELD ANTI-ICE

RIGHT RADIANTHEAT

DC TEST JACK

RIGHT GEN BUS

TO INVCONTROLDUAL FEDBUS NO. 2

VENTBLOWER POWER

AFT EVAPORATORBLOWER POWER


CONTROL ANDOVERLOADPROTECTION

MOD OPTIONALEQUIPMENT

AVIONICS NO. 3POWER RELAY(OPTIONAL)

AVIONICSBUS NO 3

BLUETESTJACK

TO INVCONTROLDUALFED BUSNO. 1

LEFT GEN BUS

RELAYPANEL

INVNO. 1

INVWARNRELAY

26 VAC115VAC

VOLTSFREQ.METER

NO

2 D

UA

L F

ED

BU

S

NO

. 3 D

UA

L F

ED

BU

S

NO

. 4 D

UA

L F

ED

BU

S

26 V

AC

115

VAC

NO

1 D

UA

L F

ED

BU

SS

UB

PA

NE

LS

LEFT LANDING LIGHT

LEFT PITOT HEAT

PROP AUTOMATICHEAT SWITCH

TAIL FLOOD LIGHTSSWITCH (OPT)

INVNO. 2

AV

ION

ICS

AV

ION

ICS

- +

LEFT GEN CONTROL

LEFT RADIANT HEAT

PILOT'SWINDSHIELDANTI-ICE

CONDENSERBLOWER

NO. 1 INVERTERCONTROL

PITCH TRIM

CIGARETTE LIGHTER


RIGHT GENERATORCONTROL


RUDDER BOOSTCONTROL

COPILOT TURN ANDSLIP


CABIN TEMP CONTROL

CABIN PRESSURELOSS (OPT)

PILOT'S TURN ANDSLIP

ENCODER ALTIMETER(OPT)

YAW DAMPER



AUTOMATIC OXYCONTROL

BRAKE DEICE (OPT)

WINDSHIELD WIPER

LEFT TORQUEMETER

RIGHT TORQUEMETER

YAWRATE

FUELCROSSFEED

RIGHT FUELPRESSUREWARNING

RIGHT FUELQUANTITY

RIGHT AUX FUELQUANTITYWARNING ANDTRANSFER


RIGHT FIREWALLVALVE

LEFT FIREWALLVALVE


LEFT AUX FUELQUANTITY WARNING AND TRANSFER

LEFT FUELQUANTITY

LEFT FUELPRESS WARN

LEFT STARTCONTROL

LEFT IGNITORPOWER


FLAP MOTOR


RIGHT STARTERCONTROL

RIGHT IGNITORPOWER

PROPELLERGOVERNOR



RIGHT FUEL VENT HEAT



RIGHT BLEED AIRWARNING

FLOURESCENT LIGHTSAND ORD WARNING

(OVHD) SUBPANEL ANDCONSOLE LIGHTS

AVIONICS & ENGINEINSTRUMENT LTS

CABIN READING LIGHTS

AUTOFEATHER

RIGHT CHIP DETECTOR

RIGHT GEN OVERHEAT(OPT)

R ICE VANE CONT

RIGHT OIL PRESSUREWARN (OPT)

RIGHT OILTEMPERATUREINDICATOR

RIGHT OIL PRESSUREINDICATOR

RIGHT FUEL FLOWINDICATOR

STALL WARNING HEAT

RIGHT PITOT HEAT

LANDING GEAR CONTROL

RECOGNITION LIGHT

NAV LIGHT SWITCH

ICE LIGHTS

TAXI LIGHT SWITCH

RIGHT LANDING LIGHTSWITCH

RIGHT ENGINE FUELCONTROL HEAT

LEFT FUEL VENT HEAT

ANNUNCIATORPOWER


LEFT BLEED AIRWARNING

STALL WARNINGSYSTEM

OVERHEAD AND SIDEPANEL LIGHTS

INSTRUMENTINDIRECT LIGHTS

FLIGHT & GYROINSTRUMENT LIGHTS

TAIL FLOOD LIGHT(OPT)

PROPSYNCHROPHASER

LEFT CHIP DETECTOR

LEFT GEN OVERHEAT

FIRE DETECTION

L ICE VANE CONT

LEFT ENGINE FUELCONT HEAT

LEFT OILTEMPERATURE

LEFT OIL PRESSUREWARNING (OPT)

LEFT OIL PRESSURE

LEFT FUEL FLOW


BEACON LIGHTSSWITCH

STROBE LIGHTSSWITCH


+ +

+

–

–

+ –

Figure 2-24. Electrical System—Super King Air B200 (BB-734, 793, 829, 854-870, 874-891,894, 896-911, 913-1438, 1440-1443, BL 37-138)




AVIONICSBUS NO. 1

TO AVIONICSMASTERCONTROL CB

ON

LEFT GENCONTROL

REVERSECURRENTPROTECTIONSHUNT

ISOLATION LIMITER

EXT POWER CONNECTOR

OFF



LEFTSTARTER

GEN

RIGHTSTARTER

GEN

LEFTSTARTRELAY

HOT BATTERY

BUS

SMALLPIN

EXT POWER RELAY

BATTERYRELAY

MA

IN B

AT

TE

RY

BU

S

ISO

LA

TIO

N B

US

BATTERY SW

LEFT FIRE EXT

THRESHOLD LT





RIGHT FIRE EXT

BATTERY

BATTERYCHARGESENSOR

VOLTLOAD

METER

LANDINGGEAR

MOTOR

VOLTLOAD

METER

SHUNT REVERSECURRENTPROTECTION

RIGHT GENCONTROL

RIGHTSTARTRELAY

ISOLATION LIMITER

TO ANNUNCIATORADVISORY LIGHT

EXTERNALPOWER PLUGENGAGED

BB 364AND SUBSEQUENT

AVIONICSBUS NO 2

DC TEST JACK

COPILOT'SWINDSHIELD ANTI-ICE

RIGHT RADIANTHEAT

RIGHT GEN BUS

TO INVCONTROLNO. 2 DUAL FEDBUS

VENTBLOWER POWER

AFT EVAPORATORBLOWER POWER

AIR CONDITIONERCLUTCH CONTROL

ANDOVERLOADPROTECTION*BB-1 THRU BB-224


AVIONICSBUS NO 3

LEFT FUELQUANTITY

RIGHT FUELQUANTITY

*L FUEL FLOW

LEFT TORQUEMETER

RIGHT TORQUEMETER

*R FUEL FLOW

INVWARNRELAY

BLUETESTJACK

TO INVCONTROLNO. 1 DUALFED BUS

LEFT GEN BUS

INVSELECT

RELAY

INVNO. 1

26 VAC

115VAC

FUEL CROSSFEED

RIGHT FUELPRESSUREWARNING

RIGHT AUX FUELQUANTITYWARNING ANDTRANSFER


RIGHT FIREWALLVALVE

LEFT FIREWALLVALVE


LEFT AUX FUELQUANTITYWARNING ANDTRANSFER

LEFT FUELPRESSUREWARNING

LEFT STARTERCONTROL

LEFT IGNITORPOWER

FLAP CONTROL& INDICATOR

FLAP MOTOR


**L FUEL FLOW

NO

2 D

UA

L F

ED

BU

S

NO

. 3 D

UA

L F

ED

BU

S

NO

. 4 D

UA

L F

ED

BU

S

26 V

AC

115

VAC

RIGHT STARTERCONTROL

RIGHT IGNITORPOWER

PROPELLERGOVERNOR



NO

1 D

UA

L F

ED

BU

SS

UB

PA

NE

LS

FUEL DRAINCOLLECTOR PUMPS

LEFT GENCONTROL

NO 1 INVERTERCONTROL

TRIM TAB

PILOT'S TURN ANDSLIP INDICATOR

OPTIONALALTIMETER

PNEUMATICSURFACE DE-ICE

LEFT FUELVENT HEATER

ANNUNCIATORPOWER


LEFT BLEEDAIR WARNING

STALL WARNINGSYSTEM

OVERHEAD ANDSIDE PANEL LIGHTS

INSTRUMENTINDIRECT LIGHT

FLIGHTINSTRUMENT LIGHT

LOGO LIGHT


LEFT BLEEDAIR CONTROL

PROPSYNCHROPHASER

FIRE DETECTION

LEFT ICE VANE

LEFT ENGINE FUELCONTROL HEAT

LEFT OIL TEMPINDICATOR

LEFT OIL PRESSUREINDICATOR

AVIONICSMASTER CONTROL

STROBE LIGHTS

BEACONLIGHTS SW

PROP DEICEAUTO HEAT SW

LEFT PITOTHEAT SW

LEFT LANDINGLIGHT SW

RIGHT GENCONTROL


RUDDER BOOSTSYSTEM

VACUUM DAMPERSYSTEM

COPILOT'S TURN &SLIP INDICATOR

WINDSHIELD WIPER

RIGHT FUELVENT HEATER


LANDING GEARPOSITION INDICATOR

RIGHT BLEEDAIR WARNING

CABIN FASTEN SEATBELT & NO SMOKINGSIGN AND CHIMES

SUBPANEL ANDCONSOLE LIGHTS

RADIO & ENGINEINSTRUMENT LIGHTS


RIGHT BLEEDAIR CONTROL

RIGHT ENGINE HEATFUEL CONTROL

RIGHT OIL TEMPINDICATOR

RIGHT OIL PRESSUREINDICATOR

STALL WARNINGHEAT

RIGHT PITOTHEAT SW

LANDING GEARCONTROL

NAV LIGHT SW

ICE LIGHT SW

TAXI LIGHT SW

RIGHT LANDINGLIGHT SW

AUTOFEATHER

CHIP DETECTOR

RIGHT ICE VANE

**R FUEL FLOW

INVNO. 2

AV

ION

ICS

AV

ION

ICS

LEFT RADIANTHEAT

PILOT'SWINDSHIELDANTI-ICE

CONDENSERBLOWER

SH

UN

T

OPTIONALEQUIPMENT

***

***SEE EXTERNAL POWER SECTION FOR DETAILS

**BB-225 & AFTER

Figure 2-25. Electrical System—Super King Air 200 (BB-2, 6-733, 735-792, 794-828, 830-853, 871-873, 892, 893, 895, 912, BL-1-36)

LIMITATIONS

GENERATOR LIMITS (250AMPERES)Maximum sustained generator load (Table 2-1) is limited as follows:

In Flight:Sea Level to 31,000 feet altitude .................. 1.00 (100%)

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

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

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

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

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

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




GENERATOR LOAD MINIMUM GAS GENERATOR RPM – N1

WITHOUT AIR WITH AIR CONDITIONING CONDITIONING

0 to 70% 52% 60% 70 to 75% 55% 60% 75 to 80% 60% 60% 80 to 85% 65% 65%

BB – 1439, 1444 AND SUBSEQUENT

0 to 75% 61% 62% 75 to 80% 61% 62% 80 to 85% 65% 65%

Table 2-1. LIMITATIONS—GROUND OPERATIONS




CIRCUIT BREAKER NAME CAPACITY PROVIDES POWER TO

1. FUEL SYSTEM

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

NO TRANSFER LIGHT

AUX TANK FLOAT SWITCH

MOTIVE FLOW VALVE

B. CROSSFEED 5 AMP CROSSFEED SWITCH

CROSSFEED VALVE

AUX FUEL TRANSFER MODULE

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

FIREWALL VALVE

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

FUEL PRESS WARNING LIGHT

AUX FUEL TRANSFER MODULE

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

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

AUX TRANSFER PCB

G. BUS FEEDERS

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

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

2. FLAP

A. MOTOR 20 AMP MOTOR RELAY AND MOTOR POWER

B. CONTROL 5 AMP FLAP POTENTIOMETER (POSITION XMTR)

SPLIT FLAP

HOBBS METER

FLAP POSITION INDICATOR

3. PROP

A. GOVERNOR 5 AMP OVERSPEED GOVERNOR TEST SWITCH

B. PROP DEICE

(1) CONTROL 5 AMP MANUAL SWITCH POWER

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

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

4. START CONTROL

A. CONTROL (L & R) 5 AMP ENGINE START SWITCH

STARTER RELAY

IGNITER AND PURGE VALVE CONTROL

AUTOIGNITION CONTROL SWITCH

B. IGNITER POWER (L & R) 5 AMP IGNITER POWER

IGNITER PURGE VALVE

Table 2-2. FUEL CONTROL CIRCUIT-BREAKER PANEL





5. ENGINE INSTRUMENTS

(BB-1484, 1486 AND

SUBSEQUENT)

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

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

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

D. TURBINE TACH (L & R) 5 AMP GAGE POWER

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

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

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

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


1. ELECTRICAL DISTRIBUTION

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

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

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

GENERATOR CONTROL PANEL

2. INVERTER CONTROL

(BB-1439, 1444-1448,

1450-1457 AND PRIOR)

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

AND CONTROL RELAY

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

AND CONTROL RELAY

3. ENGINE

A. AUTOFEATHER 5 AMP POWER LEVER ARM SWITCHES

AUTOFEATHER ARM SWITCHES (400 & 200 FT-LB)

B. CHIP DETECTOR 5 AMP BB 1-162, CHIP DETECTOR AND LIGHT

C. CHIP DETECTOR (L & R) 5 AMP BB 163 AND SUBSEQUENT

L & R CHIP DETECTOR AND

L & R CHIP DETECTOR LIGHTS

D. FIRE DETECT 5 AMP TERMINAL BOARD AMPLIFIER

(PRIOR TO BB-1444, EXCEPT 1439)

DETECTOR AND TEST SWITCH

Table 2-3. RIGHT SIDE CIRCUIT-BREAKER PANEL





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

(THROTTLE QUADRANT)

L & R FUEL CONTROL PNEUMATIC LINE HEAT

(PRIOR TO BB-1444, EXCEPT 1439)

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

COLLECTOR PUMP (BB 2-665)

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

TRANSMITTER (AC)

5 AMP BB 225 – 1483, 1485 FUEL FLOW INDICATOR

AND TRANSMITTER (DC)

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

I. ICE VANE CONTROL (L & R)

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

ICE VANE SENSE MODULE

ICE VANE ACTUATOR

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

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

MN ENG ANTI-ICE SENSE MODULE

MN ENG ANTI-ICE ACTUATOR

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

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

STBY ENG ANTI-ICE SENSE MODULE

STBY ENG ANTI-ICE ACTUATOR

L. OIL PRESS (L & R)

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

M. OIL TEMP (L & R)

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

N. TORQUEMETER (L & R)

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

O. PROP SYNC 5 AMP PROP SYNCHROPHASER CONTROL BOX

SYNCH CONTROL SWITCH

4. ENVIRONMENTAL

A. BLEED-AIR CONT (L & R) 5 AMP BLEED AIR CONTROL SWITCH

ANNUN BLEED AIR OFF LIGHT

FLOW CONTROL PACKAGE

PNEUMATIC SHUTOFF VALVE

RUDDER BOOST

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






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

MASKS 12,500 FT PRESSURE SWITCH

C. PRESS CONT 5 AMP LEFT SQUAT SWITCH

PEDESTAL PRESS CONTROL SWITCH

SAFETY VALVE DUMP SOLENOID

EVAPORATOR DOOR SOLENOID

CABIN DOOR SOLENOID

D. TEMP CONTROL 5 AMP VENT BLOWER CONTROL SWITCH

LEFT SQUAT SWITCH

AMBIENT AIR VALVES AND PCB

BB 1-450 RADIANT HEAT CONTROL SWITCH

BB 450 AND SUBSEQUENT RADIANT HEAT

POWER CIRCUIT BREAKER

CABIN TEMP MODE SELECTOR SWITCH

5. FLIGHT

A. PITCH TRIM 5 AMP PEDESTAL ELECTRIC ELEVATOR

TRIM SWITCH

TRIM MOTOR

B. RUDDER BOOST 5 AMP PEDESTAL ON/OFF SWITCH

DIFFERENTIAL PRESSURE SWITCH

RUDDER BOOST SOLENOIDS

C. TURN AND SLIP 5 AMP TURN AND SLIP INDICATOR

D. ENCODING ALTIMETER 1 AMP ALTIMETER (ENCODING)

6. LIGHTS

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

PILOT AND COPILOT CLOCK AND MAP LIGHTS

B. FLIGHT INST 7.5 AMP OVERHEAD PANEL AND TERMINAL BOARD

PILOT & COPILOT FLIGHT INST LIGHTS

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

CABIN FLUORESCENT LIGHTS

CABIN WARNING CHIME

D. INST INDIRECT 5 AMP GLARESHIELD LIGHTS

OVERHEAD PANEL

APPROACH PLATE LIGHTS

E. SIDE PANEL 5 AMP RIGHT SIDE CIRCUIT BREAKER PANEL LIGHTS

FUEL PANEL LIGHTS

AVIONICS PANEL LIGHTS

OVERHEAD PANEL OR PILOT’S RIGHT SUBPANEL






7. WARNING

A. ANNUN INDICATOR 5 AMP ANNUN N1 LOW LIGHT

ANNUN INVERTER OUT LIGHT

ICE VANE PCB

O2 PRESS SWITCH

CABIN ALT WARN PRESS SWITCH (12,500)

BATTERY CHARGE MODULE

DUCT OVERTEMP SWITCH

ALTITUDE WARNING LIGHT

B. ANNUN POWER 5 AMP 28V ANNUNCIATOR CONTROL CARD

MASTER WARNING LIGHTS

MASTER CAUTION LIGHTS

CAUTION LEGEND SWITCH

C. BLEED AIR WARNING (L & R) 5 AMP BLEED AIR WARN LIGHTS

BLEED AIR WARN PRESS SWITCH

D. LANDING GEAR INDICATOR 5 AMP GREEN GEAR DN LIGHTS

RED GEAR HANDLE LIGHTS

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

GEAR WARNING HORN SILENCE BUTTON & RELAY

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

8. WEATHER

A. BRAKE DEICE 5 AMP LEFT UPLOCK SWITCH

BATTERY CHARGE/ DEICE MODULE

BRAKE DEICE SWITCH

DEICE BLEED AIR VALVES

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

HEATER ELEMENTS

C. SURFACE DEICE 5 AMP SURFACE DEICE SWITCH

DEICE DISTRIBUTOR VALVE

TIME DELAY PCB

D. WINDSHIELD WIPERS 10 AMP OVERHEAD PANEL SWITCH

WIPER MOTOR POWER

9. AVIONICS

A. AVIONICS MASTER 5 AMP AVIONICS MASTER SWITCH TO

KEEP AVIONICS OFF

B. AVIONICS NO. 1 30 AMP AVIONICS BUS NO. 1 FEEDER

C. AVIONICS NO. 2 30 AMP AVIONICS BUS NO. 2 FEEDER




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


1. ICE PANEL

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

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

AND DEICE TIMER

25 AMP BB-1439, 1444 AND SUBSEQUENT

C. STALL WARN 15 AMP STALL WARNING HEAT CONTROL RELAY

2. LANDING GEAR

A. LANDING GEAR RELAY 5 AMP LANDING GEAR CONTROL SWITCH

RVS NOT READY ANNUNCIATOR POWER

B. LANDING GEAR RELAY

(HYDRAULIC GEAR) 2 AMP LANDING GEAR CONTROL SWITCH

HYD FLUID LOW LIGHT

RVS NOT READY ANNUNCIATOR POWER

3. LIGHTS

A. ICE 5 AMP ICE LIGHTS

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

C. NAV LIGHT 5 AMP NAV LIGHTS

D. TAIL FLOODLIGHT 15 AMP TAIL FLOODLIGHTS

OVERHEAD PANEL LIGHTS

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

RELAY AND LIGHTS (2 BULB)

7.5 AMP BB 178 AND SUBSEQUENT RECOG LIGHT

(1 BULB)

F. TAXI LIGHT 15 AMP TAXI LIGHT

G. BEACON 10 AMP BEACONS

H. STROBE 5 AMP STROBE POWER SUPPLY & STROBE TUBE

3-i

CHAPTER 3LIGHTING

CONTENTS

Page

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

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

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

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

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

EXTERIOR LIGHTS .............................................................................................................. 3-7

Landing Lights ................................................................................................................. 3-7

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

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

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

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

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

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

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

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

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

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




3-iii

ILLUSTRATIONS

Figure Title Page

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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




INTRODUCTIONThe instruments are illuminated either internally or with post-type lights. General cabinlighting consists of overhead fluorescent lights and individual passenger reading lights.A passenger FASTEN SEAT BELT–NO SMOKING sign is provided. Both the airstairand baggage area are illuminated. Exterior lights consist of landing, taxi, ice inspection,navigation, recognition, beacon, strobe, and lights for the area around the airstair door.Optional lighting is available to illuminate the vertical tail fin.

GENERALAn overhead light control panel in the cock-pit contains controls for instrument panel andcockpit lighting (Figure 3-1). Each light grouphas an individual rheostat switch labeledBRT–OFF. The MASTER PANEL LIGHTSswitch controls power to the overhead lightcontrol panel lights, fuel control panel lights,

engine instrument lights, radio panel lights,both subpanels and the console lights, pilot andcopilot instrument lights, and gyro instrumentlights. Separate rheostat switches individu-ally control the instrument indirect lighting andthe overhead floodlights.

EXIT

CHAPTER 3LIGHTING




A FREE–AIR–TEMPERATURE gage is lo-cated on the left sidewall aft of the fuel panel.For BB-1439, 1444 and subsequent, a digitaldisplay indicates the free air temperature inCelsius. Prior to BB-1444, excluding 1439, ananalog temperature display also indicates thetemperature in Celsius.

A switch on the copilot’s left subpanel labeledBRIGHT–DIM–OFF (prior to BB-1444, ex-cluding 1439, it is labeled START/BRIGHT–DIM–OFF) (Figure 3-2) controls the fluores-cent overhead cabin lights. To the right of theinterior light switch is a switch labeled NO

SMOKE & FSB–FSB–OFF. It controls the NOSMOKING–FASTEN SEAT BELT sign andthe accompanying chimes.




33N3

FO

R0

3060

90120 150 180 210

24027

030

033

0


RADIO ON

ST

EE



AIRSPEEDS (IAS)



CAUTION




PARK SLOW

FAST

OFF

MASTERPANELLIGHTS

ON

OVERHEADFLOODLIGHTS

OFFBRT


OFFBRT

AVIONICSPANELLIGHTS

OFFBRT

ENGINEINSTRUMENT

LIGHTS

OFFBRT

PILOTFLIGHTLIGHTS

OFFBRT


LIGHTS

OFFBRT

SIDEPANELLIGHTS

OFFBRT


LIGHTS

OFFBRT

COPILOTFLIGHTLIGHTS

OFFBRT

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTSAC VOLTS

FREQ

Beechcraft

0 8020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

-60

0

+60BATT AMPS

Figure 3-1. Overhead Lighting Controls

Figure 3-2. Copilot’s Left Subpanel




A pushbutton switch next to each light con-trols the individual passenger reading lightsalong the top of the cabin.

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

A threshold light located forward of the airstairdoor at floor level, and an aisle light locatedat floor level aft of the spar cover, are con-trolled by a switch next to the threshold light.When the airstair is open, these lights comeon; when it is closed and locked, they auto-matically extinguish. A flush-mounted flood-light forward of the flaps in the bottom of theleft wing and under the stair lights are also con-trolled by the threshold light switch. They il-luminate the area around the airstair when itis open and the switch is turned on.

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

Tail floodlights, if installed, are controlled bya switch located either on the overhead panelor the pilot’s right subpanel.

INTERIOR LIGHTING

COCKPIT

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

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

Pilot Flight LightsThe PILOT FLIGHT LIGHTS rheostat for thepilot’s flight instrument area controls the in-ternal or eyebrow post lights. The flight lightsare shown in Figure 3-4.

Pilot Gyro Instrument LightsFor EFIS equipped aircraft, the EADI andEHSI intensity are controlled by the EFISdimming rheostats.

For non-EFIS equipped aircraft, the PILOTGYRO INSTRUMENT LIGHTS rheostat con-trols the pilot’s gyro horizon and horizontalsituation indicator on the pilot’s flight panel.

Figure 3-3. Pilot’s Right Subpanel

Figure 3-4. Instrument and Panel Lights

Avionics Panel LightsThe AVIONICS PANEL LIGHTS rheostatcontrols the illumination of the internal avion-ics panel lights.

Overhead Subpanel andConsole LightsThe rheostat labeled OVERHEAD SUB-PANEL & CONSOLE LIGHTS controls thelighting for the overhead subpanel and thethrottle console (Figure 3-5 and Figure 3-6).

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

Copilot’s Gyro InstrumentLightsThe COPILOT GYRO INSTRUMENTLIGHTS rheostat controls any gyro instru-ments on the copilot’s flight panel.

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




Figure 3-7. Copilot’s Instrument Lights

Figure 3-5. Console Lights

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTSAC VOLTS

FREQ

Beechcraft

0 8020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

-60

0

+60BATT AMPS

Figure 3-6. Overhead Subpanel Lights




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

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

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

CABIN

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

Passenger Warning SignA sign to warn passengers not to smoke and/orto fasten their seat belts (Figure 3-11) is con-trolled by a switch on the copilot’s subpanel.The switch has three positions which are NOSMOKE & FSB–FSB–OFF. In FSB, the FAS-

Figure 3-8. OAT Gage

Figure 3-9. Free Air Temperature Switch

Figure 3-10. Fluorescent Light Switch

TEN SEAT BELT portion of the sign illumi-nates. The NO SMOKING and FASTEN SEATBELT positions are illuminated in the NOSMOKE & FSB position, with accompany-ing chimes.

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

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




Figure 3-13. Threshold, Aisle, andBaggage Lights

Figure 3-12. Reading Lights

Figure 3-11. Passenger Warning Sign




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

Passenger Oxygen SwitchWhen oxygen flows into the passenger oxy-gen system supply line, a pressure-sensitiveswitch in the line closes a circuit to illuminatethe green PASS OXYGEN ON annunciatoron the caution-advisory annunciator panel.On series beginning with 1979 models, thisswitch will also cause the cabin lights, thevestibule light, and the baggage compartmentlight to illuminate in the full-bright mode, re-gardless of the position of the cabin lightsswitch.

EXTERIOR LIGHTS

LANDING LIGHTSTwo sealed-beam landing lights are mountedon the nose gear (Figure 3-14). An individualcircuit-breaker switch in the lighting group onthe pilot’s right subpanel controls each light.The switches are labeled LANDING and ei-ther LEFT or RIGHT.

TAXI LIGHTThe single, sealed-beam taxi light is mountedon the nose gear just below the landing lights.The control circuit-breaker switch is on thepilot’s right subpanel and is labeled TAXI.

The following described lights are shown inFigure 3-15.

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

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

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

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

STROBE LIGHTSA strobe light is installed in each wingtip andalso in the tip of the tail cone. Control forthese lights is incorporated into a switch on theright of the pilot’s right subpanel and is labeledSTROBE.

TAIL FLOODLIGHTSFloodlights, which may be installed on theunderside of the horizontal stabilizer, lightthe identification on the vertical stabilizer.Control is with a switch labeled TAIL FLOOD-LIGHT located on the overhead panel, or onthe pilot’s right subpanel.

Figure 3-14. Landing and Taxi Lights

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

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




Figure 3-17. Under Step LightingFigure 3-16. Airstair Floodlight

Figure 3-15. Exterior Lights

WING ICE LIGHTS

NAVIGATION LIGHTSRECOGNITION LIGHTS BEACON LIGHT

STROBE LIGHTS TAIL FLOODLIGHTSBEACON LIGHT

NAVIGATION LIGHT

4-i

CHAPTER 4MASTER WARNING SYSTEM

CONTENTS

Page

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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




4-iii

ILLUSTRATIONS

Figure Title Page

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

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

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

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

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

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

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

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

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




4-v

TABLES

Table Title Page

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

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

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

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

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

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




INTRODUCTIONThe master warning system consists of a warning annunciator panel with red readouts cen-trally located in the glareshield, a caution-advisory annunciator panel with amber and greenreadouts located on the center subpanel, and two flasher lights in front of each pilot on theglareshield (one labeled MASTER WARNING (red) and the other MASTER CAUTION(amber). Adjacent to the warning annunciator panel on the glareshield is a PRESS TO TESTswitch, which is used to illuminate the annunciator lights and flashers (Figure 4-1).

GENERALThe annunciators are word-readout types.When a fault condition covered by the an-nunciator system occurs, a signal is gener-a ted and the appropr ia te annuncia tor i silluminated. This action, in turn, illuminateseither the WARNING or CAUTION flasher.Super King Air 200 airplanes built before

1979 have 12 legends on the warning panel and30 legends on the caution-advisory panel.Super King Air 200 models built in 1979 andafter have 16 warning legends and 36 cau-tion-advisory legends. The B200 airplaneshave 20 legends on the warning panel and 36legends on the caution-advisory panel.

TEST

CHAPTER 4MASTER WARNING SYSTEM







PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

Figure 4-1. Component Locations




DIMThe warning annunciators (red), caution an-nunciators (amber), advisory annunciators(green), amber MASTER CAUTION flashersand red MASTER WARNING flashers fea-ture both a bright and a dim mode of illumi-nation intensity. (Prior to BB-1444, except1439, the MASTER WARNING flasher doesnot have a dim mode.) The dim mode will beselected automatically when all the followingconditions are met: a generator is on the line;the MASTER PANEL switch is on; the OVER-HEAD FLOODLIGHTS are off; the PILOTFLIGHT LIGHTS are on; and the ambientlight level in the cockpit (as sensed by a pho-toelectric cell located in the overhead lightcontrol panel) is below a preset value. Unlessall these conditions are met, the bright modewill be selected automatically.

TESTThe lamps in the annunciator system shouldbe tested before every flight, and at any othertime the integrity of a lamp is in question.Depressing the PRESS TO TEST button, lo-cated to the right of the warning annunciatorpanel in the glareshield, illuminates the an-nunciator lights, both MASTER WARNINGflashers, and both MASTER CAUTION flash-ers. (The yellow NO TRANSFER lights on thefuel panel are not included in this test, sincethey do not affect flashers when a NO TRANS-FER condition exists.) Any lamp that fails toilluminate when tested should be replaced.

GLARESHIELDFLASHERS

MASTER WARNING FLASHERSIf a fault requires the immediate attention andreaction of the pilot, the appropriate red warn-ing annunciator in the warning annunciatorpanel illuminates and both MASTER WARN-ING flashers begin flashing (Figure 4-2).Illuminated lenses in the warning annuncia-

tor panel will remain on until the fault is cor-rected. However, the MASTER WARNINGflashers can be extinguished by depressingthe face of e i ther MASTER WARNINGflasher, even if the fault is not corrected. Insuch a case, the MASTER WARNING flash-ers will again be activated if an additionalwarning annunciator illuminates. When awarning fault is corrected, the affected warn-ing annunciator will extinguish, but the MAS-TER WARNING f l a she r s wi l l con t inueflashing until one of the flashers is depressedto reset the circuit.

MASTER CAUTION FLASHERSWhen an annunciator-covered fault occursthat requires the pilot’s attention, the appro-priate amber caution annunciator in the cau-tion-advisory panel illuminates, and bothMASTER CAUTION flashers begin flashing(Figure 4-2). The flashing MASTER CAU-TION lights can be extinguished by pressingthe face of either of the flashing lights to resetthe circuit. Subsequently, when any other cau-tion annunciator illuminates, the MASTERCAUTION flashers will be activated again.Most illuminated caution annunciators on thecaution-advisory annunciator panel will re-main on until the fault condition is corrected,at which time they will extinguish. The MAS-TER CAUTION flashers will continue flash-ing until one of the flashers is depressed.

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

Figure 4-2. MASTER WARNING Flasher andMASTER CAUTION Flasher

WARNING ANNUNCIATORPANEL (RED)

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

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




* OPTIONAL EQUIPMENT

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

PRIOR TO BB-453

BB-453 AND AFTER* OPTIONAL EQUIPMENT

Figure 4-3. Warning Annunciator Panel—200 Aircraft




ILLUMINATION CAUSES—200Table 4-1 lists legend nomenclatures, colors,and causes for illumination in 200 aircraft.

ILLUMINATION CAUSES—B200Table 4-2 and 4-3 list legend nomenclatures, col-ors, and causes for illumination in B200 aircraft.

NOMENCLATURE COLOR CAUSE FOR ILLUMINATION

FIRE L ENG Red Fire in left engine compartment

ALT WARN Red Cabin altitude exceeds 12,500 feet

FIRE R ENG Red Fire in right engine compartment

L FUEL PRESS Red Fuel pressure failure on left side

INST INV Red The inverter selected is inoperative

R FUEL PRESS Red Fuel pressure failure on right side

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

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

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

L CHIP DETECT Red Contamination is detected in left engine oil

* A/P DISC Red Autopilot is disconnected

R CHIP DETECT Red Contamination is detected in right engine oil

* Optional equipment

Table 4-1. WARNING ANNUNCIATOR—200 AIRCRAFT

*OPTIONAL EQUIPMENT

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





L ENG FIRE Red Fire in left engine compartment

INVERTER Red The inverter selected is inoperative

DOOR UNLOCKED Red Cabin/cargo door open or not secure


R ENG FIRE Red Fire in right engine compartment



* L OIL PRESS Red Low oil pressure left engine

* L GEN OVHT Red Left generator temperature too high


* R GEN OVHT Red Right generator temperature too high

* R OIL PRESS Red Low oil pressure right engine


* A/P FAIL Red Autopilot is disconnected



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


L ENG FIRE Red Fire in left engine compartment

INVERTER Red The inverter selected is inoperative

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


R ENG FIRE Red Fire in right engine compartment



* L OIL PRESS Red Low oil pressure left engine

* L GEN OVHT Red Left generator temperature too high


* R GEN OVHT Red Right generator temperature too high

* R OIL PRESS Red Low oil pressure right engine

L CHIP DETECT Red Contamination is detected in left engine oil


* A/P FAIL Red Autopilot is disconnected


R CHIP DETECT Red Contamination is detected in right engine oil


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

CAUTION-ADVISORYANNUNCIATOR PANEL(AMBER/GREEN)

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

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

CAUTION SWITCH(200 MODELS ONLY)If the fault indicated by an illuminated cau-tion annunciator is not corrected, and pro-vided the MASTER CAUTION flasher is not

flashing, the pilot can still extinguish the an-nunciator by momentarily moving the spring-loaded CAUTION toggle switch (if installed)down to the OFF position, then releasing itto the center position. This action will ex-tinguish all illuminated caution annuncia-tors, and will illuminate the green CAUTLGND OFF advisory annunciator in the cau-tion advisory panel; this reminds the pilotthat an uncorrected fault condition exists,but that the caution legends have all beenextinguished. The annunciator(s) previouslyextinguished with the CAUTION switch canagain be illuminated anytime by momentar-ily moving the switch up to the ON position.This action will also extinguish the greenCAUT LGND OFF annunciator. If an addi-tional fault covered by the caution annunci-ators occurs after the caution legends havebeen extinguished with the CAUTION switch,the appropriate caution annunciator for thenew fault will illuminate, and all previouslyextinguished annunciators will again illu-minate. This switch is not installed in B200airplanes.

Figures 4-6, 4-7, 4-8 and 4-9 show typicalcau t ion advisory annunc ia tor pane ls in200/B200 aircraft.




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




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

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

*Optional Equipment

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

ILLUMINATION CAUSES—200 Table 4-4 is a listing of the warning legendnomenclatures, colors, and causes for illumi-

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




Table 4-4. CAUTION ADVISORY ANNUNCIATOR—200 AIRCRAFT


L DC GEN Amber Left generator off line

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

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

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

R DC GEN Amber Right generator off line

CABIN DOOR Amber Cabin door open or not secure

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

EXT PWR Amber External power connector is plugged in

BATTERY CHG Amber Excessive charge rate on the battery

DUCT OVERTEMP Amber Duct air too hot

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

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

FUEL CROSSFEED Green Crossfeed has been selected

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

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

L ICE VANE EXT Green Ice vane extended

* BRAKE DEICE ON Green Brake deice has been selected

LANDING LIGHT Green Landing lights on with landing gear up

PASS OXYGEN ON Green Oxygen is available to the passengers

R ICE VANE EXT Green Ice vane extended

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

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

CAUT LGND OFF Green Caution annunciator is turned off

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

R IGNITION ON Green Right starter/ignition switch is in the engine/ignition mode or right autoignition system is armed and right enginetorque is below 400 ft-lbs

* Optional Equipment

ILLUMINATION CAUSES—B200Tables 4-5 and 4-6 list the warning legendsnomenclatures, colors, and causes for illumi-

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




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



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

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




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



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

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

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


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

L ICE VANE EXT Green Ice vane extended

*BRAKE DEICE ON Green Brake deice has been selected

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

PASS OXY ON Green Oxygen is available to the passengers

R ICE VANE EXT Green Ice vane extended

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




R IGNITION ON Green Right starter/ignition switch is in the engine/ignition mode or right autoignition system is armed and right enginetorque is below 400 ft-lbs

* Optional Equipment† Not required when Type II synchrophaser is used




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



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



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


R CHIP DETECT Amber Metal contamination in the right engine oil is detected

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



R ENG ICE FAIL Amber Right engine anti-ice malfunction. Ice vane has not attained proper position

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

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


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

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

* BRAKE DEICE ON Green Brake deice has been selected

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

PASS OXY ON Green Oxygen is available to the passengers

ELEC HEAT ON Green Cabin electric heat is on

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

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




R IGNITION ON Green Right starter/ignition switch is in the engine/ignition modeor right autoignition system is armed and right engine torque is below 400 ft-lbs

* Optional Equipment

5-i

CHAPTER 5FUEL SYSTEM

CONTENTS

Page

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Fuel Crossfeed System ................................................................................................. 5-11

Fuel Gaging System ...................................................................................................... 5-12

Fueling .......................................................................................................................... 5-13

Antisiphon Valve ........................................................................................................... 5-13




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

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

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

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

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

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

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

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




5-iii

ILLUSTRATIONS

Figure Title Page

5-1 Fuel System Schematic for the Super King Air 200 and B200(BB-666 and Subsequent) ........................................................................................ 5-3

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

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

5-4 Fuel Pressure Warning Lights .................................................................................. 5-6

5-5 Fuel Control Panel.................................................................................................... 5-7

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

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

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

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

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

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

TABLES

Table Title Page

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




INTRODUCTIONThe airplane fuel system consists of two separate wing fuel systems connected with acommon crossfeed line and solenoid-operated crossfeed valve. Each wing system is fur-ther divided into a main and an auxiliary system. The main system employs a total of386 gallons of usable fuel; the auxiliary system, 158 gallons. At 6.7 pounds per gallon,these totals convert to 2,586 pounds in the main system and 1,058 pounds in the auxil-iary system. Total usable fuel is 544 gallons, or 3,644 pounds.

GENERALEach main fuel system is fueled through afiller opening on top of each wing at the outerwingtip. Fuel flows by gravity to the nacelletank. Each auxiliary fuel system is fueledthrough its own filler port. An antisiphonvalve at each filler point prevents fuel lossshould the filler cap be improperly secured orlost in flight. The auxiliary fuel system in

each wing consists of a rubber bladder-typetank mounted in each wing center sectionfrom which auxiliary fuel is transferred by ajet pump to the nacelle tank in the main fuelsystem. Although the main fuel system is fu-eled first, the fuel in the auxiliary tank is nor-mally exhausted before the fuel in the mainfuel system is automatically selected.

0

2

4 6

8

10

MAINFUEL

LBS X 100

CHAPTER 5FUEL SYSTEM




Additionally, the fuel system incorporates afully automatic vent system; a capacitancefuel gaging system on each side which providesseparate quantity readings for each main andauxiliary fuel system; and a fuel filter systemincorporating a filter bypass to enable fuelfeed to the engine in the event of filter icingor clogging.

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

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

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

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

The fuel system is covered in this chapter upto the high-pressure, engine-driven fuel pump,at which point fuel system operation becomesa function of the engine. Refer to Chapter 7,POWERPLANT, for additional information.

FUEL ROUTING INTO THEENGINEAfter exiting the main fuel system, fuel passesthrough the normally open firewall shutoffvalve. Just downstream of this valve is thelow-pressure, engine-driven boost pump. Fromthis pump, fuel is subsequently routed to thefirewall fuel filter and pressure switch, througha fuel heater which utilizes heat from engineoil, to the engine fuel pump, on to the fuel con-trol unit (FCU), and then through the fuel flowtransmitter (prior to BB-1401, the fuel flowtransmitter is upstream of the fuel heater).Fuel is then directed through the dual fuelmanifold to the fuel sprayer nozzles and intothe annular combustion chamber. Fuel is alsotaken from just downstream of the firewallfuel filter to supply the auxiliary tank trans-fer system with motive fuel flow.

MAJOR COMPONENTLOCATIONS ANDFUNCTIONS

MAIN AND AUXILIARY FUELSYSTEMSEach fuel system is divided into a main and anauxiliary fuel system, with a total usable fuelcapacity of 544 gallons. See Figure 5-1 forModels 200 and B200, SN BB-666 and sub-sequent. See Figure 5-2 for Model 200 priorto SN BB-666.

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

The filler cap for the main fuel system is lo-cated on top of the leading edge of the wing,near the tip; the cap has an antisiphon valve.

The auxiliary fuel system consists of a fueltank located in each wing center section, witha total usable capacity of 79 gallons per side.




5-3FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

L

F AUXILIARY

WING LEADING EDGE

WING LEADING EDGE

INTEGRAL (WET WING) BOXSECTION

BOXSECTION

DRAINVALVE

VENTFLOATVALVE

AIR INLET

DRAIN

RECESSED VENT

HEATED RAM VENT

FLAME ARRESTER

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

DRAIN**

TRANSFER JET PUMP

STRAINER, DRAIN,AND FUEL SWITCHF

FUEL CONTROL UNIT

ENGINE FUEL MANIFOLD

PRESSURE TANK

FIREWALL FUEL FILTER

ENGINE-DRIVEN BOOST PUMP (LP)DRAIN VALVE (FIREWALL)

FIREWALL SHUTOFF VALVE

STANDBY BOOST PUMP (30 PSI)

NACELLE TANKVENT FLOAT VALVE

CROSSFEED VALVE (NC)

P3 BLEED AIR LINE

ENGINE FUEL PUMP (HP)

FUEL HEATERAIR FILTER

*FUEL FLOW TRANSMITTER AND INDICATORLEFT FUEL PRESSURE ANNUNCIATOR PRESSURE SWITCH

FUEL CONTROL PURGE VALVEGRAVITY FLOW CHECK VALVE

STRAINER AND DEFUELING DRAIN VALVETRANSFER CONTROL MOTIVE FLOW VALVE (NC)

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

LEGEND

FUEL

FUEL AT STRAINER OR FILTER

FUEL UNDER LOW PUMP PRESSURE

HIGH-PRESSURE FUEL

FUEL CROSSFEED

FUEL RETUREN

GRAVITY FEED

FUEL VENT

FILLERPROBESSUCTION RELIEF VALVECHECK VALVEFUEL FLOW TRANSMITTERFUEL PRESSURE ANNUNCIATOR

F

L

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

5-4FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

F

F

L

AUXILIARY

WING LEADING EDGE WING LEADING EDGE

INTEGRAL (WET WING) BOX SECTION

DRAIN VALVE

VENT FLOAT VALVE

AIR INLET

DRAIN

BOX SECTION

RECESSED VENT

HEATED RAM VENT

FLAME ARRESTER

FROM FUEL NOZZLE MANIFOLDFUEL DRAIN COLLECTOR TANKFLOAT SWITCHFUEL DRAIN RETURN PUMP

FUEL HEATERPRESSURE SWITCH (10 PSI)

FIREWALL FUEL FILTER

FLAME ARRESTERENGINE-DRIVENBOOST PUMP (LP, 30 PSI)

FIREWALL SHUTOFF VALVE

DRAIN VALVE FIREWALLSTANDBY BOOST PUMP (30 PSI)NACELLE TANK

VENT FLOAT VALVE

CROSSFEED VALVE (NC)

ENGINE FUEL PUMP

FUEL CONTROL UNIT

(HP)

FUEL FLOW TRANSMITTER

STRAINER AND DEFUELING DRAIN VALVE

FUEL CONTROLPURGE VALVE

TRANSFER CONTROLMOTIVE FLOW VALVE (NC)

GRAVITY FLOW CHECK VALVEPRESSURE SWITCH

(6 PSI)

DRAIN

TRANSFER JET PUMP

AIR INLET (PRIOR TO SI 1021)

STRAINER, DRAIN,AND FUEL SWITCH

LEGEND

FUEL

FUEL AT STRAINER OR FILTER

FUEL UNDER LOW PUMP PRESSURE

HIGH-PRESSURE FUEL

FUEL CROSSFEED

FUEL RETUREN

GRAVITY FEED

FUEL VENT

FILLERPROBESSUCTION RELIEF VALVECHECK VALVEFUEL FLOW TRANSMITTERFUEL PRESSURE ANNUNCIATOR

F

L

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




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

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

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

AUXILIARY FUEL TRANSFERSYSTEMWhen auxiliary fuel is available, this systemautomatically transfers fuel from the auxiliarytank to the nacelle tank. No pilot action is in-volved. The jet transfer pump in the auxiliarytank operates on the venturi principle using thefuel and boost pump for motive flow. The en-gine-driven or electric low-pressure boostpump routes fuel through the normally-closedmotive flow valve, the jet pump, and into thenacelle tank. Fuel moving through the jet pumpventuri creates suction in the jet pump whichdraws fuel from the auxiliary tank.

During engine start, a 30- to 50-second timedelay is built into the automatic transfer sys-tem to allow all the fuel pressure to be used

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

Fuel pressure supplied by either the engine-driven boost pump or the electric standby boostpump (normally 25 to 30 psi) will open a fuelpressure-sensing switch and extinguish thered FUEL PRESS warning light (Figure 5-4).A minimum pressure of 10 ± 1 psi is requiredto extinguish the light. This same FUEL PRESSswitch will also send a signal to the auxiliaryfuel transfer printed circuit board indicatingthat motive flow is available for fuel transfer.If there is fuel in the auxiliary tank, this cir-cuit board will open the motive flow valvewithin 30 to 50 seconds. With the motive flowvalve now open, fuel is permitted to flowthrough the auxiliary transfer line. If the fuelpressure in this auxiliary transfer line is atleast 4 to 6 psi, a normally-closed pressureswitch will open and extinguish the amber NOTRANSFER light on the fuel panel. When theauxiliary tank empties, a float switch in the aux-iliary tank transmits a signal to close the mo-tive flow valve. This normally occurs after a30- to 60-second time delay, to prevent cy-cling of the motive flow valve due to sloshingfuel. This will not illuminate the NO TRANS-

WING LEADING EDGE(13 GALLONS)

INTEGRAL WET WING(35 GALLONS)

WING LEADING EDGE(40 GALLONS)

BOX SECTION(25 GALLONS)

BOX SECTION(23 GALLONS)

NACELLETANK

(57GALLONS)

AUXILIARY(79 GALLONS)

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

FER light because there is no more fuel left totransfer.

If the motive flow valve or its associated cir-cuitry should fail, it will go to the normally-closed position. Loss of motive flow pressurewith fuel remaining in the auxiliary tank willilluminate the amber NO TRANSFER lighton the applicable side of the fuel control panel.The motive flow valve may be manually en-ergized to the open position by placing theAUXILIARY TRANSFER switch, normally inthe AUTO position, to the OVERRIDE posi-tion (Figure 5-5). This procedure will bypassthe automatic feature in the auxiliary transfersystem and send DC power directly to the mo-tive flow valve.

On BB-32 and subsequent airplanes and on ear-lier models complying with Service Bulletin0703-286, power bypasses the AUXILIARYTRANSFER switch and the amber NO TRANS-FER light will not extinguish unless the motiveflow valve has opened (Figure 5-6).

On SN BB-2 through BB-31, selecting theOVERRIDE position of the switch takes powerfrom the NO TRANSFER light, causing it toextinguish. Even though the light is extin-guished, the valve may or may not open. Theauxiliary fuel level must be monitored to en-sure that it is decreasing.

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

FIREWALL SHUTOFF VALVEThe fuel system incorporates two in-line motor-driven firewall shutoff valves, one on each side.Each is controlled by a corresponding (guarded)switch near the circuit breakers on the fuel con-trol panel (Figure 5-5). The switches are plac-arded LEFT and RIGHT FIREWALL SHUTOFFVALVE, OPEN, and CLOSED. A red guard(guarded open) over each switch prevents in-advertent activation to the closed position.




Figure 5-4. Fuel Pressure Warning Lights




STANDBY PUMPON

OFF OFF

SEE MANUAL FOR FUEL CAPACITY

AUX TRANSFEROVERRIDE

AUTO

NO

TRANSFER

CROSSFEED FLOW

FUEL QUANTITYMAIN

AUXILIARY

ENGINE ENGINE

+

LEFT RIGHT

STANDBY PUMPON

OFF


AUTO

NO

TRANSFER

OPEN

CLOSED

FIREWALLSHUTOFF

VALVE

LEFT

RIGHT

FUEL SYSTEM CLOSED

ITT PROPTACH

TURBINETACH

FUELFLOW

OILPRESS

OILTEMP

5 5 5 5 5 5 5

5 5 5 5 5 5 5

TORQUE

FIREWALLVALVE

AUXTRANS

FER

AUXTRANS

FER

QTYIND

PRESSWARN

5 10 5 5 5 5

STANDBYPUMP

STANDBYPUMP

CROSSFEED

QTYIND

FIREWALLVALVE

5 5 5 10 5

PRESSWARN

RIGHTLEFT

OPEN

ENGINE INSTRUMENTS

FIREWALLSHUTOFF

VALVE

NO 4

BUSFEEDERS

BUSFEEDERS

NO 3

50

50

RIGHT

PROPDEICE

LEFT

25

25

CONTROL

MOTOR

PROPDEICE

FLAP

20

5

GOV

CONTROL

PROP

5

5

RIGHT

IGNITORPOWER

LEFT

5

5

RIGHT

STARTCONTROL

LEFT

5

5

NO 4

NO 3

50

50

LBS X 100QTY0

6

4

122

10

8

14LBS X 100

QTY0

6

4

122

10

8

14

MAIN TANKONLY

MAIN TANKONLY

FUEL FUEL

STANDBY PUMPON

OFF OFF

SEE MANUAL FOR FUEL CAPACITY


AUTO

NO

TRANSFER

CROSSFEED FLOWENGINE ENGINE

FUEL QUANTITYMAIN

AUXILIARY

+

LEFT RIGHT

STANDBY PUMPON

OFF


AUTO

NO

TRANSFER

OPEN

CLOSED


FUEL SYSTEM CIRCUIT BREAKERS

CLOSED

FIREWALLVALVE

AUXTRANS

FER

AUXTRANS

FER

QTYIND

PRESSWARN

STANDBYPUMP

STANDBYPUMP

CROSSFEED

QTYIND

FIREWALLVALVE

5 10 5 5 5 5 5 5 5 10 5

PRESSWARN

RIGHTLEFT

OPEN


BUS FEEDERS

50 50

RIGHT

PROP DEICE

LEFT

25 25

CONTROL

FLAP

205

GOV

PROP

5

RIGHT

IGNITOR

LEFT

5 5

RIGHT

CONTROL

LEFT

5 5

NO 4NO 3

50 50

LBS X 100QTY0

6

4

122

10

8

14LBS X 100

QTY0

6

4

122

10

8

14

MAIN TANKONLY

MAIN TANKONLY

START

PROP PROP

5

FUEL FUEL

MOTOR CONTROL POWER

PRIOR TO BB-1486, EXCEPT BB-1484

BB-1484, BB-1486 AND AFTER

PARKING BRAKEOFF

COLLINS

PARKING BRAKEOFF

COLLINS

Figure 5-5. Fuel Control Panel

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

ENGINE-DRIVEN BOOST PUMPThe low-pressure, engine-driven boost pumpis mounted on a drive pad on the aft accessorysection of the engine. The boost pump deliv-ers low-pressure fuel to the engine high-pres-sure fuel pump, thus preventing cavitation.The boost pump is protected against contam-ination by a strainer, and has an operating ca-pacity of 1,250 pph at a pressure of 25 to 30psi. Since it is engine driven, the pump oper-ates any time the gas generator (N1) is turn-ing and p rov ides su ff i c i en t fue l t o t hehigh-pressure pump for all flight conditions.An exception exists with aviation gasolinewhere flight above 20,000 feet altitude re-quires both standby boost pumps to be oper-ational and crossfeed to be operational.

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

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

When using aviation gas in climbs above20,000 feet, the first indication of insufficientfuel pressure will be an intermittent flicker ofthe red FUEL PRESS lights. Fuel flow andtorque may also indicate wide fluctuation.These conditions may be eliminated by acti-vation of the standby pumps.

CAUTION




AUXTRANSFER

L FUEL PRESS LOL FUEL PRESS LO

AUXTRANSFER

PCB

FROMCROSSFEED

SWITCH

MOTIVE FLOWVALVE

BOOST PUMPPRESSURESWITCH

FROM ENGINEDRIVEN BOOSTPUMP

PRESSUREWARNING

JET TRANSFERPUMP

FROMAUX TANK

TONACELLE TANK

CB

CB

AUX TRANSFERSWITCH

OVER RIDE

AUTO

NOT EMPTY

FLOAT SWITCH

EMPTY

MOTIVE FLOWPRESSURE SWITCH

TOENGINE

N.C.

Figure 5-6. Auxiliary Fuel Transfer System

STANDBY BOOST PUMPAn electrically-driven, low-pressure standbyboost pump located in the bottom of each na-celle tank performs three functions:

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

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

3. It is used during crossfeed operation.

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

Electrical power for standby pump operation iscontrolled by lever-lock switches on the fuelcontrol panel (Figure 5-5) and DC power issupplied from the dual fed buses (prior to BB-1098, except 1096 the standby boost pumps arealso powered by the hot battery bus). Theswi tches a re labe led STANDBY PUMPON–OFF. With the master switch on, power issupplied from the No. 3 (left) or No. 4 (right)bus feeders through the STANDBY PUMP cir-cuit breakers on the fuel control panel to thepumps.

Prior to BB-1098, except 1096, battery powerfrom the hot battery bus is also available forstandby boost pump operation. Fuses locatedin the right wing center section adjacent to thebattery box protect this circuit. These circuitsuse diodes to prevent failure of one circuitfrom disabling the other circuit. During shut-down, both STANDBY PUMP switches andthe CROSSFEED FLOW switch must be po-sitioned to OFF to prevent battery discharge.

FIREWALL FUEL FILTERFuel is filtered through a firewall-mounted 20-micron filter, which incorporates an internalbypass. The bypass opens to permit uninter-rupted fuel supply to the engine in case of fil-ter icing or blockage. In addition, a screenstrainer is located at each tank outlet beforefuel reaches the fuel boost and auxiliary trans-fer pumps.

LOW FUEL PRESSURESWITCHMounted on top of the firewall fuel filter is afuel pressure-sensing switch. In the event ofan engine-driven boost pump failure or anyother failure resulting in low pressure in thefuel line, the respective fuel pressure switchwill close, causing the red FUEL PRESS lighton the warning annunciator panel to illuminate(Figure 5-4). This light illuminates any timepressure decreases below 10 ± 1 psi. The lightwill normally be extinguished by switching onthe standby boost pump on that side.

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

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

• Prior to BB-225 by 26-volt AC power

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

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

With BB-1401 and subsequent, the transmit-ters were moved downstream of the fuel con-t ro l un i t t o on ly ind ica te fue l u sed fo rcombustion. Prior to BB-1401, the fuel flowtransmitters were installed in the fuel line up-stream of the fuel heater and were affected byFCU fuel purge flow.




FUEL HEATERFuel is heated prior to entering the fuel controlunit by an oil-to-fuel heat exchanger. An engineoil line is in proximity with the fuel line and,through conduction, a heat transfer occurs. Thepurpose of heating the fuel is to remove any iceformation which may have occurred or pre-clude any ice from forming, and which may re-sult in fuel blockage at the fuel control unit (seeLIMITATIONS at the end of this chapter). Thefuel heater is thermostatically controlled tomaintain a fuel temperature of 70° to 90°F undernormal conditions. If the fuel temperature risesabove 90°, the fuel will automatically bypass thefuel heater. If the fuel is extremely cold, and theoil temperature is too low, the unit may not becapable of preventing icing in the FCU. The oilvs. fuel temperature graph in the LIMITATIONSsection will specify under what conditions icingmay occur. The fuel heater is automatic and re-quires no pilot action.

HIGH-PRESSURE ENGINEFUEL PUMPThe high-pressure engine fuel pump is enginedriven and is mounted on the accessory drivein conjunction with the fuel control unit. Thisgear-type pump supplies the fuel pressureneeded for a proper spray pattern in the com-bustion chamber. Failure of this pump resultsin an immediate flameout.

FUEL MANIFOLDCLEARING

FUEL PURGE SYSTEM

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

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

In normal operation, the P3 air generated by theengine is held within the purge tank by theinput check valve and fuel pressure whichholds the dump valve shuttle closed. When theengine is shut down, fuel pressure on the dumpvalve shuttle decreases. The shuttle valve openswhen P3 pressure is greater than fuel manifoldpressure. This allows P3 air to enter the fuelmanifolds, forcing the remaining fuel in themanifolds into the burner can. Since combus-tion has not ceased, this small amount of fuelfrom the manifolds is now burned, which mayresult in a small rise in ITT and N1. Refer toChapter 7, POWERPLANT, for additional in-formation on the purge system.




Figure 5-7. Fuel Flow Gages

FUEL DRAIN COLLECTORSYSTEM

(BB-2 through BB-665)Fuel from the engine flow divider drains intoa collector tank mounted below the aft engineaccessory section. An internal float switchactuates an electric pump, which delivers thefuel to the fuel purge line just aft of the fuelpurge shutoff valve. A check valve in the lineprevents the backflow of fuel during enginepurging. A vent line plumbed from the top ofthe collector tank is routed through an in-lineflame arrester and downward to a drain mani-fold on the underside of the nacelle (Figure 5-2). The system receives power from the No. 1dual-fed bus. This fuel delivery is completelyautomatic.

FUEL CROSSFEED SYSTEMCrossfeeding is conducted only during sin-gle-engine operation, when it may be nec-

essary to supply fuel to the operative enginefrom the fuel system on the opposite side(Figure 5-9). The simplified crossfeed con-t ro l pos i t ions are labeled CROSSFEEDFLOW and OFF (Figure 5-5). The STANDBYPUMP switches must be positioned to OFFfor crossfeeding.

The auxiliary transfer switch must bepositioned to the AUTO position onthe side being crossfed. If auxiliaryfuel supply is required from the in-operative engine side, the firewallvalve must be opened provided en-gine shutdown was not due to a fuelleak or fire.

Movement of the CROSSFEED switch LEFTor RIGHT, will directly affect four circuitsand may indirectly cause a fifth indication:

CAUTION




FUELPUMP

BOOST PUMP PRESSURE

HIGH-PRESSURE FUEL

ENGINE BLEED AIR

LEGEND

FILTER

CHECKVALVE

PURGE TANK

CHECKVALVE

PRIOR TO BB-1401, FUEL FLOWTRANSMITTER IS UPSTREAM OF FUEL HEATER

TOPNEUMATICS

TOFLOW

PACKAGE

PURGELINE

DUMPVALVE

POPPETVALVE

FROMP3 AIR

ENGINEFUELCONTROL

UNIT

FUELHEAT

FIRESEAL

FUELFLOW *

*

Figure 5-8. Fuel Purge System

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

2. The CROSSFEED valve will open.

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

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

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

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

Illumination of the green FUEL CROSSFEEDlight on the caution/advisory panel indicatescrossfeed has been selected, not that the cross-feed valve has moved. The Before EngineStarting checklist contains a crossfeed test toensure operation of this valve. During this test,the pilot should ensure that both red FUELPRESSure lights extinguish once the CROSS-FEED switch is moved LEFT or RIGHT, in-dicating the valve has opened.

FUEL GAGING SYSTEMA capacitance-type fuel gaging system mon-itors fuel quantity in either the main or aux-iliary fuel system for each side. Two fuelgages, one for each wing fuel system, are onthe fuel control panel (Figure 5-5).




MOTIVE FLOWVALVE

MOTIVE FLOWVALVE

CROSSFEEDVALVE

STANDBY BOOST PUMP


FIREWALLSHUTOFF

VALVE

LOW PRESSUREENGINE-DRIVEN

FUEL PUMP

Figure 5-9. Fuel Crossfeed System

Quantity is read directly in pounds. An errorof 3% maximum may be encountered in the sys-tem. The readings are compensated for densitychanges caused by temperature variations.

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

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

Power is supplied through the capacitanceprobes to the quantity indicator.

FUELINGFuels and fueling considerations are coveredin LIMITATIONS.

The procedure for blending anti-icing additivewith fuel is accomplished during fueling, and

is covered in the NORMAL PROCEDURESsection of the Flight Manual.

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

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

An air inlet at the wingtip vents the integral(wet cell) tank, the auxiliary tank and for BB-479 and subsequent, the nacelle tank. Prior toBB-479 a dedicated nacelle air inlet was lo-cated at the trailing edge of the wing behindthe nacelle.




Figure 5-10. Fuel Crossfeed Advisory Light

FUEL DRAINSThere are five sump drains and a firewall fil-ter drain in each wing. Drain locations areshown in Table 5-1.

LIMITATIONS

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

Aviation gasoline Grades 80 Red (formerly80/87), 91/98, 100LL Blue (same as 100LGreen in some countries), 100 Green (formerly100/130), and 115/145 Purple are emergencyfuels. Emergency fuels may be mixed withrecommended fuels in any ratio. However,when aviation gasoline is used, operation islimited to 150 hours between engine over-hauls. The number of gallons taken aboard for

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

The pilot must be familiar with the consump-tion rate of his airplane and record the num-ber of gallons taken aboard for each engine.

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

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

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

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




Table 5-1. DRAIN LOCATIONS

DRAINS LOCATION

Leading edge tank Outboard of nacelle underside of wing

Integral tank Underside of wing forward of aileron

Firewall fuel filter Underside of cowling forward of firewall

Sump strainer Bottom center of nacelle forward of the wheel well

Gravity feed line Outboard side of nacelleAft of wheel well (Prior to BB-1193)

Auxiliary tank At wingroot just forward of the flap

NOTEWindmi l l i ng t ime need no t becharged against this time limit.

The maximum allowable fuel imbalance is1,000 pounds. Check the Aircraft FlightManual Supplements for maximum imbalanceduring autopilot operation.

APPROVED FUEL ADDITIVEAnt i - i c i ng add i t i ve con fo rming t oSpecification MIL-I-27686 or MIL-I-85470are the only approved fuel additives.

Engine oil is used to heat the fuel on enteringthe fuel control. Since no temperature mea-surement is available for the fuel at this point,it must be assumed to be the same as the OAT.Figure 5-11 is supplied for use as a guide inpreflight planning, based on known or forecastoperating conditions, to allow the operator tobecome aware of operating temperatures atwhich icing of the fuel control could occur. Ifoil temperature versus OAT indicates that iceformation could occur during takeoff or inflight, anti-icing additive per MIL-I-27686

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

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

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

CAUTION

CAUTION




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

FUELING CONSIDERATIONSDo not put any fuel into the auxiliary tanks un-less the main tanks are full.

The airplane must be statically grounded to theservicing unit, and the servicing unit mustalso be grounded.

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

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

ZERO-FUEL WEIGHTThe maximum zero-fuel weight of the SuperKing Air 200 is 10,400 pounds. The maxi-mum zero-fuel weight of the B200 is 11,000pounds.




7-i

CHAPTER 7POWERPLANT

CONTENTS

Page

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

OVERVIEW ............................................................................................................................ 7-1

ENGINE .................................................................................................................................. 7-2

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

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

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

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

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

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

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

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

Indication ......................................................................................................................... 7-8

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

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

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

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

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

Fuel System Operation..................................................................................................... 7-9

Fuel Control Unit (FCU)................................................................................................ 7-13

ENGINE IGNITION SYSTEM ............................................................................................ 7-15

General........................................................................................................................... 7-15




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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Autofeathering ............................................................................................................... 7-32

Unfeathering .................................................................................................................. 7-34




7-iii

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

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

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

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

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

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

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

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

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




7-v

ILLUSTRATIONS

Figure Title Page

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

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

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

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

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

7-6 Oil Pressure/Temperature Gages .............................................................................. 7-8

7-7 Chip Detection Lights .............................................................................................. 7-9

7-8 Oil System Schematic ............................................................................................ 7-10

7-9 Fuel Low-Pressure Lights ...................................................................................... 7-11

7-10 Fuel Flow Gages..................................................................................................... 7-11

7-11 Fuel Schematic ....................................................................................................... 7-12

7-12 Simplified Fuel Control Schematic........................................................................ 7-14

7-13 Engine Start and Ignition Switches ........................................................................ 7-15

7-14 Engine Autoignition Switches................................................................................ 7-15

7-15 Ignition System Schematic..................................................................................... 7-16

7-16 Propellers................................................................................................................ 7-17

7-17 PROP GOV TEST Switch ..................................................................................... 7-19

7-18 Propeller Governor Test Schematic ....................................................................... 7-20

7-19 Propeller Onspeed Schematic ................................................................................ 7-22

7-20 Propeller Overspeed Schematic ............................................................................. 7-22

7-21 Propeller Underspeed Schematic ........................................................................... 7-23

7-22 Powerplant Control Levers .................................................................................... 7-24

7-23 Beta and Reverse Control....................................................................................... 7-25




7-24 Propeller Control Lever.......................................................................................... 7-26

7-25 Friction Control Knobs .......................................................................................... 7-26

7-26 ITT Gages............................................................................................................... 7-27

7-27 Torque Gages ......................................................................................................... 7-27

7-28 Propeller RPM Gages............................................................................................. 7-28

7-29 Engine RPM Gages ................................................................................................ 7-28

7-30 Propeller Synchroscope and Switches (Type II) .................................................... 7-28

7-31 Type II System Schematic...................................................................................... 7-29

7-32 Type I System Schematic ....................................................................................... 7-30

7-33 Propeller Synchroscope and Switch (Type I)......................................................... 7-31

7-34 Sync Light .............................................................................................................. 7-31

7-35 AUTOFEATHER Switch ....................................................................................... 7-31

7-36 Autofeather Lights ................................................................................................. 7-32

7-37 Autofeather System Schematic (Both Power Levers at Approximately90% N1; Right Engine has Failed) ......................................................................... 7-33

7-38 Autofeather Test Schematic (Left Power Lever Below 200 ft-lb;Right Power Lever Above 400 ft-lb) ..................................................................... 7-34

7-vi FOR TRAINING PURPOSES ONLY



7-vii

TABLES

Table Title Page

7-1 Engine Operating Limits(PT6A-42 Engine BB-1439, 1444 and Subsequent) .............................................. 7-35

7-2 Engine Operating Limits (PT6A-42 Engine Prior to BB-1439,1444 and Subsequent) ............................................................................................ 7-35

7-3 Engine Operating Limits (PT6A-41 Engine) ......................................................... 7-36

7-4 Powerplant Instrument Markings........................................................................... 7-38




INTRODUCTIONThis chapter deals with the powerplant of the Super King Air 200. All values, such as forpressures, temperatures, rpm, and power are used for illustrative meanings only. Actualvalues must be determined from the appropriate sections of the approved flight manual.

Information in this chapter must not be construed as being equal to or superseding anyinformation issued by or on behalf of the various manufacturers or the Federal AviationAdministration.

OVERVIEWThe Super King Air 200 (Figure 7-1) is pow-ered by two wing-mounted, turboprop en-gines, manufactured by Pratt and WhitneyAircraft of Canada Limited, a Division ofUnited Technologies. The engines drive three-or four-blade, constant-speed propellers which

incorporate full feathering and full reversingcapabilities in addition to ground fine/Betamode control for ground operation. On theground, the propeller is feathered when the en-gine is shut down and unfeathered when theengine is restarted.

#1 DCGEN

CHAPTER 7POWERPLANT




ENGINE

GENERALThe engines used on the Super King Air 200are designated PT6A-41, while the B200 usesPT6A-42.

The PT6A (Figure 7-2) is a free-turbine, re-verse-flow, lightweight turboprop engine, ca-pable of developing 850-shaft horsepower(903 equivalent shaft horsepower [ESHP]).

PT6 engine development began about 1960.The first certificated engine, the PT6A-6, en-tered service in 1962, rated at 450-shaft horse-power. Since then, the output of the PT6A hasalmost tripled (with no apparent outwardchanges), to 1,300-shaft horsepower on thePT6A-68A.

MAJOR SECTIONSFor the purpose of this chapter, the engine(Figure 7-3) is divided into seven major sections.

1. Air Intake Section

2. Compressor Section

3. Combustion Section

4. Turbine Section

5. Exhaust Section

6. Reduction Gear Section

7. Accessory Drive Section




Figure 7-1. Super King Air 200




FUEL PUMP/FCU

TACHOMETER-GENERATOR(NG)

OIL SCAVENGEPUMPS AND

FUEL BOOST PUMP

OPTIONALACCESSORY

DRIVE

STARTER-GENERATOR

PROPELLEROVERSPEEDGOVERNOR

PROPELLERGOVERNOR

TACHOMETER-GENERATOR

(NF)

AFT

FRONT

Figure 7-2. PT6A Engine

7-4FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

Figure 7-3. Engine Cutaway




Air Intake SectionThe compressor air intake consists of a circularscreen-covered, aluminum casting. Air is di-rected to the air intake by the nacelle air scoopon the lower side of the nacelle. The functionof the air intake section is to direct airflow tothe gas generator compressor.

Compressor SectionThis section consists of a four-stage compres-sor assembly, made up of three axial stages andone centrifugal stage. The function of the com-pressor is to compress and supply air for com-bustion, combustion cooling, pressurizationand pneumatic services, compressor bleed valveoperation, and bearing sealing and cooling.

Compressor Bleed ValvesAt low N1 rpm, the compressor axial stagesproduce more compressed air than the cen-trifugal stage can use. Compressor bleed valvescompensate for this excess airflow at low rpm

by overboarding, or bleeding axial stage air toreduce backpressure on the centrifugal stage(Figure 7-4). This pressure relief helps preventcompressor stall of the centrifugal stage.

The compressor bleed valves, one on eachside of the compressor, are pneumatic pis-tons, which reference the pressure differentialbetween the axial and centrifugal stages.Looking forward, the low-pressure valve is lo-cated at the 9 o’clock position and the highpressure at 3 o’clock. The function of thesevalves is to prevent compressor stalls andsurges in the low N1 rpm range.

At low N1 rpm, both valves are in the open po-sition. At takeoff and cruise N1 rpm, above ap-proximately 90%, both bleed valves will beclosed. If both compressor bleed valves wereto stick closed below approximately 90% N1,a compressor stall would result.

If one or both valves were to stick open, theITT would increase and torque decrease whileN1 rpm remained constant.

CONTROL PRESSURE

PISTON DAMPER(SPRING LOAD)

SLEEVE

AMBIENT PRESSURE

FINALORIFICE

PRIMARYORIFICE

DELIVERYAIR PASSAGE

P3

P2.5

Figure 7-4. Compressor Bleed Valves

Combustion SectionThe PT6 engine utilizes an annular combustionchamber. Two, high-energy igniter plugs areinstalled in the combustion chamber, as wellas 14 equally-spaced simplex fuel nozzles.

Turbine SectionThe PT6A uses three reaction turbines: a free,two-stage axial propeller (power) turbine anda single-stage compressor turbine. The two-stage power turbine extracts energy from thecombustion gases to drive the propeller and itsaccessories through the planetary reductiongears. This combination is defined as NP. Thesingle-stage compressor turbine extracts en-ergy from the combustion gases to drive thegas generator compressor and the accessorygear section. This combination is defined asN1.

Exhaust SectionThis section is located immediately aft of thereduction gear section and it consists of an an-nular exit plenum, a heat-resistant cone, andtwo exhaust outlets at the 9 o’clock and 3o’clock positions.

Reduction Gear SectionThe reduction gear section at the front of theengine is a two-stage, planetary type. The pri-mary function of the reduction gear section isto reduce the high rpm of the free turbine tothe value required for propeller operation.

The reduction gear section is also used fortorquemeter operation and includes drive sec-tions for the propeller governor (with fueltopping governor sensing), the propeller over-speed governor, and a propeller tachgenera-tor (Figure 7-2).

Accessory SectionThe accessory drive section forms the aft por-tion of the engine. The accessory section isdriven by the compressor turbine through ashaft that extends aft through the oil tank tothe accessory gearbox.

The function of the accessory section is todrive the engine and airplane accessories,which include:

• Fuel control unit (FCU) and high-pres-sure fuel pump

• Lubricating pump/scavenge pumps

• N1 tachgenerator

• DC starter-generator

• Refrigerant compressor (right engineonly)

• Low-pressure fuel boost pump

Other drive pads are provided for optional op-erator equipment (Figure 7-2).

OPERATING PRINCIPLESWhen the engine is rotating (Figure 7-5), air isinducted through the nacelle air scoop to the en-gine air intake. Airflow is turned 180° in a for-ward direction and is then progressivelyincreased in pressure by a three-stage axial-flow and single-stage centrifugal-flow com-pressor. It is then directed forward throughdiffuser ducts towards the forward side of thecombustion chamber. The airflow is again turned180° and enters the combustion chamber, wheremetered fuel is added to the air by 14 fuel spraynozzles. Two high-energy igniter plugs ignitethe gas mixture. The expanding gases moverearward through the combustion chamber andturn 180° forward to enter the turbine section.The compressor turbine extracts sufficient en-ergy from the expanding gases to drive the four-stage compressor and the accessory gear section.The remaining two stages of the free power tur-bine extract the maximum amount of the re-maining energy from the combustion gases todrive the propeller and the propeller accessoriesthrough the reduction gearbox. The two-stagepower turbine is a free turbine and is only aero-dynamically (not mechanically) connected to thegas generator. The gases from the turbine con-tinue forward into an exhaust plenum wherethey are directed to the atmosphere by exhaustnozzles at the 9 o’clock and 3 o’clock positionson the exhaust section of the engine.







POWER SECTION

COMPRESSOR SECTION

POWER SECTION

MODULE 1

MODULE 2

GAS GENERATOR SECTION

Figure 7-5. Engine Gas Flow and Stations

ENGINE LUBRICATIONSYSTEM

GENERALThe engine lubrication system is a completelyself-contained and fully automatic system. Itprovides for cooling and lubrication of theengine bearings and the reduction and acces-sory drive gears, and for operation of the pro-peller control system, the torquemeter system,the torque limiter, and the fuel heater system.

The engine oil system is a dry-sump systemconsisting of pressure, scavenge, and cen-trifugal air breather systems.

OIL TANKThe oil tank forms an integral part of the en-gine, located between the aft end of the com-pressor air inlet and the forward end of theaccessory gearbox.

A filler and dipstick are located at the 11o’clock position on the accessory case. The oiltank is vented to a centrifugal breather to pro-vide for air-oil separation.

PUMPSThe oil pumps consist of one pressure ele-ment and four scavenge elements. The pres-sure pump supplies lubrication pressure to thebearings and the accessory system drive gears.In addition, the pressure pump supplies oil tothe propeller control system, the torquemetersystem, reduction gears and the torque limiter.

OIL COOLERAn oil radiator is located inside the lower na-celle for oil cooling. The oil cooling systemis fully automatic and uses a thermal sensorto control the position of a door that regulatesthe flow of air through the oil cooler.

INDICATION

Engine Oil PressureEngine oil pressure is sensed by a transmitterin the pressure pump outlet line and suppliedto a combination, pressure-temperature gage(Figure 7-6) on the engine instrument panel.The oil pressure system requires DC power.

Engine Oil TemperatureOil temperature is sensed by a resistance bulband transmitted to the same combination pres-sure/temperature gage (Figure 7-6) on the en-gine instrument panel. The power supply forthe gage is from the DC power system.

Chip DetectionFor BB-1439, 1444 and subsequent, the cau-tion annunciator panel contains two amberlights marked L CHIP DETECT and R CHIPDETECT (Figure 7-7). Prior to BB-1444, ex-cept 1439, these are red lights on the warningannunciator panel. They are operated by amagnetic chip detector located at the bottomof each reduction gearbox.

When either light illuminates, it indicates thatferrous metal particles in the oil have been at-tracted to the chip detector magnets.




Figure 7-6. Oil Pressure/TemperatureGages

FUEL HEATEROil scavenged from the accessory gearcase isdirected through an oil-to-fuel heater prior toits return to the oil tank.

OPERATIONWhen the engine is running, the oil pressurepump (Figure 7-8) draws oil from the tank, de-velops a higher pressure with the oil, and di-rects pressure oil through various filters tothe engine bearings, the accessory and re-duction drive gears, the propeller governor,and the engine torquemeter system. Oil pres-sure is regulated and limited by a relief valve.Oil pressure and temperature are sensed andtransmitted to the cockpit gages. All oil isscavenged to the accessory gearcase except thereduction gearcase oil, which goes directly tothe oil cooler. A screened scavenge pump re-turns the gearcase oil to the tank through theoil-fuel heater; another scavenge pump scav-enges oil from the reduction gearcase and re-turns this oil to the tank through the oil cooler.

ENGINE FUEL SYSTEMGENERALThe engine fuel system consists of an oil-to-fuel heater, an engine-driven, high-pressurefuel pump, an engine-driven, low-pressure boost pump, a fuel control unit (FCU), a flow

divider, and two fuel manifolds each withseven simplex fuel nozzles.

INDICATION

Fuel PressureThe warning annunciator panel red lightsmarked L FUEL PRESS and R FUEL PRESS(Figure 7-9) are operated by pressure switchesthat sense outlet pressure at the engine-drivenboost (LP) pump. The lights will come on toindicate abnormally low (10 ± 1 psi) fuel pres-sure to the (HP) engine pump.

Fuel FlowFuel flow information is sensed by a transmit-ter in the engine fuel supply line and suppliedto the fuel flow gages (Figure 7-10) on the cen-ter instrument panel. For BB-225 airplanes andsubsequent, they are DC powered. Prior to BB-225, these fuel flow gages are AC powered.

FUEL SYSTEM OPERATION

The fuel control system for PT6A engines isessentially a fuel governor that increases ordecreases fuel flow to the engine to maintainselected engine operating speeds. At firstglance, the system may appear quite com-plicated. The engine fuel control system con-sists of the main components shown in Figure7-11. They are the primary low-pressure boost





Figure 7-7. Chip Detection Lights

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TO OIL PRESSUREINDICATOR

TO OIL TEMPERATUREINDICATOR

TO OIL PRESSUREANNUNCIATOR

FROM COOLER

TO COOLER

OIL TANK BREATHEROIL DIPSTICK

DIVERTERVALVE

FUELHEATER

SCAVENGEPUMP

ACCESSORYGEARBOX DRAIN

BYPASSVALVE

OILTANK

OIL FILTER AND CHECK VALVE

BYPASS VALVE

OVERPRESSURERELIEF VALVE

TANK DRAINTORQUEMETER

& TORQUE LIMITER

TORQUEMETER PRESSURE

(INDICATOR)CHIP DETECTOR

OIL SUPPLYTO PROPELLER

PROPELLER GOVERNORAND BETA CONTROL

TORQUEMETER OILCONTROL VALVE

PRESSURE OIL

PROPELLER SUPPLY OIL

SCAVENGE OIL

BREATHER AIR

TORQUEMETER PRESSURE

LEGEND

PRESSUREREGULATING VALVE

* OPTION

*

Figure 7-8. Oil System Schematic

pump, oil-to-fuel heat exchanger, high-pres-sure fuel pump, fuel control unit, fuel cutoffvalve, fuel flow transmitter, flow divider, anddual fuel manifold with 14 simplex nozzles.

The low-pressure boost pump is engine-drivenand operates when the gas generator shaft(N1) is turning, to provide sufficient fuel headpressure to the high-pressure pump to main-tain proper cooling and lubrication. The oil-to-fuel heat exchanger uses warm engine oilto maintain a desired fuel temperature at thefuel pump inlet to prevent icing at the pumpfilter. This is done with automatic temperaturesensors and requires no action by the pilot.

Fuel enters the engine fuel system throughthe oil-to-fuel heat exchanger, and then flowsinto the high-pressure engine-driven fuelpump, and on into the fuel control unit (FCU).

The high-pressure fuel pump is an engine-driven gear-type pump with an inlet and out-

let filter. Flow rates and pressures will varywith gas generator (N1) rpm. Its primary pur-pose is to provide sufficient pressure at the fuelnozzles for a good spray pattern at all modesof engine operation. The high-pressure pumpsupplies fuel at approximately 800 psi to thefuel side of the FCU.

Two valves included in the FCU ensure con-sistent and cool engine starts. When the igni-tion or start system is energized, the purgevalve is electrically opened to clear the FCUof vapors and bubbles. The excess fuel flowsback to the nacelle fuel tank. The spill valve,referenced to atmospheric pressure, adjuststhe fuel flow for cooler high-altitude starts.

Between the FCU fuel valve and the enginecombustion chamber, and part of the FCU, aminimum pressurizing valve cuts off fuel flowduring starts until fuel pressure builds suffi-ciently to maintain a proper spray pattern inthe combustion chamber. About 70 psi is re-quired to open the minimum-pressurizingvalve. The engine-driven high-pressure fuelpump maintains this required pressure. If thepump should fail, the valve would close andthe engine would flame out.

Downstream from the minimum pressurizingvalve in the FCU is the fuel cutoff valve. Thecondition lever controls this valve, either openor closed. There is no intermediate position ofthis valve. For starting, fuel flows initiallythrough the flow divider valve to the primary




Figure 7-9. Fuel Low-Pressure Lights

Figure 7-10. Fuel Flow Gages

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FUEL PRESS

FIREWALLFUEL FILTER

ANDMANUAL

SHUT-OFF VALVE

ENGINEDRIVENBOOSTPUMP

TRANSDUCER

ENGINEDRIVEN

FUEL PUMP(HIGH

PRESSURE)

FUEL FLOWTRANSMITTER

FLOWDIVIDER

P3 PURGE**TANK

P3 AIR

P3 AIR

NP

FUELTOPPING

GOVERNOR

N1

OIL-TO-FUELHEAT

EXCHANGER

FUELCONTROL

UNIT

POWERAND

CONDITIONLEVERS

COCKPITGAGE

*

TOFUELTANK

PURGE LINE

* PRIOR TO BB-1401, FUEL FLOW TRANSMITTER LOCATED UPSTREAM OF OIL-TO-FUEL HEAT EXCHANGER

** PRIOR TO BB-666, FUEL DRAIN COLLECTOR TANK

Figure 7-11. Fuel Schematic




fuel spray nozzles in the combustion chamber.As the engine accelerates through approxi-mately 40% N1, fuel pressure is sufficient toopen the transfer valve to the secondary fuelnozzles. At this time all 14 nozzles are deliv-ering atomized fuel to the combustion cham-ber. This progressive sequence of primary andsecondary fuel nozzle operation providescooler starts. On engine startups, there is a def-inite surge in N1 speed when the secondary fuelnozzles cut in.

In order to improve cold weather starting, SB-3214 changed seven primary and seven sec-ondary nozz le s to 10 p r imary and foursecondary. At a later date, SB-3250 changedthe nozzles back to seven and seven, but witha different arrangement and an improvedburner can.

During engine shutdown on BB-666 and sub-sequent, any fuel left in the manifold is forcedout through the nozzles and into the combus-tion chamber by purge tank pressure. As thefuel is burned, a momentary surge in N1 rpmshould be observed. The entire operation is au-tomatic and requires no input from the crew.On BB-2 through BB-665, an EPA collectortank is used instead of the purge tank system.

FUEL CONTROL UNIT (FCU)

GeneralThe fuel control unit (Figure 7-12), which isnormally referred to as the FCU, has multiplefunctions, but its main purpose is to meter theproper fuel amount to the nozzles in all modesof engine operation. It is calibrated for start-ing flow rates, acceleration, and maximumpower. The FCU compares gas generator speed(N1) with the power lever setting and regulatesfuel to the engine fuel nozzles. The FCU alsosenses compressor section discharge pressure,compares it to rpm, and establishes accelera-tion and deceleration fuel flow limits.

Fuel flow to the engine is dependent on the po-sition of the fuel cutoff valve, which is man-ually operated by the condition lever in thecockpit. In addition, the minimum pressuriz-

ing valve prevents fuel flow to the engine untilthe fuel pressure has increased enough to en-sure proper atomization of the fuel at the noz-zles. Once the minimum pressure valve hasopened, fuel will flow to the flow divider andthe fuel nozzles.

Aside from opening and closing the fuel cutoffvalve, the condition lever adjusts N1 speed fromLOW IDLE to HIGH IDLE. The power lever,by adjusting the governor position in the FCU,adjusts the fuel-metering valve to allow moreor less fuel to the spray nozzles. In summary,the power lever controls fuel to the engine byadjusting the governor position, which in turnrepositions the fuel-metering valve in the FCU.

FCU OperationThe pneumatic section of the FCU determinesthe flow rate of fuel to the engine for all op-erations. The power levers control enginepower from idle through takeoff power by op-eration of the gas generator (N1) governor inthe FCU. Increasing N1 rpm results in in-creased engine power.

For explanation purposes, consider the N1 gov-ernor bellows as a diaphragm. P3 air is intro-duced into the bellows in a manner that sets upa differential pressure on each side of the di-aphragm. Therefore, any change in P3 pres-sure will move the diaphragm. When pressureis increased, the fuel-metering valve attachedto the bellows will move in an opening direc-tion to increase fuel flow and increase N1 rpm.

As P3 pressure decreases, fuel flow also de-creases which reduces the N1 rpm. The N1governor increases or decreases P3 pressure inthe bellows by varying the opening of relieforifices in the bellows.

The FCU controls engine power by maintain-ing the requested N1 rpm through the N1 gov-ernor. If actual N1 rpm is lower than the desiredsetting, the N1 governor closes the P3 orifice,allowing pressure to increase. As the pressureincreases, the diaphragm moves to open themetering valve, increasing fuel flow, which inturn increases N1 rpm to the speed requested by

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N1

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Figure 7-12. Simplified Fuel Control Schematic




the governor. When N1 rpm reaches the desiredspeed, the governor adjusts the P3 orifice to re-duce pneumatic pressure to match the fuel pres-sure required to maintain the desired N1 rpm.

The fuel topping (power turbine) governorprotects against power turbine overspeed. Ifan overspeed occurs, and the propeller goes be-yond 106% of the requested propeller rpm,the fuel topping governor vents air to reducefuel flow. Reducing fuel flow decreases N1speed and accordingly power turbine speed.With propellers in reverse, the fuel-toppinggovernor restricts fuel flow to approximately95% of the requested propeller rpm.

ENGINE IGNITIONSYSTEMGENERALThe engine ignition system is a high-energy,capacitance type consisting of a dual-circuitigniter box and two igniter plugs in the com-bustion chamber. The ignition system is di-vided into starting ignition and autoignition.

STARTING IGNITIONA three-position lever lock switch for each en-gine (Figure 7-13) controls this system. The

switch is located on the left switch panel. It hasthree marked positions: ON–OFF–STARTERONLY. The ON position (UP) is lever locked andit provides for engine cranking and ignition op-eration. The STARTER ONLY position is a mo-mentary (spring loaded to center hold down)position and it only provides for engine motor-ing. In this position, the igniters do not function.

AUTOIGNITIONThe autoignition system is controlled by atwo-position switch for each engine markedARM and OFF (Figure 7-14). Turning on anAUTO IGNITION switch arms the ignitercircuit to an engine torque switch that isnormally open when the engine is develop-ing more than 400 foot-pounds of torque. Thesystem must be armed prior to takeoff andfor all phases of flight, and it should beturned off only after landing. If engine torquedrops to 400 foot-pounds or less when theautoignition is armed, the ignition systemwill energize to prevent engine flameout ifthe loss of power was caused by a momen-tary fuel or air interruption.

INDICATIONGreen annunciator lights marked L and R IG-NITION ON tells the pilot that the igniters arereceiving power.

Figure 7-13. Engine Start and IgnitionSwitches

Figure 7-14. Engine Autoignition Switches

OPERATION

Starting IgnitionWhen DC power is available, turning on theignition and engine start switch (Figure 7-15)will apply DC power to the ignition ON light,FCU purge valve, and to the ignition exciter.The exciter, which operates at three cyclesper second, will apply high-energy power tothe igniter plugs in the combustion chamber.

AutoignitionWhen the AUTO IGNITION switch (Figure7-14) is at the ARM position, the ignitionsystem is inactive as long as engine torque isabove 400 foot-pounds. If torque decreases to

400 foot-pounds, the torque switch will closeand apply DC power to the ignition ON light,the FCU purge valve, and to the ignition ex-citer. Ignition will be continuous until powerincreases above 400 foot-pounds.

PROPELLERGENERALThe PT6A engine drives a three- or four-blade,oil-operated propeller (Figure 7-16). A threeblade Hartzell propeller is used on BB-2through B-1192 and has a blade angle rangeof +90° to –9°. A three blade McCauley pro-peller is used on BB-1193 through 1438, and1440 through 1443, and has a blade angle




DC POWER

L IGNITER POWER

ONOFFSTARTER ONLYIGNITION

ANDENGINE STARTER

ARMOFF

AUTOIGNITION

CLOSE400 FT-LBS

TORQUE SW

IGN ON

IGN EXCITER

IGNITER PLUGS

Figure 7-15. Ignition System Schematic




3-BLADE PROPELLER

4-BLADE PROPELLER

Figure 7-16. Propellers

range of +86.8° to –10°. Airplanes BB-1439,1444 through 1508 have a four-b ladedMcCauley propeller with a blade angle range of+87.5° to –10°. Airplanes BB-1509 and subse-quent have a four-bladed Hartzell propellerwith a blade angle range of +87.9° to –11°. Thepropeller control system provides for constant-speed operation, full feathering, reversing, andBeta mode control. Feathering is induced bycounterweights and springs.

If an engine flames out in flight or if the pilotselects the condition lever to CUTOFF, thepropeller will not feather because of the wind-milling effect and governor action. Featheringin flight should be manually selected by usingthe propeller control lever.

A conventional oil-operated propeller gover-nor achieves normal propeller operation inthe constant speed range. A preset oil-operatedoverspeed governor is provided in case of fail-ure of the normal propeller governor. In ad-dition to the normal and overspeed propellergovernors, a fuel topping function, integralwith the primary governor, provides protec-tion against propeller overspeed, as well aslimiting rpm in the reverse ranges.

FEATHERINGFeathering is a function of counterweights at-tached to each blade root and spring forces inthe propeller cylinder.

UNFEATHERING ANDREVERSINGUnfeathering and reversing functions are doneby hydraulic (engine oil) pressure developedby a high-pressure oil pump, which is an in-tegral part of the propeller primary governor.

The Hartzell or McCauley propeller installedin the Super King Air operates in two modes:the propeller-governing constant-speed mode,and the ground fine/Beta-reverse propellerblade angle control mode.

BASIC PRINCIPLESConstant-speed propellers operated in threeconditions under the control of a propellergovernor. These conditions are:

• Onspeed

• Overspeed

• Underspeed

OnspeedOnspeed is defined as the condition of oper-ation in which the selected rpm and actualrpm are the same.

OverspeedOverspeed is the condition of operation inwhich the actual rpm is greater than the se-lected rpm.

UnderspeedUnderspeed is the condition of operation inwhich the actual rpm is less than the se-lected rpm.

CONTROLSpeed (rpm) control is a function of the pro-peller governor. This unit is engine-drivenand operates on the principle of balancing twoopposing forces, both of which are variables.These forces are speeder spring force and fly-weight force.

Speeder Spring ForceSpeeder spring force is a function of, and var-ied by, the position of the propeller controllever.

Flyweight ForceFlyweight force is a function of, and varied by,propeller rpm through a reduction gear section.




If the speeder spring force is greater than fly-weight force, the propeller would be operat-ing in an underspeed condition.

If the flyweight force is greater than speederspring force, the propeller would be operatingin an overspeed condition.

When the speeder spring and flyweight forcesare equal, the propeller is onspeed.

Unbalance of speeder spring and flyweightforces is used to position a pilot valve to ac-complish the following:

• Direct governor boosted high oil pres-sure to the propeller servo piston to re-duce the blade angle.

• Shut off governor-boosted high oil pres-sure to the propeller servo piston andconnect the piston chamber to the oilsump, allowing the counterweights andpropeller spring force to increase theblade angle, to include feather if de-sired. When the speeder spring and fly-weight forces are equal, the pilot valveis positioned appropriately to maintaina constant blade angle.

OVERSPEED CONTROLThe normal rpm control range of the primarygovernor is from 1,600 rpm to 2,000 rpm; thelatter is 100% rpm.

If the primary governor fails to limit rpm to2,000, a second (overspeed) governor, driven bythe reduction gearbox, operates in parallel withthe primary governor. This is called the over-speed governor. The overspeed governor has apreset speeder spring tension which limits pro-peller rpm to the preset limit of 2,120 rpm (priorto BB-1444, except 1439; 2,080 rpm), which is106% (prior to BB-1444, except 1439; 104%)of the primary governor maximum setting. If thepropeller blades stick or move too slowly fail-ing to limit rpm, a fuel topping section of theprimary governor will limit rpm to 106% of thepropeller rpm selected by the propeller control

lever (2,120 being the highest setting, propellerlevers full forward).

Test SystemThe overspeed governor incorporates a testsystem controlled by a two-position switch(Figure 7-17) for both propellers. The switchis marked PROP GOV TEST. The switch is lo-cated on the pilot’s left subpanel (BB-2 through162 had two switches).

A solenoid valve is associated with each over-speed governor. The valve is energized whenthe PROP GOV TEST switch is moved to theTEST position. When energized, the valve ap-plies governor pump pressure to change thefixed value of the overspeed governor as listedabove, to a range of from 1,830-1,910 rpm.

Operating PrinciplesWith the engine running and the propellercontrol lever full forward, moving the gov-ernor test switch to TEST will open a solenoidvalve and admit primary governor pump pres-sure to a hydraulic reset valve on the over-speed governor. Movement of the reset valvewill raise the pilot valve, simulating an over-speed, and allow governor pump pressure todrain to the reduction gearcase through thepilot valve of the overspeed governor. If thepower lever is advanced, the rpm should sta-bilize at the TEST reset value of the overspeedgovernor, which is between 1,830 and 1,910rpm (Figure 7-18).




Figure 7-17. PROP GOV TEST Switch

FUEL TOPPING (POWERTURBINE) GOVERNORIf a mechanical failure causes the propeller tolock or stick, it will not respond to oil pres-sure changes. The primary and overspeed gov-ernors, although still operating normally, willbe unable to control propeller rpm with oilpressure. The fuel topping governor (FTG), anintegral part of the primary governor, acts toreduce fuel flow, which in turn reduces pro-peller rpm. With a locked propeller (fixedpitch propeller), a power reduction will con-trol rpm as long as airspeed is not increasedexcessively.

The fuel-topping governor is designed to ventair pressure from the FCU, which results in afuel flow reduction. The propeller rpm atwhich the FTG activates is determined by pro-peller control lever position. With the pro-peller locked, the FTG will reduce fuel flowwhen the overspeed reaches approximately106% of the selected propeller rpm.

The FTG utilizes the same flyweights andpilot valve mechanism of the primary gover-nor. If the primary governor fails, the fuel-top-ping governor will not be operational. Theresultant overspeed will, however, be con-trolled by the backup overspeed governor.




PROP GOVTEST SWITCH

TRANSFERGLAND

SPEED SETTING SCREW

SPEEDERSPRING

FLYWEIGHT

RESTRICTOR

FINE PITCH

ENGINE OILSUPPLY

OIL DUMP TOREDUCTION BOX

OIL DUMP TOREDUCTION BOX

PROPELLEROVERSPEEDGOVERNOR

CONSTANTSPEEDPROPELLERGOVERNOR

RELIEFVALVE

GOVERNORPUMP

FLYWEIGHT

PROPELLER LEVER

SPEEDERSPRING TEST

SOLENOID

Figure 7-18. Propeller Governor Test Schematic

REVERSE OPERATIONWhen full reverse is selected, the power leverssend three commands:

1. Spool the compressor to 83% ± 5% N1with a fuel flow increase.

2. Decrease the propeller blade angle to –9°or –10°.

3. Reset the FTG to 95% of the rpm se-lected by the propeller lever.

The maximum allowable propeller speed in re-verse is 1,900 rpm; however, this is not anoverspeed limitation for the propeller or powerturbine. The 1,900-rpm limit, which is con-trolled by the FTG, assures that the propellerdoes not attain 2,000 rpm, which brings thepropeller on speed and begins to interfere withthe reverse operation.

BETA MODE CONTROLBeta control defines a range of operation inwhich the pilot can reduce the residual idlethrust of the propeller by reducing blade angle.This reduction in blade angle and, therefore,propeller thrust, is accomplished by liftingthe power levers aft into the ground fine rangeon BB-1439, 1444 and subsequent. For prioraircraft this is accomplished by lifting thepower levers aft to a position just above thered and white lines (reverse range) on thethrottle quadrant.

The propeller used in the King Air Series in-cludes a Beta valve, which forms an integralpart of the propeller governor. The pilot canmechanically position this valve, within a lim-ited (ground) range, described above, to effectpropeller blade angle changes. Propeller servopiston movement is fed back to the valve bya mechanical follow-up system to null theBeta valve when the blades reach the desiredangle, and blade angle will remain constantuntil the pilot selects another angle.

PROPELLER OPERATINGPRINCIPLES

OnspeedWhen the upward force of the governor fly-weights (Figure 7-19) is equal to the downwardforce of the speeder spring, the governor pilotvalve is positioned to shut off the governorpump pressure from the propeller piston andisolate the propeller cylinder from the gearcasedrain. This, in effect, hydraulically locks theblades at a specific angle. This condition doesnot prevail for very long as changes in altitude,temperature, airspeed, and inherent leakage atthe prop transfer sleeve require blade anglechanges. In effect, in any constant-speed con-dition, the governor is hunting through a verynarrow range to maintain the selected rpm.

OverspeedWhen an overspeed occurs, the governor fly-weight force (Figure 7-20) exceeds the speederspring force. This occurs when the propeller hasaccelerated above the selected rpm. The in-creased flyweight force will raise the governorpilot valve and reduce oil pressure at the propellerpiston, allowing the counterweights and springto increase blade angle and decelerate the pro-peller until an onspeed condition occurs.

UnderspeedWhen an underspeed condition occurs, the pro-peller decelerates below the selected rpm andthe speeder spring force overcomes the force ofthe flyweights (Figure 7-21). As a result, thepilot valve moves down and allows the gover-nor pump to apply oil pressure to the propellerservo piston, resulting in a decrease in bladeangle. This allows the propeller to accelerateuntil the flyweight force equals the speederspring force and pressure is again restrictedfrom the propeller servo piston.







GOVERNORPUMP

OIL

PRIMARY PROP GOVERNOR1,600 - 2,000 RPM

OVERSPEED

REVERSELEVER

HYDRAULICOVERSPEEDGOVERNOR

TOCASE

TOCASE

BETAVALVE

TRANSFERGLAND

LOW PITCH(HIGH OIL PRESSURE)

AUTOFEATHER SOLENOID (NC)

PROPLEVER

PILOTVALVE

2,120 RPMNORMAL(2,080 PRIOR TO BB-1444, EXCEPT 1439)

Figure 7-20. Propeller Overspeed Schematic

TRANSFERGLAND



TOCASETO

CASE


REVERSELEVER

BETAVALVE

OIL


PILOTVALVE


GOVERNORPUMP

PROPLEVER

Figure 7-19. Propeller Onspeed Schematic

POWERPLANT POWERCONTROLThe powerplant (engine-propeller combina-tion) is controlled by the interaction of threelevers (Figure 7-22): a condition lever, a powerlever, and a propeller control lever.

Condition LeverThe condition lever (Figure 7-22) is me-chanically connected to the FCU to operatea fuel cutoff valve that shuts off meteredfuel to the fuel manifold.

The condition levers located on the powerlever quadrant (last two levers on the rightside) are in the center pedestal and have threedesignated positions: FUEL CUTOFF, LOWIDLE, and HIGH IDLE. The FUEL CUT-OFF position will shut off fuel to the en-g ine , and the LOW IDLE pos i t ion wi l lestablish a fuel flow that will sustain 61%(56% in B200s prior to BB-1444, except

1439; or 52% in the PT6A-41) gas genera-tor, or N1 rpm. HIGH IDLE will establish afuel flow that will sustain 70% N1 rpm. Thereis a progressive increase in fuel flow as thecondition lever is moved from LOW IDLE toHIGH IDLE, and any rpm may be selected be-tween LOW IDLE and HIGH IDLE.

Power LeversPower levers (Figure 7-22) are located on thepower lever quadrant (first two levers on the leftside) on the center pedestal and they are me-chanically interconnected through a cam box tothe FCU, the Beta valve and follow-up mecha-nism, and the fuel topping (NP) governor. Thepower lever quadrant permits movement of thepower lever in the forward thrust (Alpha) rangefrom idle to maximum thrust and in the groundfine (Beta prior to BB-1444, except 1439) or re-verse range from idle to maximum reverse. Adetent in the power lever quadrant at the IDLEposition prevents inadvertent movement of thelever into the ground fine (Beta prior to BB-




GOVERNORPUMP

OIL


UNDERSPEED

REVERSELEVER


TOCASETO

CASE

BETAVALVE

TRANSFERGLAND



PROPLEVER

PILOTVALVE


Figure 7-21. Propeller Underspeed Schematic





PRIOR TO BB-1444, EXCEPT 1439

POWERLEVERS

POWERLEVERS

CONDITIONLEVERS

CONDITIONLEVERS

PROPELLERLEVERS

PROPELLERLEVERS

Figure 7-22. Powerplant Control Levers

1444, except 1439) or reverse range. The pilotmust lift the power levers up and over this de-tent to select ground fine/Beta or reverse.

The function of the power levers in the forwardthrust (Alpha) range is to establish a gas gen-erator rpm through the gas generator gover-nor (N1) and a fuel flow that will produce andmaintain the selected N1 rpm. In the groundfine (Beta) range, the power levers are usedto reduce the propeller blade angle, thus re-ducing residual prop thrust. In the reverserange, the power lever functions to:

1. Select a blade angle proportionate to theaft travel of the lever.

2. Select a fuel flow that will sustain theselected reverse power.

3. Reset the fuel topping governor (NP)from its normal 106% to a range of ap-proximately 95%.

Ground Fine (Beta) andReverse ControlThe geometry of the power lever linkage (Figure7-23) through the cam box is such that powerlever increments from idle to full forward thrusthave no effect on the position of the Beta valve.When the power lever is moved from idle intothe reverse range, which requires the powerlevers to be lifted over a second gate in BB-1439, 1444 and subsequent, it positions theBeta valve to direct governor pressure to thepropeller piston, decreasing blade angle throughzero into a negative range (Figure 7-23). Thetravel of the propeller servo piston is fed backto the Beta valve to null its position and, in ef-fect, provide many negative blade angles all theway to full reverse. The opposite will occurwhen the power lever is moved from full reverseto any forward position up to idle, therefore pro-viding the pilot with manual blade angle con-trol for ground handling.




GOVERNORPUMP

OIL


POWER/REVERSELEVER


TOCASEDRAINTO

CASE DRAIN

BETAVALVE

TRANSFERGLAND


NC

PROPLEVER

POWERLEVER

PILOTVALVE

2,080 RPMNORMAL

REV IDLE LO HI

FX LO HI

APPROXIMATELY1870 RPM IN TEST

MODE

Figure 7-23. Beta and Reverse Control

Propeller Control LeverThe propeller control lever (Figure 7-24) op-erates in the throttle quadrant (the two cen-ter levers) on the center pedestal and it ismechanically connected to the primary pro-peller governor. In the forward thrust, orconstant-speed range, the propeller controllever selects rpm from low rpm to high rpm(1,600 to 2,000 rpm) by changing the set-ting of the primary propeller governor. Thepropeller control lever is also used to featherthe propeller by moving the lever aft into thefeather detent position. This action posi-tions the primary propeller governor’s pilotvalve to dump oil from the propeller servopiston chamber and allows the propellercounterweights and springs to move the pro-peller blades to the full feather position. Adetent (requiring more force to overcome)at the low rpm position, prevents inadver-tent movement of the propeller lever into thefeather range.

Friction ControlFour friction locks (Figure 7-25) are locatedon the center pedestal. Turning the knobscounterclockwise will reduce friction on thepowerplant control levers. Clockwise rota-tion will increase friction or lock the levers inany desired position.

ENGINE INSTRUMENTATION

Engine Temperature (ITT)GagesEngine operating temperature at station T5 issensed by eight thermocouple probes locatedbetween the gas generator turbine and the firststage power turbine. The probes are connectedin parallel to provide the best average reading.

Interstage turbine temperature (ITT) mea-surement is calibrated to provide a very ac-curate reading. This is done by a temperaturetrimmer located on top of the engine. Thistemperature trimmer is connected in parallelwith the ITT harness, and it is factory preset.

The temperature sensed by the thermocouplesis sent to gages (Figure 7-26) on the center in-strument panel calibrated in degrees Celsiusand designated ITT. On BB-1484, 1486 andsubsequent, the gages use DC power. Prior toBB-1486, excluding BB-1484, the gages areself-energizing and do not require DC power.

Engine Power (Torque)Engine power is a measurement of that por-tion of the power developed by the engine thatis transmitted to the propeller. This power ismeasured in foot-pounds and is designated asengine TORQUE.

The ring gear of the first-stage planetary re-duction gearbox is fixed in rotary direction,




Figure 7-24. Propeller Control Lever

Figure 7-25. Friction Control Knobs

but it can move a limited amount in axial di-rection because of helical splines. Therefore,the first-stage ring gear is a reaction memberthat reacts to an increase or decrease of appliedtorque by moving aft as engine torque is in-creased and moving forward as engine torqueis decreased. This axial motion of the ringgear is balanced by oil pressure in a meteredchamber called a torquemeter chamber.

The pressure in the torquemeter chamber issensed by a transmitter and sent to a gage

(Figure 7-27) on the engine instrument panelthat is calibrated in foot-pounds of torquetimes 100. The torquemeter chamber receivesa supply of oil at a relatively constant pressurefrom the engine lubricating system. On BB-1484, 1486 and subsequent, the gages use DCpower. Prior to BB-1486, except 1484, it ispowered by the 26-volt AC bus.

Torque LimiterEngine torque is automatically limited to apreset value by a torque limiter that is suppliedwi th a t o rque p re s su re s igna l f rom thetorquemeter.

At a predetermined torque pressure of 2,368to 2,447 foot-pounds, the torque limiter willbleed off and change the pneumatic servo pres-sures in the fuel control unit. This action re-duces metered fuel flow and, consequently, gasgenerator power to the preset limit of thetorque limiter. The system is designed only toprotect the nose gearbox and reduction gear-ing from excessive torque. It will not preventa pilot from exceeding the certified maximumtorque of 2,230 foot-pounds.




B200 — BB-1484, 1486 AND AFTER

B200 — PRIOR TO BB-1486, EXCEPT 1484

200

Figure 7-26. ITT Gages



Figure 7-27. Torque Gages

Propeller RPM Propeller rpm output is sent to a gage (Figure7-28) on the engine instrument panel cali-brated directly in propeller revolutions perminute. On BB-1484, 1486 and subsequent,DC power is required. Prior to BB-1486, ex-cept BB-1484, these gages do not requireaircraft DC electrical power, as they are op-erated by tachgenerators.

Engine RPM (N1)Engine or gas generator (N1) rpm is also sentto a gage (Figure 7-29) on the engine in-strument panel. The gage is calibrated inpercentage of design 100% rpm. On BB-1484, 1486 and subsequent, DC power is re-quired. It does not require aircraft electricalpower prior to BB-1484 including BB-1485as they are powered by tachgenerators.

SYNCHROSCOPEA synchroscope (Figure 7-30) with black andwhite cross patterns is located on the lower rightcorner of the pilot’s instrument panel to aid inmanual propeller synchronization. The disc willrotate in the direction of the higher rpm engine.The disc will stop rotating when the enginesare synchronized. Input signals to the synchro-scope are from the propeller tachgenerators.






Figure 7-28. Propeller RPM Gages



Figure 7-29. Engine RPM Gages

Figure 7-30. Propeller Synchroscope andSwitches (Type II)




SYNCHROPHASING

GeneralTwo synchrophasing systems are available. Theyare identified as Type I and Type II systems.

Type II System (BB-935 andSubsequent)The Type II synchrophaser system is anelectronic system, certified for takeoff andlanding (Figure 7-31). It functions to matchthe rpm of both propellers and establish ablade phase relationship between the rightand left propellers to reduce cabin noise toa minimum.

The system can not reduce rpm of either pro-peller below the datum selected by the pro-peller control lever. Therefore, there is noindicating light associated with the Type IIsystem.

ControlThe system is controlled by a two-positionswitch (Figure 7-30) located on the lower rightside of the pilot’s instrument panel.

Operation Type II SystemTurning the control switch on will supply DCpower to the electronic control box. Input sig-nals representing propeller rpm are receivedfrom magnetic pickups on each propeller. Thecomputed input signals are corrected to a com-mand signal and sent to an rpm trimming coillocated on the propeller governor of the slowengine and its (propeller) rpm is adjusted tothat of the other propeller.

NOTEIf the synchrophaser is on and failsto synchronize the propellers, turn itoff, then manually synchronize thepropellers and turn it back on.

DC BUS

SYNCHROSCOPE

PROP SYNCOFF

SYNC CONTROLLER

RPM SENSORRPM SENSOR

PROPELLER GOVERNOR (PRIMARY)

PROPELLER SPINNER

Figure 7-31. Type II System Schematic

Type I System BB-2 throughBB-934

The Type I system uses the master-slave con-cept (Figure 7-32). The left propeller is themaster propeller and the right propeller is theslave. The system functions to adjust the rpmof the right propeller to that of the left, withina limited rpm range and at the same time it pro-vides a specific blade phase relationship be-tween the left and right propellers. The overalleffect of the synchrophaser system is to reducenoise level in the cabin to a low value.

ControlSystem control is achieved by a two-positionswitch (Figure 7-33) on the lower right side

of the pilot’s instrument panel. Being a mas-ter slave system, it should be off during groundoperation, takeoff, and landing, because if themaster engine fails, the rpm of the slave en-gine will decrease a limited amount. The pro-pellers should be manually synchronizedbefore turning the system on.

An amber light (Figure 7-34) on the caution/advisory panel will come on if the synchro-nizer system is on and the landing gear is se-lected down.

Operation Type I SystemWhen the synchrophaser switch is on (Figure7-33), DC power is available to the control box.Input signals are received by the control box




DC BUS

SYNCHROSCOPE

SYNC CONTROLLER

RPM MONOPOLE RPMMONOPOLE

SLAVE

MSYNCACTUATOR

LDG GRUP

DOWNSYNC ON

MASTER

PROPELLER OVERSPEED GOVERNOR

Figure 7-32. Type I System Schematic




from monopoles on each propeller overspeedgovernor. These signals represent propellerrpm. The pulse rate difference of the signalsis corrected to a command signal, which istransmitted to an actuator on the right engineprimary governor housing. The actuator, inturn, trims the right propeller governor tomatch its rpm to the left (master) propeller.This adjustment does not affect the positionof the propeller control lever. When turned off,the stepping motor or actuator will run to a neu-tral position.

PROPELLERFEATHERINGThe Hartzell and McCauley propeller instal-lations on the King Air are full-featheringpropellers.

The propeller servo piston is spring-loadedto FEATHER. The counterweights attachedto each blade near the root are supplementedby feathering springs. The centrifugal forcesexerted by the counterweights and spring

forces tend to induce high blade angles ortoward feather.

Feathering is normally accomplished with thepropeller control lever (Figure 7-22). Movingthis lever aft to the FEATHER position will me-chanically raise the governor pilot valve anddump oil from the propeller cylinder. The coun-terweights and springs will then rapidly featherthe propeller.

Also, if the engine is shut down on the groundusing the condition lever, the oil pressure de-creases and the centrifugal force of the coun-terweights plus the springs will eventuallyfeather the propeller. However, this is not a rec-ommended procedure. The prop should befeathered with the prop control lever.

AUTOFEATHERINGAn autofeather system is available in theevent of engine failure. This system willrapidly feather the affected propeller byopening a solenoid valve on the overspeedgovernor and will dump propeller controloil. The counterweights and springs willrapidly feather the propeller.

ControlAutofeather is controlled by a single switch(Figure 7-35) for both propellers. The switchis marked ARM, OFF, and TEST.

ArmingTurning the switch to the ARM position ap-plies power to a microswitch in each powerlever quadrant. The switches will close whenthe power levers are advanced to a position thatshould produce approximately 90% N1 rpm.

Figure 7-34. Sync Light

Figure 7-33. Propeller Synchroscopeand Switch (Type I)

Figure 7-35. AUTOFEATHER Switch

When this occurs, electrical power is finallytransmitted to torque switches.

Once engine torque is over 400 foot-pounds, theopposite engine’s autofeather annunciator willilluminate.

IndicationTwo green lights (Figure 7-36) on the cau-tion/advisory panel marked L and R AUTOFEATHER will illuminate if the autofeathersystem is armed, the power levers are ad-vanced to approximately 90% N1 rpm orgreater, and the engines are developing powerin excess of 400 foot-pounds of torque.

TestingThe TEST position of the autofeather systemi s u s e d t o b y p a s s t h e p o w e r l e v e r m i -croswitches and induce arming at a muchlower power setting to test the integrity ofthe torque switches, the arming relays, thedump solenoid valve, and the arming lightswithout high power settings. The autofeathersystem is designed for use only during crit-ical power periods such as takeoff, approach,and landing, and it should be turned off underall other operating conditions.

OPERATING PRINCIPLESAssume that the autofeather system is armedfor takeoff. As the power levers are advanced,the microswitches will close at a position in thequadrant representing 90% N1 rpm. Electricalpower will now be applied to engine torque-sensitive switches (two for each engine). Oneswitch on each engine is set to open at ap-proximately 200 foot-pounds of torque andthe second switch on each engine opens at 400foot-pounds of torque. When passing through90% N1, a green AUTOFEATHER light foreach engine should be on, indicating a fullyarmed condition for both engines.

AUTOFEATHERINGIf an engine fails (Figure 7-37) (for example,during takeoff), a torque switch will closewhen torque decays to 400 foot-pounds and theAUTOFEATHER light of the operating en-gine will extinguish, indicating that its auto-feather circuit is disarmed. Then as torque onthe failing engine decays to 200 foot-pounds,a second torque switch closes. The armingrelay will be energized, and the dump valvelocated on the overspeed governor will opento dump propeller servo oil and produce rapidfeathering. In addition, the autofeather lightfor the failed engine will extinguish.




Figure 7-36. Autofeather Lights

Autofeather TestThe TEST position (Figure 7-38) of the AUTO-FEATHER switch bypasses the power lever90% N1 switches. With both engines set to ap-proximately 500 foot-pounds of torque, mov-ing the switch to the TEST position and reducingpower slowly on one engine, the opposite en-gine’s autofeather light should extinguish atapproximately 400 foot-pounds of torque.Continued power reduction should cause theother autofeather light to extinguish at 200 foot-pounds, then begin flashing as the feather/un-feather cycle begins. The propeller will notcompletely feather during the testing proce-dure, since the engine is still producing torque.

NOTEIf the condition levers are not set atLOW IDLE, it may not be possibleto reduce torque below 200 foot-pounds, which would result in thepropeller not cycling during test.

When the autofeather system is activated, adump valve on the overspeed governor is en-ergized open, connecting the propeller servopiston chamber directly to the drain line, dump-ing propeller oil into the reduction gearcase.The counterweights and springs will move theblades to the full-feathered position.




L AUTOFEATHERL AUTOFEATHER

R AUTOFEATHERR AUTOFEATHER

NC

DUMPVALVE

NC

DUMPVALVE

TORQUESWITCH

TORQUESWITCH

TORQUESWITCH

TORQUESWITCH

200

400

400

200

TEST

C/B

AUTOFEATHER

ARM

LEFTPOWERLEVER

SWITCH

RIGHTPOWERLEVER

SWITCH* CLOSED AT HIGH N1

OFF

ARMINGRELAY

AUTOFEATHERLIGHTS

ARMINGRELAY

Figure 7-37. Autofeather System Schematic (Both Power Levers at Approximately 90% N1; Right Engine has Failed)

UNFEATHERINGWith the prop levers set full forward, pro-peller unfeathering occurs automatically withoil pressure as the engine is started and theblade angle will decrease to the datum set bythe Beta/reverse mechanism (approximately18°). As there are no unfeathering pumps in-stalled in the King Air 200, the engine mustbe operating to unfeather the propeller.

LIMITATIONS(POWERPLANT)

GENERALThe limitations contained in Section II of thePilot’s Operating Handbook and FAA-ap-proved Flight Manual must be observed inthe operation of the Super King Air.




L AUTOFEATHERL AUTOFEATHER

R AUTOFEATHERR AUTOFEATHER

NC

DUMPVALVE

NC

DUMPVALVE

TORQUESWITCH

TORQUESWITCH

TORQUESWITCH

TORQUESWITCH

400

200

TEST

C/B

AUTOFEATHER

ARM

LEFTPOWERLEVER

SWITCH

RIGHTPOWERLEVER

SWITCH* CLOSED AT HIGH N1

OFF

ARMINGRELAY

AUTOFEATHERLIGHTS

400

200

Figure 7-38. Autofeather Test Schematic (Left Power Lever Below 200 ft-lb;Right Power Lever Above 400 ft-lb)




POWERPLANTManufacturer: Pratt & Whitney Aircraft ofCanada LTD, Engine Model No. PT6A-41/42.

ENGINE OPERATING LIMITSThe following limitations in Tables 7-1, 7-2and 7-3 shall be observed. Each column pre-

sents limitations. The limits presented do notnecessarily occur simultaneously. Refer toPratt & Whitney Maintenance Manual for spe-cific actions required if limits are exceeded.

APPROVED FUELSSee Chapter 5, FUEL SYSTEM.

OPERATING TORQUE MAXIMUM GAS GENERATOR PROP OIL OILCONDITION SHP FT-LB OBSERVED RPM N1 RPM PRESS. TEMP

(1) ITT °C RPM % NP PSI (2) °C (3) (4)

STARTING --- --- 1,000 (5) --- --- --- --- -40 (min)

LOW IDLE --- --- 750 (6) 22,875 61 (min) (13) 60 (min) -40 to 99

HIGH IDLE --- --- --- --- (7) --- --- -40 to 99

TAKEOFF ANDMAX CONT 850 2,230 800 38,100 101.5 2,000 100 to 135 0 to 99

MAX CRUISE 850 2,230 (8) 800 38,100 101.5 2,000 100 to 135 0 to 99

CRUISE CLIMB ANDREC (NORMAL) CRUISE 850 2,230 (8) 770 38,100 101.5 2,000 100 to 135 0 to 99

MAX REVERSE (9) 850 --- 750 --- 88 1,900 100 to 135 0 to 99

TRANSIENT --- 2,750 (5) 850 38,500 (10) 102.6 (10) 2,200 (5) --- 0 to 104 (11)

Table 7-1. ENGINE OPERATING LIMITS (PT6A-42 ENGINE BB-1439, 1444 AND SUBSEQUENT)


(1) ITT °C RPM % NP PSI (2) °C (3) (4)

STARTING --- --- 1,000 (5) --- --- --- --- -40 (min)

LOW IDLE --- --- 750 (6) 21,000 56 (min) --- 60 (min) -40 to 99

HIGH IDLE --- --- --- --- (7) --- --- -40 to 99

TAKEOFF ANDMAX CONT 850 2,230 800 38,100 101.5 2,000 100 to 135 0 to 99

MAX CRUISE 850 2,230 (8) 800 38,100 101.5 2,000 100 to 135 0 to 99

CRUISE CLIMB ANDREC (NORMAL) CRUISE 850 2,230 (8) 770 38,100 101.5 2,000 100 to 135 0 to 99

MAX REVERSE (9) 850 --- 750 --- 88 1,900 100 to 135 0 to 99

TRANSIENT --- 2,750 (5) 850 38,500 (10) 102.6 (10) 2,200 (5) --- 0 to 104 (11)

Table 7-2. ENGINE OPERATING LIMITS (PT6A-42 ENGINE PRIOR TO BB-1439,1444 AND SUBSEQUENT)





(1) ITT °C RPM % NP PSI (2) °C (3) (4)

STARTING --- --- 1,000 (5) --- --- --- --- -40 (min)

LOW IDLE --- --- 660 (6) 19.500 52 (min) --- 60 (min) -40 to 99

HIGH IDLE --- --- --- --- (7) --- --- -40 to 99

TAKEOFF (12) 850 2,230 750 38,100 101.5 2,000 105 to 135 10 to 99

MAX CONT ANDMAX CRUISE 850 2,230 (8) 750 38,100 101.5 2,000 105 to 135 10 to 99

CRUISE CLIMB ANDREC CRUISE 850 2,230 (8) 725 38,100 101.5 2,000 105 to 135 0 to 99

MAX REVERSE (9) --- --- 750 --- 88 1,900 105 to 135 0 to 99

TRANSIENT --- 2,750 (5) 850 38,500 (10) 102.6 (10) 2,200 (5) --- 0 to 104 (11)

Table 7-3. ENGINE OPERATING LIMITS (PT6A-41 ENGINE)

FOOTNOTES:

1. Torque limit applies within range of 1,600-2,000 propeller rpm (N2). Below 1,600 propeller rpm torque is limited to 1,100 ft-lbs.

2. When gas generator speeds are above 27,000 rpm (72% N1) and oil temperatures are between 60°C and 71°C, normal oil pres-sures are: 105 to 135 psi below 21,000 feet; 85 to 135 psi at 21,000 feet and above.

During extremely cold starts, oil pressure may reach 200 psi. Oil pressure between 60 and 85 psi is undesirable; it should be tolerated only forthe completion of the flight, and then only at a reduced power setting not exceeding 1,100 ft-lbs torque. Oil pressure below 60 psi is unsafe; itrequires that either the engine be shut down, or that a landing be made at the nearest suitable airport, using the minimum power required tosustain flight. Fluctuations of ± 10 psi are acceptable.

3. A minimum oil temperature of 55°C is recommended for fuel heater operation at takeoff power.

4. Oil temperature limits are -40°C and 99°C. However, temperatures of up to 104°C are permitted for a maximum time of 10 minutes.

5. These values are time limited to five seconds.

6. High ITT at ground idle may be corrected by reducing accessory load or increasing N1 rpm.

7. At approximately 70% N1.

8. Cruise torque values vary with altitude and temperature.

9. This operation is time limited to one minute.

10. These values are time limited to 10 seconds.

11. Values above 99°C are time limited to 10 minutes.

12. These values are time limited to five minutes.

13. 1,100 rpm for McCauley Propeller, 1,180 rpm for Hartzell Propeller.

PROPELLER

Manufacturer:• Prior to BB-1193 and BL-37

Hartzell Propeller, Inc. Diameter 98.5 inches

• BB-1193 through 1438,BB-1440 through 1443, BL-37 through 138 McCauley Propeller Diameter 98.0 inches

• BB-1439, 1444 through 1508, McCauley Propeller Diameter 94.0 inches

• BB-1509 and subsequentHartzell Propeller Diameter 93.0 inchesRotational Speed Limits

Rotational Speed Limits:• 2,200 rpm (Transient)—Not exceeding

five seconds

1,900 rpm—Reverse

2,000 rpm—All other conditions

Propeller Rotational OverspeedLimitsThe maximum propeller overspeed limit is2,200 rpm and is time-limited to five seconds.

Sustained propeller overspeeds faster than2,000 rpm indicate failure of the primarygovernor. Flight may be continued at pro-peller overspeeds up to 2,120 rpm (2,080rpm prior to BB-1444, except 1439) pro-vided torque is limited to 1,800 foot-pounds.Sustained propeller overspeeds faster than2,120 rpm (or 2,080 as indicated above) in-dicate failure of both the primary governorand the secondary governor, and such over-speeds are unapproved.

POWERPLANT INSTRUMENTMARKINGSThe powerplant instrument markings are givenin Table 7-4.

STARTER LIMITSUse of the starter is limited to:

40 seconds ...................................... ON60 seconds .................................... OFF

Then, if necessary:

40 seconds ...................................... ON60 seconds .................................... OFF

Then, if necessary:

40 seconds ...................................... ON30 minutes .................................... OFF







* Starting Limit (Dashed Red Line) 1,000°C** A dual band, yellow-green arc extends form 85 to 100 psi, indicating the extended range of normal oil pressures for operation

at 21,000 feet or above*** 1,180 to 2,000 rpm (Hartzell Propeller), 1,100 to 2,000 rpm (McCauley Propeller)

Table 7-4. POWERPLANT INSTRUMENT MARKINGS

RED LINE YELLOW ARC GREEN ARC RED LINEINSTRUMENT MINIMUM CAUTION NORMAL MAXIMUM

LIMIT RANGE OPERATING LIMIT

INTERSTAGE TURBINETEMPERATURE (ITT) * --- --- 400°C to 800°C 800°C

TORQUEMETER --- --- 0 to 2,230 ft-lb 2,230 ft-lb

PROPELLERTACHOMETER (N2) --- --- *** 2,000 rpm

GAS GENERATORTACHOMETER (N1) --- --- 61 to 101.5% 101.5%

OIL TEMPERATURE --- --- 0°C to 99°C 99°C

OIL PRESSURE ** 60 psi 60 to 100 psi 85 psi to 135 psi 135 psi

BB-2 THROUGH 1485, EXCEPT 1484




TORQUEMETER --- --- 400 ft-lb to 2,230 ft-lb 2,230 ft-lb

PROPELLERTACHOMETER (N2) --- --- 1,600 rpm to 2,000 rpm 2,000 rpm

GAS GENERATORTACHOMETER (N1) --- --- --- 101.5%

OIL TEMPERATURE --- --- 10°C to 99°C 99°C

OIL PRESSURE ** 60 psi --- 100 psi to 135 psi 200 psi

PT6A-41 ENGINE




TORQUEMETER --- --- 400 to 2,230 ft-lb 2,230 ft-lb


8-i

CHAPTER 8FIRE PROTECTION

CONTENTS

Page

INTRODUCTION ................................................................................................................... 8-1

FIRE DETECTION ................................................................................................................. 8-1

General ............................................................................................................................. 8-1

Indicators.......................................................................................................................... 8-4

FIRE EXTINGUISHING ........................................................................................................ 8-4

General ............................................................................................................................. 8-4

Controls, Indicators, and Operation ................................................................................. 8-4

Limitations ....................................................................................................................... 8-4

TESTING OF THE SYSTEMS............................................................................................... 8-6

PORTABLE FIRE EXTINGUISHERS ................................................................................... 8-7

QUESTIONS ........................................................................................................................... 8-8




8-iii

ILLUSTRATIONS

Figure Title Page

8-1 Fire Detection System—BB-1439, 1444 and After ................................................. 8-2

8-2 Fire Detection System—BB-2 through 1443 Except 1439 ...................................... 8-3

8-3 Fire-Extinguishing System....................................................................................... 8-5

8-4 Gage Location .......................................................................................................... 8-6

8-5 Portable Fire Extinguisher........................................................................................ 8-7

TABLES

Table Title Page

8-1 Temperature vs. Pressure Data ................................................................................. 8-6




INTRODUCTIONThe two engines each have independently operating fire-detection systems. A temper-ature-sensing cable or three flame detectors per engine (operating through an amplifier)turn on the appropriate warning light. Separate fire-extinguishing systems are availableas an option. Crew activation is required to release the extinguishing chemical agent intothe nacelle with the fire.

FIRE DETECTION

GENERALOn BB-1439, BB-1444 and after, the systemconsists of a temperature-sensing cable foreach engine; two red warning annunciators, LENG FIRE and R ENG FIRE; a test switch onthe copilot’s left subpanel and a circuit breakerlabeled FIRE DET on the right side panel (No.1 dual fed bus) (Figure 8-1).

Prior to BB-1444 except BB-1439, three pho-toconductive detectors per engine each feedone control amplifier to activate the appro-priate annunciator. The left amplifier controlsa red warning light labeled FIRE L ENG; theright amplifier controls a red warning light la-beled FIRE R ENG (Figure 8-2). The detec-

FIRE PULL

FIREWARN

CHAPTER 8FIRE PROTECTION







A

B

L LEXT DET

R R

OFFTEST SWITCHENG FIRE SYS

DETAIL B(WITH FIRE

EXTINGUISHER)

LDET

R


DETAIL B(WITHOUT FIREEXTINGUISHER)

DETAIL CLEFT OR RIGHT

ENGINE FIRE

B

D

C

A

28 VDC

TEST SWITCH

SENSOR RESPONDERSIMPLIFIED CIRCUIT RESPONDER ALARM

SWITCH (N.O.)

SENSOR ELEMENT

ISOLATOR

INTEGRITY SWITCHN.C. — HELD CLOSED BYNORMAL SENSOR PRESSURE

SENSOR ELEMENTSENSOR

RESPONDER

FIRE SENSORELEMENT

C

PRINTEDCIRCUITCARDS

FIRE SENSORELEMENT

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

R ENG FIREPUSH TO EXT

OKD

L ENG FIREPUSH TO EXT

OKD

DETAIL A

Figure 8-1. Fire Detection System—BB-1439, 1444 and After




L 3EXT

DET

R 2


DETAIL B(WITH FIRE

EXTINGUISHER)

B

A

FIRE DETECTORS

CONTROLAMPLIFIERS

FIRE DETECTORS

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION


OKD


OKD

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION


OKD


OKD

APPROACHPLATE

BRT

PRESS

TO

TEST

AIRPLANES WITH 12-STATION WARNING ANNUNCIATOR PANEL

AIRPLANES WITH 20-STATION WARNING ANNUNCIATOR PANEL

1

DETAIL A

Figure 8-2. Fire Detection System—BB-2 through 1443, Except 1439

tor system operates at a high preset thresholdlevel, but occasionally the system may be setoff by sunlight if it enters nacelle openings atthe appropriate angle to reach the detectors.Power is supplied from the No. 1 dual-fed busthrough a circuit breaker on the right sidepanel.

INDICATORSWhen the temperature-sensing cable is acti-vated or the light threshold is reached (priorto BB-1444, except 1439) indicating a possi-ble fire, the appropriate light on the warningannunciator panel comes on. Assuming theintegrity of the wiring or sensor cable has notbeen compromised and the fire goes out, thelight will extinguish. Both systems can againdetect the outbreak of fire.

With the fire-extinguishing system installed,fire warning is indicated by the L or R ENGFIRE PUSH TO EXT switchlights located onthe glareshield at each end of the warning an-nunciator. Fire warning is also simultaneouslyindicated by the red warning annunciators.

FIRE EXTINGUISHING

GENERALFire in either engine compartment is smoth-ered by engulfing the nacelle compartmentwith bromotrifluoromethane (CBrF3) pres-surized with dry nitrogen. There are threespray bars per engine compartment (Figure8-3), each one supplied by one common fireextinguisher supply cylinder per engine. Onesquib per bottle incorporates a pyrotechniccartridge which releases the entire contents.The squib is fired by depressing the switch-light on the glareshield. Each engine has itsown independent system, but both circuitbreakers (fuses prior to 1098, except 1096) arefed from the hot battery bus.

CONTROLS, INDICATORS, ANDOPERATIONA three-lens control indicator is located on theglareshield when the optional extinguishersystem is incorporated (Figure 8-3). The threelenses are:

• Red—L (or R) ENG FIRE PUSH TOEXT

• Amber—D

• Green—OK

The red L (or R) ENG FIRE PUSH TO EXTlens indicates a detected fire. The three-lenscontrol indicator is pushed to activate the ap-propriate extinguisher.

The amber D lens indicates that the extin-guisher has been discharged, and the supplycylinder is empty.

The green OK lens confirms circuit continu-ity during the test function.

When a red warning light indicates a fire andit is confirmed by the pilot, the appropriate (Lor R) engine should be shut down and the fire-extinguishing switchlight depressed. This firesthe appropriate squib, releasing the contentsthrough the tubing. When the bottle is dis-charged, the amber D light illuminates.

The pressure gages, one located on each fire-extinguishing supply cylinder, reflect the con-tents of the bottle. They can be read only whileon the ground because they are located in thewheel wells. See Figure 8-4 and Table 8-1 fortemperatures vs. pressure data and for thegage location.

LIMITATIONSThe detection system is operable when elec-trical power is applied to the aircraft. But theextinguishing system can be discharged at anytime since it is operated from the hot batterybus. Therefore, even though the airplane maybe parked with the engines off, the fire-ex-tinguishing system may be discharged.







PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION


OKD


OKD

DETAIL A

L LEXT DET

R R


DETAIL B(WITH FIRE

EXTINGUISHER)

DETAIL C

EXPLOSIVESQUIB

FIRE EXTINGUISHERSUPPLY CYLINDER

PRESSUREGAGE

C

R MONITORMODULE

L MONITORMODULE

C

A

B

Figure 8-3. Fire-Extinguishing System

Each engine has its own self-contained extin-guishing system, which can be used only oncebetween recharging. This system cannot be usedto extinguish a fire in the opposite engine.

TESTING OF THESYSTEMSThe rotary test switch allows ground or in-flight testing of the detection system (Figures8-1 and 8-2).

For BB-1439, 1444 and subsequent, when theswitch is placed in the DET L or DET R po-sition, the illumination of the correspondingENG FIRE light assures the integrity of thecable and continuity of the electrical wiring.

Prior to BB-1444, except 1439, the four-posi-tion rotary test switch allows each of the de-tection sections to be individually tested. Fortesting the detection systems, an output voltageis supplied to the control amplifier, simulatinga signal from the detectors. Each of the three de-tector circuits is tested individually, causingthe appropriate panel lights to illuminate.




Figure 8-4. Gage Location

TEMPERATURES -40°/-40° -29°/-20° -18°/0° -6°/20° 4°/40° 16°/60° 27°/80° 38°/100° 48°/120°°C/°F

PSI MINIMUM 190 220 250 290 340 390 455 525 605

to to to to to to to to to

PSI MAXIMUM 240 275 315 365 420 480 550 635 730

Table 8-1. TEMPERATURE VS. PRESSURE DATA

NOTE: PRESSURES ARE EXTRACTED FROM THE BEST AVAILABLE INFORMATION AND SHOULD ONLY BE USED AS A GUIDE.

During testing, the pilot’s and copilot’s redMASTER WARNING light flashes, and, ifthe optional extinguisher system is installed,the red lenses placarded L ENG FIRE–PUSHTO EXT and R ENGINE FIRE–PUSH TOEXT illuminate. Failure of the fire detectionannunciators in any of the test positions indi-cates a malfunction in that system. When thelight fails to come on during testing, a no-gosituation exists. Should there be no responsein any position, check the circuit breaker.

For testing the extinguishing systems, the cir-cuitry of the squibs is checked for continuityby rotating the TEST SWITCH FIRE DETand FIRE EXT through the two (LEFT andRIGHT) EXT positions (Figure 8-3). Theamber D light and the green OK light shouldilluminate, indicating that the bottle charge de-tector circuitry and squib-firing circuitry areoperational and that the squib is in place(Figure 8-3).

PORTABLE FIREEXTINGUISHERSThere are two portable fire extinguishers in-side the airplane. One is in the cabin, the otheris in the cockpit. One is normally installed onthe floor on the left side of the airplane for-ward of the airstair entrance door, just aft ofthe rearmost seat; the other is underneath thecopilot’s seat (Figure 8-5).




Figure 8-5. Portable Fire Extinguisher

9-i

CHAPTER 9PNEUMATICS

CONTENTS

Page

INTRODUCTION ................................................................................................................... 9-1

GENERAL............................................................................................................................... 9-1

System Description and Location .................................................................................... 9-1

Bleed-Air Warning System .............................................................................................. 9-3

Bleed-Air Control ............................................................................................................ 9-5

Door Seal System............................................................................................................. 9-7

Flight Hourmeter.............................................................................................................. 9-7

LIMITATIONS ........................................................................................................................ 9-7

QUESTIONS ........................................................................................................................... 9-8




9-iii

ILLUSTRATIONS

Figure Title Page

9-1 Pneumatic and Vacuum Systems Diagram............................................................... 9-2

9-2 Bleed-Air Ejector ..................................................................................................... 9-3

9-3 Suction Gage and Pressure Gage.............................................................................. 9-3

9-4 Bleed-Air Warning System Diagram....................................................................... 9-4

9-5 L & R BL AIR FAIL Warning Lights ...................................................................... 9-5

9-6 BLEED AIR VALVE Switches ................................................................................ 9-5

9-7 Pneumatic Plastic Tubing ......................................................................................... 9-5

9-8 Bleed-Air Control Diagram...................................................................................... 9-6

9-9 Cabin Door Air Seal ................................................................................................. 9-7

9-10 Hourmeter................................................................................................................. 9-7




INTRODUCTIONThe Super King Air utilizes an engine bleed-air pneumatic system to provide bleed airfor the door system (door seal line), the ice protection systems (surface deice), thebleed-air warning system, the rudder boost, the hourmeter, and the brake deice system.Also, pneumatic air that is exhausted overboard via a venturi creates a negative pres-sure that is used by the vacuum system.

GENERAL

SYSTEM DESCRIPTION ANDLOCATION

Pneumatic and VacuumSystemsHigh-pressure bleed air regulated to 18 psi,supplies pressure for the surface deice system

and the vacuum source (Figure 9-1). Vacuumfor the flight instruments, pressurization con-

VALVE

L R

COBLEED AIR

515

20

AIR

CHAPTER 9PNEUMATICS







LEGENDHIGH PRESSURE BLEED AIRREGULATED BLEED AIRVACUUM

TODEICEBOOTS

EXHAUSTOVERBOARDEJECTOR

VACUUMREGULATOR

RIGHT BLEED-AIR WARNING SYSTEM

RIGHTP AIR

LEFT BLEED-AIR WARNING SYSTEM

LEFTP AIR

AIRSTAIRDOOR SEAL

LINE

LANDING GEARRESERVOIR

(HYDRAULIC GEAR ONLY)

PNEUMATIC PRESSUREGAGE

(IN COCKPIT)

PRESSURESWITCHRIGHT

SQUATSWITCH

LEFTSQUAT

SWITCH

CLOSED ONGROUND

(NO)

CHECK VALVECHECK VALVE

PNEUMATICAIR VALVE

(NO)

PNEUMATICAIR VALVE

(NO)

LEFT BRAKEDEICEVALVE(NC)

RIGHT BRAKEDEICEVALVE(NC)

18 PSIPRESSURE

REGULATOR

DEICEDISTRIBUTOR

VALVE

PRESSURATIONCONTROLLER,OUTFLOW AND

SAFETY VALVES

GYROINSTRUMENTS

GYROSUCTION

(IN COCKPIT)

HOURS

FLIGHT

1/10

60 PSID

P SWITCH

RUDDERBOOSTSYSTEM

VALVEL SERVO

R SERVO

RUDDER BOOST RUDDER BOOST

LEFT NC

RIGHT NC

15 PSIREGULATOR

4 PSIREGULATOR

Figure 9-1. Pneumatic and Vacuum Systems Diagram




troller, and surface deice originates througha venturi (bleed air ejector) which is exhaustedoverboard (Figure 9-2). One engine can sup-ply sufficient bleed air for all associated sys-tems. In addition, the brake deice systemreceives bleed air that is tapped off down-stream of each instrument air valve (Figure 9-1). Refer to Chapter 10, ICE AND RAINPROTECTION, for more information on thebrake deice system. Refer to Chapter 15,FLIGHT CONTROLS, for information on therudder boost system.

Engine bleed air is ducted from each engine toits respective L or R flow control unit mountedon the firewall. A pressure supply line tees offthe engine bleed-air line forward of the fire-wall and flow control unit. This supply line con-tains pneumatic pressure to operate the surfacedeicer, rudder boost, door seal, brake deice(hot brakes) system hydraulic reservoir (BB-1193 and after, including BB-1158 and 1167),

and the flight hourmeter. An ejector changespressure to a vacuum to operate gyro instru-ments, pressurization controller, and outflowand safety valves. The flow control unit reg-ulates the mixture of engine bleed air for pres-surization with ambient air. Pressurization airis routed through the wings and, finally, intothe cabin where it is used for heating, cooling,and pressurization.

A suction gage (Figure 9-3), which is cali-brated in inches of mercury and is located onthe copilot’s right subpanel, indicates gyrosuction. To the right of the suction gage is apneumatic pressure gage (Figure 9-3) whichindicates air pressure available to the deice dis-tributor valve, vacuum system, bleed air warn-ing, rudder boost, hourmeter, and door seal.The pneumatic pressure gage is calibrated inpounds per square inch (psi).

BLEED-AIR WARNING SYSTEMThe bleed-air warning system is installed toalert the pilot when a pressurization line orpneumatic line ruptures, exhausting hot enginebleed air into the airframe.

Whenever the temperature from this rupturereaches approximately 204°F (Figure 9-4),the plastic tubing melts, which results in theillumination of either the L BL AIR FAIL orthe R BL AIR FAIL warning lights (Figure 9-5). A severe bleed-air leak could result in a de-

Figure 9-3. Suction Gage and Pressure Gage

Figure 9-2. Bleed-Air Ejector




UNREGULATED PNEUMATIC BLEED AIR

REGULATED PNEUMATIC BLEED AIR

BLEED AIR WARNING LINE

LEGEND

ENVIRONMENTAL BLEED AIR

VACUUM

ENGINE P3BLEED-AIR

CONNECTOR

ENGINE P3BLEED-AIR

CONNECTOR

AMBIENTAIR

AMBIENTAIR

FIREWALLFIREWALL

PLUGS

PLUGS

PLUG

MANIFOLD(18 PSI REGULATOR)

BLEED-AIR WARNINGSWITCHES

RH N.O. PNEUMATICAIR VALVE

ENVIRONMENTALMIXING PLENUM

FLOW CONTROLVALVE

AIR-TO-AIRHEAT EXHCHANGER

BLEED-AIRBYPASS VALVE

BLEED-AIRBYPASS VALVE

AIR INLET AIR INLET

REAR SPAR

LH N.O. PNEUMATICAIR VALVE

RH ENGINEBLEED AIRINLET

EJECTOR

PRESSURE REGULATOR

MANIFOLD

AIR SOURCE—RHBLEED AIR WARNING

LH ENGINEBLEED AIR INLET

AIR SOURCE—LHBLEED AIR WARNING

Figure 9-4. Bleed-Air Warning System Diagram




crease in engine torque and an increase in ITT.Therefore, whenever the applicable BLEEDAIR VALVE switch (Figure 9-6) is placed intoINST and ENVIR OFF position, the pilotshould monitor the engine instruments for anincrease in torque and a decrease in ITT. Thisindicates that the leak has been isolated, if itwas a severe leak.

However, regardless of engine instruments,any time the bleed-air warning light illumi-nates, the respective bleed-air valve must bepos i t i oned i n t he INSTrumen t andENVIRonmental OFF position.

The plastic tubing (Figure 9-7) lies alongsidethe insulated pressurization air lines and theuninsulated pneumatic lines. Excessive heatfrom a ruptured bleed-air line causes the plas-tic tubing to fail and could damage surround-ing systems, or weaken the structure. Thepressure released in the plastic tubing closesa pressure switch located underneath the floor

below the copilot’s feet. When this switch(one of two switches) closes, the applicable BLAIR FAIL light illuminates.

NOTEThe bleed-air warning annunciatorwill not extinguish after closing thebleed-air valves. When the bleed-aircontrol switch is in the OPEN posi-tion, it requires DC power to open theflow control unit shutoff valve. Whenthe switch is in the INST & ENVIROFF position, it requires DC powerto close the pneumatic instrumentair valve. Both positions receive theirpower from the bleed-air control CB.

BLEED-AIR CONTROLBleed air entering the cabin, used for pressur-ization and environmental functions, is con-trolled by the two BLEED AIR VALVESswitches which are marked OPEN, ENVIR OFF,and INST & ENVIR OFF. When the switch isin the OPEN position, both the environmentalflow control unit and the pneumatic instrumentair valve open. When the switch is in the ENVIROFF position, the environmental flow controlunit closes and the pneumatic instrument airvalve remains open. In the INST & ENVIR OFFposition, both the environmental and pneumaticflow valves are closed (Figure 9-8).

Figure 9-5. L & R BL AIR FAIL WarningLights

Figure 9-6. BLEED AIR VALVE Switches

Figure 9-7. Pneumatic Plastic Tubing

9-6FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

NC

NC

NC

NO

NO

NO

NO

6-8 SEC

NO

NO

NO

CABINPRESSDUMP

TEST

PRESS

R BL AIR OFFR BL AIR OFF

L BL AIR OFFL BL AIR OFF

BLEED AIR VALVESOPEN

INSTR & ENVIR OFF

ENVIROFF


INSTR & ENVIR OFF

ENVIROFF

LH FLOWCONTROLSHUTOFF

LHPNEUMATICBLEED AIRSHUTOFF

RH FLOWCONTROLSHUTOFF

RHPNEUMATICBLEED AIRSHUTOFF

RAM AIRDOORSOLENOID

CABINPRESETSOLENOID

CABINPRESSURESAFETYVALVE

DOORSEALSOLENOID

TIMEDELAYPCB

PNEUMATIC ANDENVIRO OFF

PNEUMATIC ANDENVIRO OFF

ENVIROOFF

ENVIROOFF

OPEN

TEST

PRESS

DUMP

LH GEARSAFETYSWITCH

OPEN

RH AMBIENTAIR SHUTOFFVALVE

LH AMBIENT AIR SHUTOFFVALVE

RH GEAR SAFETY SWITCH

CABIN AIR TEMP

UP

UP

DN

DN5A

DUALFED

BUS NO. 1

DUALFED

BUS NO. 2

Figure 9-8. Bleed-Air Control Diagram

DOOR SEAL SYSTEMThe entrance door to the cabin utilizes airfrom the pneumatic system to inflate the doorseal (Figure 9-9) after the airplane lifts off.Bleed air is tapped off the manifold down-stream of the 18-psi pressure regulator. Afterthe tap, the regulated air passes through a 4-psi regulator and to the normally-open valvethat is controlled by the left landing gear safetyswitch.

FLIGHT HOURMETERThe FLIGHT hourmeter (Figure 9-10) pro-vides a readout of the airplane’s flight time.The meter is located on the copilot’s rightsubpanel. In order for it to operate, pneumaticbleed air must be supplied, and DC powermust be available through the flap control cir-cuit breaker. In addition, weight must be re-moved from the right landing gear strut toaffect the squat switch.

LIMITATIONSThe pneumatic system limitations are as follows:

• Pneumatic gage indicates, within a greenarc, the normal operating range of 12 to20 psi, and the maximum operating limit(red line) of 20 psi.

• Vacuum (suction) gage indicates, withina narrow green arc, the normal suctionfrom 15,000 to 30,000 feet MSL of 3.0to 4.3 in. Hg, or from 15,000 to 35,000feet MSL of 2.8 to 4.3 in. Hg. A widegreen arc indicates the normal vacuumrange from sea level to 15,000 feet MSLof 4.3 to 5.9 in. Hg.




Figure 9-9. Cabin Door Air Seal

R

10080 FLIGHT

HOURS 1/10

500

0

1000

SUPPLY PRESMADE IN U

OXYG

Figure 9-10. Hourmeter

10-i

CHAPTER 10ICE AND RAIN PROTECTION

CONTENTS

Page

INTRODUCTION................................................................................................................. 10-1

GENERAL ............................................................................................................................ 10-1

ICE PROTECTION—PNEUMATIC SOURCE ................................................................... 10-2

Wing and Horizontal Stabilizer Deice System .............................................................. 10-2

Controls, Indicators and Operation................................................................................ 10-5

Brake Deice System....................................................................................................... 10-6

Control and Indicator ..................................................................................................... 10-6

Operation ....................................................................................................................... 10-6

ICE PROTECTION—ELECTRICAL SOURCE.................................................................. 10-8

Windshield Heat............................................................................................................. 10-8

Controls.......................................................................................................................... 10-8

Operation ....................................................................................................................... 10-8

Propeller Heat ................................................................................................................ 10-8

Controls, Indicators and Operation.............................................................................. 10-10

Pitot Heat ............................................................................................................................ 10-11

Controls and Operation................................................................................................ 10-11

Stall Warning Vane Heat.............................................................................................. 10-12

Fuel Vent Heat ............................................................................................................. 10-12

MISCELLANEOUS SYSTEMS POWERPLANT............................................................. 10-13

Controls, Indicators and Operation.............................................................................. 10-16




WINDSHIELD WIPERS .................................................................................................... 10-17

Controls and Operation................................................................................................ 10-17

WING ICE LIGHTS ........................................................................................................... 10-18

Location and Control ................................................................................................... 10-18

LIMITATIONS.................................................................................................................... 10-18

QUESTIONS....................................................................................................................... 10-22




10-iii

ILLUSTRATIONS

Figure Title Page

10-1 Weather-Protected Airplane Surfaces .................................................................... 10-2

10-2 Ice and Rain Protection Controls and Indicators(BB-1439, 1444 and Subsequent) .......................................................................... 10-3

10-3 Ice and Rain Protection Controls and Indicators(Prior to BB-1444, Except BB-1439)..................................................................... 10-4

10-4 Wing and Horizontal Stabilizer Deice Boots System Control ............................... 10-5

10-5 Brake Deice System ............................................................................................... 10-7

10-6 Windshield Anti-Ice System .................................................................................. 10-9

10-7 Propeller Boots Heat-Control and Indicator ........................................................ 10-10

10-8 Pitot Probes and Heat Controls ............................................................................ 10-12

10-9 Stall Warning Vane and Heat Controls ................................................................ 10-12

10-10 Heated Fuel Vent and Control.............................................................................. 10-13

10-11 Powerplant Intake Ice Protection (BB-1439, 1444 and Subsequent) .................. 10-14

10-12 Powerplant Intake Ice Protection (Prior to BB-1444, Except BB-1439)............. 10-15

10-13 Engine Intake Inertial Vane Positions and Bypass Door ..................................... 10-17

10-14 Windshield Wiper Control ................................................................................... 10-17

10-15 Wing Ice Inspection Light and Control................................................................ 10-18




INTRODUCTIONIce, rain, and fogging can adversely affect a flight. Several systems have been includedon the Super King Air to protect those surfaces susceptible to the effects of weather.

Three sources of energy are used to prevent or to break up ice formations on the airplane’ssurfaces: engine bleed-air (pneumatics), electrical power, and engine exhaust.

GENERAL Surfaces kept ice-free by engine bleed-air(pneumatics) are:

• Wing and horizontal stabilizer leadingedge surfaces (inflatable boots)

• Brakes

Surfaces kept ice- and/or water-free by elec-trical energy are:

• Propellers

• Both pitot tubes

• The stall warning vane

• Both windshield panes

• Fuel vents

CHAPTER 10ICE AND RAIN PROTECTION




Surfaces kept ice-free by engine exhaustgases are:

• The air inlets for both engines

Figure 10-1 illustrates the location of the sur-faces so protected.

Heated pitot tubes, stall warning vane, wind-shield panes, fuel vents, and the engine inletlips prevent ice from forming and are com-ponents of the anti-ice systems.

The inflatable boots on the wings and hori-zontal stabilizer and the electrically-heatedpropeller deicers remove accumulated ice andare considered to be the deice system.

Also, to prevent ice from accumulating on theengine compressor intake screen, an inertialvane separating system is installed.

The ice and rain controls and indicators are lo-cated on the main instrument panel (Figures10-2 and 10-3).

ICE PROTECTION—PNEUMATIC SOURCE

WING AND HORIZONTALSTABILIZER DEICE SYSTEMThe leading edges of the wings and horizon-tal stabilizer are protected against an accu-mulation of ice buildup. Inflatable bootsattached to these surfaces are inflated whennecessary by pneumatic pressure to breakaway the ice accumulation and are deflated bypneumatic-derived vacuum. The vacuum isalways supplied while the boots are not in useand are held tightly against the skin.

Never take off or land with the bootsinflated. Do not operate deice bootswhen OAT is below –40°C (–40°F).

CAUTION




Figure 10-1. Weather-Protected Airplane Surfaces




PARKING BRAKEOFF

ENGINE ANTI-ICE

ON

MAIN

OFF

ACTUATORSTANDBY

LEFT RIGHT

COLLINS

0

10 20

30PROP AMPSSTALLWARN

BRAKEDEICE

ICE PROTECTIONPROPWSHLD ANTI-ICE

NORMAL

SURFACEDEICE

SINGLE PITOT

LANDING TAXI ICE NAV RECOG

OFF

OFF

LIGHTS

MANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

LEFT RIGHT

OFF

OFF

AUTO MANUAL FUEL VENT

HI

IN

PULL

OFF

1520

33N3

FO

R0

3060

90120 150 180 210

24027

030

033

0


RADIO ON

ST

EE



AIRSPEEDS (IAS)



CAUTION




PARK SLOW

FAST

OFF

MASTERPANELLIGHTS

ON

OVERHEADFLOODLIGHTS

OFFBRT


OFFBRT

AVIONICSPANELLIGHTS

OFFBRT

ENGINEINSTRUMENT

LIGHTS

OFFBRT

PILOTFLIGHTLIGHTS

OFFBRT


LIGHTS

OFFBRT

SIDEPANELLIGHTS

OFFBRT


LIGHTS

OFFBRT

COPILOTFLIGHTLIGHTS

OFFBRT

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTSAC VOLTS

FREQ

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft



PARK SLOW

FAST

ON

OFF

0

10 20

30PROP AMPS

LANDINGGEAR

STALLWARN

BRAKEDEICE


NORMAL

SURFACEDEICE

SINGLE PITOT

OFFMANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

RELAY

OFF

OFF


HI

2

OFF

ENGINE ANTI-ICE

ON

MAIN

OFF

ACTUATORSTANDBY

LEFT RIGHT

0 8020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

-60

0

+60BATT AMPS

0 8020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

-60

0

+60BATT AMPS

0 8020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTSAC VOLTS

FREQ

Figure 10-2. Ice and Rain Protection Controls and Indicators(BB-1439, 1444 and Subsequent)




PARKING BRAKEOFF

ENGINE ANTI-ICE

ON

MAIN

OFF

ACTUATORSTANDBY

LEFT RIGHT

COLLINS


OFF

LIGHTS

LEFT RIGHT

IN

PULL

0

10 20

30PROP AMPS

33N3

FO

R0

3060

90120 150 180 210

24027

030

033

0


RADIO ON

ST

EE



AIRSPEEDS (IAS)



CAUTION




PARK SLOW

FAST

OFF

MASTERPANELLIGHTS

ON

OVERHEADFLOODLIGHTS

OFFBRT


OFFBRT

AVIONICSPANELLIGHTS

OFFBRT

ENGINEINSTRUMENT

LIGHTS

OFFBRT

PILOTFLIGHTLIGHTS

OFFBRT


LIGHTS

OFFBRT

SIDEPANELLIGHTS

OFFBRT


LIGHTS

OFFBRT

COPILOTFLIGHTLIGHTS

OFFBRT

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

100

40 60 8020

0 3010 20

0

PUSH

FOR VOLTSDC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft



PARK SLOW

FAST

STALLWARN

BRAKE


NORMAL

DEICECYCLESINGLE PITOT

OFFMANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

OFF

OFF

AUTO INNER FUEL VENT

HI

OFF OUTER

DEICE

ICE VANE

AUTOFEATHER

EXTEND

RETRACT

OFF

ARM

LEFT RIGHT

0 8020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTSAC VOLTS

FREQ

Figure 10-3. Ice and Rain Protection Controls and Indicators(Prior to BB-1444, Except BB-1439)




Each wing has an inboard and an outboardboot. The horizontal section of the tail hasonly one boot from the left and right segmentsof the horizontal stabilizer. The vertical sta-bilizer is not, nor does it have to be, deiced(Figure 10-1).

CONTROLS, INDICATORS ANDOPERATIONThe three-position switch in the ice protectiongroup labeled DEICE CYCLE SINGLE–OFF–MANUAL controls the operation of the boots.

This switch is spring-loaded to the center OFFposition. When approximately one-half to oneinch of ice has accumulated, the switch should

be selected to the SINGLE cycle (up) positionand released (Figure 10-4). Pressure-regu-lated bleed air from the engines’ compressorssupply air through a distributor valve to inflatethe wing boots. After an inflation period of sixseconds, an electronic timer switches the dis-tributor to deflate the wing boots with vacuum,and a four-second inflation begins in the hor-izontal stabilizer boots. After these boots havebeen inflated and deflated, the cycle is com-plete, and all boots are again held down tightlyagainst the wings and horizontal stabilizer byvacuum. The spring-loaded switch must beselected up again for another cycle to occur.

Each engine supplies a common bleed-airmanifold. To ensure the operation of the sys-tem if one engine is inoperative, a check valve

ENGINE P3BLEED AIR

SOURCE

ENGINE P3BLEED AIRSOURCE

BLEED AIR FLOWCONTROL UNIT

BLEED AIR FLOWCONTROL UNIT

DEICEBOOT

DEICEBOOT

DEICEBOOT

DEICEBOOT

BRAKE DEICEVALVE BRAKE DEICE

VALVE

DEICE BOOT

PNEUMATICSHUTOFF

VALVE

PNEUMATICSHUTOFFVALVE

PNEUMATICCONTROLASSEMBLY

VACUUM REGULATOR

LEGEND

VACUUM

PRESSURE OR VACUUM

PRESSURE

STALLWARN

BRAKEDEICE


NORMAL

SURFACEDEICE

SINGLE PITOT

OFFMANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

OFF

OFF

FUEL VENT

HIOFF

AUTO MANUAL

OFF

Figure 10-4. Wing and Horizontal Stabilizer Deice Boots System Control

is incorporated in the bleed-air line from eachengine to prevent the loss of pressure throughthe compressor of the inoperative engine.

If the boots fail to function sequentially, theymay be operated manually by selecting thedown position of the same DEICE CYCLEswitch. Depressing and holding it in the MAN-UAL (down) position inflates all the boots si-multaneously. When the switch is released, itreturns to the (spring-loaded) OFF position,and each boot is deflated and held by vacuum.

A single circuit breaker located on the copi-lot’s side panel, receiving power from the No.1 dual-fed bus, supplies the electrical opera-tion of both boot systems.

The boots operate most effectively when ap-proximately one-half to one inch of ice hasformed. Very thin ice will crack and couldcling to the boots and/or move aft into un-protected areas.

When operated manually, the boots shouldnot be left inflated longer than necessary toeliminate the ice, as a new layer of ice maybegin to form on the expanded boots and be-come unremovable.

If one engine is inoperative, the loss of its pneu-matic pressure does not affect boot operation.

Refer to LIMITATIONS in this chapter foradditional information.

Electrical power to the boot system is requiredto inflate the boots in either single-cycle ormanual operation, but with a loss of this power,the vacuum will hold them tightly against theleading edge.

BRAKE DEICE SYSTEMThe disc brakes may freeze when they are ex-posed to water and snow because the carrierlining and the disc are always in contact.

An optional brake deice system provides en-gine P3 bleed air directed onto the brake as-semblies by a distributor manifold on each

main landing gear. If installed, this high-pres-sure and high-temperature air is routed througha solenoid control valve in each main wheelwell, through a flexible hose on the main gearstrut, and to the distribution manifold aroundthe brake assembly (Figure 10-5).

The brake deice system can be used on theground or in flight to prevent or melt away anyice accumulation.

CONTROL AND INDICATORThe BRAKE DEICE switch in the anti-icegroup on the pilot’s right subpanel (Figures10-2 and 10-3) activates the valves, allow-ing the pneumatic air to enter the brake man-ifolds. When this switch is activated, bothsolenoid valves are opened, and the greenBRAKE DEICE ON light on the caution ad-visory annunciator panel illuminates to ad-vise that both solenoids are being activatedto the open position (Figure 10-5). The lightdoes not, however, ensure that the valveshave actually opened. Conversely, if theBRAKE DEICE switch is turned off, the lightshould extinguish. However, it is possiblethat the valves are stuck in the open position.Confirmation that the valves are openingand closing can be made by observing a slightincrease or decrease in ITT when BRAKEDEICE is cycled. The circuit breaker for thebrake deice system is located on the copilot’ss ide panel in the weather group labeledBRAKE DEICE.

OPERATIONWith the landing gear extended, the brakedeice system may be operated on a continu-ous basis, provided that the limitations listedin that section are observed.

During ground operation, the simultaneoususe of the hot brake system and the wing deiceboots system may cause the red BLEED AIRFAIL lights on the warning annunciator panelto flash momentarily because of the substan-tial drop in pneumatic pressure. This is nor-mal, and the light should not remain on.




10-7FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

18 PSIPRESSURE

REGULATOR

18 PSI

PNEUMATICPRESSURE

PNEURIGHTP3AIR

PNEULEFT

P3AIR

LEFTBRAKEDEICE

MANIFOLD

RIGHTBRAKEDEICEMANIFOLD

VDC

BRAKEDEICE C/B

N.C.N.C.

BRAKE DEICEN.C. VALVES

GEARUPLOCK

10MIN

BRAKE DEICETIMER PCB

BRAKEDEICE

DUAL FEDBUS NO.1

Figure 10-5. Brake Deice System

A minimum of 85% power on each engine isnecessary to maintain proper boot inflation ifthe hot brake system is on.

A 10-minute timer is activated when the gearis retracted, which allows sufficient time forthe brakes to dry.

The system should not be used continuouslyabove 15°C ambient temperature. Both in-strument (pneumatic) valves must be open foruse of the system.

The brake deice system is the single biggestuser of engine bleed air. During an enginefailure, the rudder boost system may be in-operative when the brake deice system is in usebecause there isn’t enough differential pres-sure to activate the system.

ICE PROTECTION—ELECTRICAL SOURCE

WINDSHIELD HEATBoth windshields are heated by resistance wireembedded in the glass. A thermal sensor withinthe lamination monitors the glass temperatureand feeds a control signal into a controller unit.The controller regulates the current flow to theembedded wire. Normally, a constant temper-ature of +95°F to +105°F is maintained (Figure10-6). However, at cold temperatures and highairspeeds, the system may not be able to main-tain an ice-free windshield.

The windshields can be operated at two heatlevels. Normal heating supplies heat to thebroadest area. High heating supplies a higherintensity of heat to a smaller but more essen-tial viewing area.

CONTROLSEach windshield heat system is separatelycontrolled by a toggle switch labeled WSHLDANTI-ICE on the pilot’s right subpanel. Eachswitch has three positions: OFF (center posi-tion), NORMAL (upper position), and HI(down position). Each switch must be lifted

over a detent before it can be moved to the HI(down) position, preventing inadvertent se-lection of the HI position when moving theswitch from NORMAL to OFF.

The two control units receive power throughtwo 5-amp control circuit breakers located ona panel on the forward pressure bulkhead, notaccessible by the crew in flight. The windowheaters are each supplied by 50-amp circuitbreakers located in the power distributionpanel under the floor forward of the main spar.

OPERATIONEither or both windshields may be heated atany time, as overheating is prevented by ther-mal sensors. Each window is fed from the leftor right generator bus through a circuit breakerlocated in the power distribution panel underthe floor forward of the main spar. The panelswitch closes a relay, which supplies currentto the windshields, subject to the control of thetemperature controller and thermal sensors.

Windshield heat may be used at any time, butit causes erratic operation of the magnetic com-pass, and could result in distorted visual cues.

PROPELLER HEATAn electrically-heated boot on each blade, de-ices the propellers (BB-2 through 815, 817-824, 991; BL 1-29, these boots were dividedinto an inner and outer segment). The boot,firmly cemented in place, receives currentfrom a slip ring and brush assembly on the pro-peller shaft. The slip ring transmits current tothe deice boot. The centrifugal force of thespinning propeller and airblast breaks the iceparticles loose from the heated blades.

Propeller deice must not be operatedwhen the propellers are static.

The boots are heated in a preset sequence, whichis an automatic function controlled by a timer.

CAUTION







STALLWARN

BRAKEDEICE


NORMAL

SURFACEDEICE

SINGLE PITOT

OFFMANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

OFF

OFF


HIOFF

HEATINGWIRES

OVERTEMPSENSOR OVERTEMP

SENSOR95° TO 105°

CB

CB

CB

CB

TEMPERATURECONTROLLER

TEMPERATURECONTROLLER

Figure 10-6. Windshield Anti-Ice System

On models BB-816, 825-990, 992 and subse-quent; BL 30 and subsequent, the following se-quence is followed:

• For 90 seconds—entire right propeller

• For 90 seconds—entire left propeller

On models prior to BB-2 through 815, 817-824, 991; BL 1-29, the most common timer isused and this sequence is followed:

• For 30 seconds—right outer elements

• For 30 seconds—right inner elements

• For 30 seconds—left outer elements

• For 30 seconds—left inner elements

Once the system is turned on for automatic op-eration, it cycles continuously.

For both versions, manual bypass of the timeris possible. Refer to LIMITATIONS in thischapter for additional information on pro-peller deicing.

Figure 10-7 shows the control and circuitbreakers for the two configurations.

CONTROLS, INDICATORS ANDOPERATIONThe propeller deice boots are controlled by acircuit-breaker type switch and a two-positionPROP toggle switch. When the possibility ofice buildup exists, the PROP AUTO switch la-beled AUTO–OFF should be set to the AUTOposition, initiating the timer sequencing ofthe boots. An ammeter labeled PROP AMPSon the copilot’s left subpanel indicates the




STALLWARN

BRAKEDEICE


NORMAL


OFFMANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

OFF

OFF


HIOFF

COLLINS

BB-2 THROUGH 815, 817-824, 991

BB-816, 825-990, 992 AND AFTER

0

10 20

30PROP AMPS

STALLWARN

BRAKEDEICE


NORMAL

PITOT

OFFMANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

OFF

OFF


HIOFF OUTER

DEICECYCLESINGLE

0

10 20

30PROP AMPS

4-BLADE PROPELLER 3-BLADE PROPELLER

Figure 10-7. Propeller Boots Heat-Control and Indicator

current flow to the propeller elements (Figure10-7).

Normal current flow within the green arc is 18to 24 amperes for all 4-bladed airplane versions(14 to 18 for 3-bladed versions). The amme-ter may flicker as the timer sequences to thenext combination of boots, but this flicker isvery difficult to see.

The ammeter should be monitored to makecertain that current flow is approximately thesame for all timer positions. Variations couldindicate that uneven heating is occurring, re-sul t ing in possible propel ler vibrat ions.However, loss of one heating element (whenthe prop ammeter indicates a less than greenarc value) does not mean that the entire sys-tem must be turned off. (Refer to the appro-priate section of the Flight Manual.)

A manual backup of the automatic sequenc-ing is installed in case the timer fails to oper-ate properly. The PROP MANUAL–OFFswitch (or on earlier aircraft as listed above,the PROP INNER–OUTER switch), providescurrent to the boots (Figure 10-7). On air-planes with a single boot element per pro-peller, with the PROP AUTO switch in theOFF position, holding the PROP MANUALswitch in the MANUAL position for approx-imately 90 seconds deices both props at thesame time, applying heat to all the boots. Onairplanes with a two-segment boot per pro-peller, the spring-loaded switch must be heldto the OUTER position until the ice has beendislodged from both propellers’ outer boots.Then it must be held to the INNER position todeice both propellers’ inner boots.

The PROP AMPS ammeter does not registercurrent flow in the MANUAL mode of oper-ation. The increased load, however, can beobserved on the airplane loadmeters.

The automatic and manual deice circuits haveseparate circuit breakers. A single circuit-breaker switch is utilized for the automaticmode and is located on the pilot’s right sub-

panel in the ice group. The manual system’scircuit breakers are located on the fuel controlcircuit-breaker panel, located on the pilot’s leftside panel in the PROP DEICE group. Thecontrol circuit breaker is for the INNER/OUTER switch, depending on the model. ThePROP LEFT and PROP RIGHT circuit break-ers control power to the prop elements in themanual mode.

Although this system is called aprop deice system, pilot manage-ment of the system should be as ananti-ice system.

PITOT HEATA heating element in each pitot probe pre-vents ice and moisture buildup. There is nothermal protection for the heating system ex-cept its own circuit-breaker switch.

CONTROLS AND OPERATIONEach pitot heater has its own circuit-breakerswitch that can be left in the ON position dur-ing flight (Figure 10-8).

The two circuit-breaker switches are fed offseparate dual-fed buses. The left is on the No.1 and the right is on the No. 2 dual-fed bus.

It is recommended that the pitot heat not be op-erated on the ground except for testing or forshort intervals to remove ice or snow fromthe mast. However, it should be turned on fortakeoff when icing conditions are suspected.

Prolonged use of pitot and stall warn-ing heat on the ground will damagethe heating elements.

CAUTION

CAUTION




STALL WARNING VANE HEATHeat is applied to both the mounting plate andthe vane. There is no thermal protection of theheating element except its own control cir-cuit-breaker switch.

Control and OperationA circuit-breaker switch labeled STALLWARN in the ICE group controls the heatingfunction. Due to the left landing gear squatswitch, the current flow to the heater is min-imal while the airplane is on the ground. Inflight, full current is supplied (Figure 10-9).

The heating elements protect the lefttransducer vane and faceplate fromice. However, a buildup of ice on thewing may change or disrupt the airflowand prevent the system from accu-rately indicating an imminent stall.

FUEL VENT HEAT

Controls and OperationElectric heaters prevent ice formation in the fuelvent system. Each wing fuel system has its ownanti-ice system, operated by the two switchesin the ICE group labeled FUEL VENT (Figure10-10). They should be used whenever icingconditions are anticipated or encountered.

A fuel heater prevents ice formation in thefuel control unit. An engine oil line withinthe fuel heater is in proximity to the fuel linesand, through conduction, a heat transfer oc-curs, melting any ice particles which may haveformed in the fuel.

WARNING




STALL WARNING VANE

HEAT CONTROLS

Figure 10-9. Stall Warning Vane andHeat Controls

PITOT PROBES

HEAT CONTROLS

Figure 10-8. Pitot Probes and HeatControls

On earlier models prior to 1979, each P3pneumatic fuel control l ine is protectedagainst ice by an electrically-heated jacketwhich receives electric current if the enginecondition levers are moved out of fuel cut-off range. On later models and all B200 air-planes, a heated jacket and a filter is installedfor this purpose.

MISCELLANEOUSSYSTEMSPOWERPLANTThe engine air inlet lips are heated by engineexhaust gases to prevent the formation of ice(Figures 10-11 and 10-12). On airplanes BB-1266, BL-129 and subsequent, hot engine ex-haust flows from the left stack, through the lip,and exits out the right stack. Prior to BB-1266,hot engine exhaust is routed downward andinto each end of the inlet lip and eventuallyducted out through the bottom of the lip.

The system is automatic and does not requirepilot action.

To prevent the engine compressor inlet screenfrom accumulating ice, an inertial vane sep-arating system is installed. When the ice vanesare lowered, they deflect the airstream slightlydownward, creating a venturi effect. At thesame time, an inertial vane bypass door underthe cowling is also opened, allowing an exit.

As the ice particles or water droplets enter theengine inlet, the airstream with these particlesis accelerated because of the venturi effect.These frozen moisture particles, due to thegreater mass and, therefore, greater momentum,accelerate past the screen area and vent overboardthrough the bypass door. However, the airstreammakes the sudden turn easier because the air isfree of the ice particles which are being de-flected rearward and overboard.

The inertial vane and the inertial vane bypassdoor are closed for normal flying conditions,thus directing the air into the powerplant in-take and oil cooler.




HEATED FUEL VENT

HEAT CONTROLS

Figure 10-10. Heated Fuel Vent andControl




OFF

ENGINE ANTI-ICE

AUTOFEATHER

ON

MAIN

OFF

ACTUATORSTANDBY

OFF

OFF

PROP GOVTESTARM

LEFT RIGHT

TEST

OFF

ARM

ENG AUTOIGNITION

PILOTAIR

PULLON

STALLWARN

BRAKEDEICE


NORMAL



OFF

OFF

LIGHTS

MANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

LEFT RIGHT

OFF

OFF


HI

DEFROSTAIR

PULLON

RIGHTLEFT

OFF

FLAPS

DOWN

TAKEOFFAND

APPROACH

UP20

60

80

4

6

2 41

1

0.5

.5

2CABIN CLIMBTHDS FT PER MIN

71

2

345

6

0PSI

CABINALT

00 FT

40 5

10

1520

25

30

35

0

LANDINGGEAR

RELAY

2

Figure 10-11. Powerplant Intake Ice Protection (BB-1439, 1444 and Subsequent)




OFF

ICE VANE

AUTOFEATHER

EXTEND

TEST

OFF

RETRACT

ICE VANEMANUAL

OFF

PROP GOV TEST

ARM

LEFT RIGHT

TEST

OFF

ARM

ENG AUTO IGNITION

PILOTAIR

PULLON STALL

WARNBRAKE


NORMAL



OFF

OFF

LIGHTS

MANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

LEFT RIGHT

OFF

OFF


HI

DEFROSTAIR

PULLON

RIGHTLEFT

OFF

FLAPS

DOWN

TAKEOFFAND

APPROACH

UP20

60

80

4

6

2 41

1

0.5

.5


71

2

345

6

0PSI

CABINALT

00 FT

40 5

10

1520

25

30

35

0

ICE VANECONTROLS

LEFTENG

RIGHTENG

PULL

MECHANICAL BACKUPT-HANDLES

OUTER

DEICE

Figure 10-12. Powerplant Intake Ice Protection (Prior to BB-1444, Except BB-1439)

CONTROLS, INDICATORS ANDOPERATIONTo extend or retract the ice vanes, the ENGANTI-ICE toggle switches (prior to BB-1444,except 1439, ICE VANE toggle switches) aremoved to the appropriate position. They arelocated on the pilot’s left subpanel (Figures 10-11 and 10-12).

In the ice-protection mode, the extended po-sition of the vane and the bypass door is in-dicated by the green annunciator lights, LENG ANTI-ICE and R ENG ANTI-ICE (priorto BB-1444, 1439; L ICE VANE EXT and RICE VANE EXT) on the caution-advisorypanel. When retracted, the lights extinguish.

In addition, two amber lights labeled L and RENG ANTI-ICE (prior to BB-1444, except1439 L ICE VANE and R ICE VANE) are pro-vided on the caution-advisory panel. If eitherengine’s inertial vane and inertial vane by-pass door have not attained the selected posi-tion (either open or closed) within 15 seconds,the appropriate light illuminates.

For BB-1439, 1444 and subsequent a backupsystem consists of dual actuators and con-trols. Illumination of the L and R ENG ANTI-ICE (amber) annunciators indicates that thesystem did not operate to the desired posi-tion. Immediate illumination of the L or RENG ANTI-ICE (yellow) annunciator indi-cates loss of electrical power, whereas de-layed illumination indicates an inoperativeactuator. In either event, the STANDBY ac-tuator should be selected.

Prior to BB-1444, except 1439, a mechani-cal backup system is provided for manuallylowering or raising the vanes. It is actuatedby pulling the T-handles just below the pilot’ssubpanel. Airspeed should be decreased to160 knots or less to reduce the forces op-posing manual operation. Normal airspeedmay be resumed after the ice vanes havebeen positioned. During manual system use,the electric switch position should matchthe manual handle position for correct an-

nunciator indication provided the appropri-ate circuit breaker has already been pulledvia the checklist. When the vane is success-fully positioned with the manual system, theamber annunciator light(s) will extinguish.

Prior to BB-1444, except 1439

Once the manual override system hasbeen engaged (i.e., any time the man-ual ice vane T-handle has been pulledout), do not attempt to retract or ex-tend the ice vanes electrically, evenif the T-handle has been pushed backin, until the override linkage in theengine compartment has been prop-erly reset on the ground. However,the pilot can raise or lower the vanesrepeatedly, any time, with the man-ua l sy s t em engaged . (See t heRaytheon Maintenance Manual forresetting procedure.)

NOTELowering the ice vanes will resultin a slight ITT rise and a significantloss of torque at normal cruise powersettings.

The circuit breakers for the ice vanes are lo-cated on the copilot’s right side panel in theengine group and are labeled MAIN ENGANTI-ICE and STBY ENG ANTI-ICE (priorto BB-1444, except 1439 only one circuitbreaker exists labeled ICE VANE CONTROL).

The movable vane and the bypass door must belowered into the airstream when operating in vis-ible moisture at +5°C or colder. Retractionshould be accomplished at +15°C and above toensure adequate engine oil cooling.

The vanes must be either retracted or ex-tended. There are no intermediate positions(Figure 10-13).

CAUTION




WINDSHIELD WIPERS

CONTROLS AND OPERATIONThe dual wipers are driven by a mechanism op-erated by a single electric motor, all locatedforward of the instrument panel.

The windshield wiper switch is located on theoverhead light control panel (Figure 10-14).It provides the wiper mechanism with twospeeds and a park position. The wipers maybe used either on the ground or in flight, asrequired. The wipers must not be operated ona dry windshield.




33N3

FO

R0

3060

90120 150 180 210

24027

030

033

0


RADIO ON

ST

EE



AIRSPEEDS (IAS)



CAUTION




PARK SLOW

FAST

OFF

MASTERPANELLIGHTS

ON

OVERHEADFLOODLIGHTS

OFFBRT


OFFBRT

AVIONICSPANELLIGHTS

OFFBRT

ENGINEINSTRUMENT

LIGHTS

OFFBRT

PILOTFLIGHTLIGHTS

OFFBRT


LIGHTS

OFFBRT

SIDEPANELLIGHTS

OFFBRT


LIGHTS

OFFBRT

COPILOTFLIGHTLIGHTS

OFFBRT

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

40 60 8020

0 3010 20

0

PUSH

FOR VOLTS

100

DC VOLTS

% LOAD

Beechcraft

400 410420380

390

100 130110 120

PUSH

FOR VOLTSAC VOLTS

FREQ

Beechcraft



PARK SLOW

FAST

0 8020

0 3020

0

PUSH

FOR VOLTS

100

DC VOLTS

Beechcraft

-60

0

+60BATT AMPS

Figure 10-14. Windshield Wiper Control

Figure 10-13. Engine Intake Inertial Vane Positions and Bypass Door

INERTIAL VANE RETRACTED INERTIAL VANE BYPASS DOOR EXTENDED

INERTIAL VANE BYPASS DOOR EXTENDED

Windshield wipers may be damagedif used on a cracked outer panel.

The circuit breaker is on the copilot’s right CBpanel in the WEATHER group.

WING ICE LIGHTS

LOCATION AND CONTROLThe wing lights are located on the outboardside of each nacelle. The circuit-breaker switchis located on the pilot’s right subpanel in theLIGHTS group above the ICE group (Figure10-15).

LIMITATIONSSafe operation in icing conditions is dependentupon pilot knowledge regarding atmosphericconditions conducive to ice formation, famil-iarity with the operation and limitations of theinstalled equipment, and the exercise of goodjudgment when planning a flight into areaswhere possible icing conditions might exist.

When icing conditions are encountered, theperformance characteristics of the airplanewill deteriorate.

Increased aerodynamic drag increases fuelconsumption, thereby reducing the airplane’srange and making it more difficult to maintainspeed.

Decreased rate of climb must be anticipated,not only because of the decrease in wing andempennage efficiency, but also because ofthe possible reduced efficiency of the pro-pellers and increase in gross weight. Also,the use of the inertial ice vanes may result inlost performance.

Abrupt maneuvering and steep turns at lowspeeds must be avoided because the airplanewill stall at higher than published speeds withice accumulation. On final approach for land-ing, increased airspeed must be maintainedto compensate for this increased stall speed.After touchdown with heavy ice accumula-tion, landing distances may be as much astwice the normal distance due to the increasedlanding speed.

During descent, a minimum of 85% power oneach engine is necessary to maintain properboot inflation if the airplane is equipped withand using the hot brake system.

Use of the brake deice system in flight will re-sult in an ITT rise of approximately 20°C.ITT limitations must be observed when settingclimb and cruise power.

The brake deice system should not be operatedcontinuously above +15°C OAT.

CAUTION




WING ICE INSPECTION LIGHT

CONTROL

Figure 10-15. Wing Ice Inspection Lightand Control

If the landing gear is retracted, the systemmay not be operated longer than 10 minutes,which is one timer cycle. The annunciatorlight should be monitored. If it does not au-tomatically go out after approximately 10 min-utes following gear retraction, the systemshould be manually turned off.

Both engine bleed-air sources must be in op-eration to use the brake deice system on bothsides.

A minimum speed of 140 KIAS is necessaryto prevent ice formation on the underside ofthe wing, which cannot be adequately deiced.

Windshield heat may be used at any time, butit causes erratic operation of the magnetic com-pass, and could result in distorted visual cues.

Windshield wipers may be damagedif used on a cracked outer panel.Heating elements may be inopera-tive in area of crack.

During sustained icing conditions, 226 KIASis the maximum effective airspeed due to thelimitations of the windshield heating system.

In flight, the boots should be cycled onceevery time the ice accumulation is approxi-mately one-half to one inch thick.

Should either engine fail in flight, there is suf-ficient air for the entire deice operation (ex-cept for the hot brake operation). Should theautomatic cycling of the boots fail, the MAN-UAL position should be used for inflation.

While in flight, the engine ice vanes must beextended and the appropriate annunciatorlights monitored:

• Before visible moisture is encounteredat OAT +5°C and below.

• At night when freedom from visiblemoisture is not assured and the OAT is+5°C or below.

If the amber ENG ANTI-ICE (prior to BB-1444, except 1439, ICE VANE) annunciatorsilluminate upon extension (Figures 10-11 and10-12), the ice vanes may not have positionedproperly.

The STBY actuators (or prior to BB-1444,except 1439, manual control) should be usedto retract or to extend them. A reliable backupcheck on the position is to closely monitorengine torque. Normal torque may be re-gained with the power levers, observing theITT limits.

If in doubt, extend the vanes. Engineicing can occur even though no sur-face icing is present. If freedom fromvisible moisture cannot be assured,engine ice protection should be ac-tivated. Visible moisture is moisturein any form: clouds, ice crystals,snow, rain, sleet, hail, or any com-bination of these. Ice vanes should beretracted at +15°C and above to as-sure adequate engine oil cooling.Operation of strobe lights will some-times show ice crystals not normallyvisible.

Prior to BB-1444, except 1439, once the icevanes have been actuated manually, do not at-tempt to retract or extend them electricallyuntil they have been reset, as this may causedamage to the system.

During flight in icing conditions, fuel ventheat, pitot heat, prop deice, windshield heat,and stall warning heat should all be on.

The wing ice lights should be used as re-quired in night flight to check for wing iceaccumulation.

CAUTION

CAUTION




Due to distortion of the wing air-foil, stalling airspeeds should beexpected to increase as ice accu-mulates on the airplane. For thesame reason, stall warning devicesare not accurate and should not berelied upon. Maintain a comfort-able margin of airspeed above thenormal stall airspeed when ice is onthe airplane. In order to prevent iceaccumulation on unprotected sur-faces of the wing, maintain a min-i m u m o f 1 4 0 k n o t s d u r i n goperations in sustained icing con-ditions. In the event of windshieldicing, reduce airspeed to 226 KIASor below to ensure maximum wind-shield heat effectiveness.

NOTEThe wing ice lights operate at a hightemperature and therefore should notbe used for prolonged periods whilethe airplane is on the ground.

If either BLEED AIR FAIL light illuminatesin flight, the bleed-air switch on the affectedengine must be closed to the INST & ENVIROFF position. This will isolate the brake deicesystem on that side. Therefore, the brake deicesystem must be selected to OFF. BLEED AIRFAIL lights may momentarily illuminate dur-ing simultaneous wing boot and brake deiceoperation at low N1 speeds. If lights immedi-ately extinguish, they may be disregarded.

The wipers must not be operated on a drywindshield.

Windshield wipers may be damagedif used on a cracked outer panel.

While in flight, the propeller deice systemmay be operated continuously in automaticmode without overheating.

Propeller deice must not be operatedwhen the propellers are static.

The PROP AMPS should read 18 to 24 amperesfor 4-bladed models and 14 to 18 amperes for3-bladed models. Procedures differ for vari-ous abnormal readings on the PROP AMPSammeter.

For a reading of zero amperes, the PROPAUTO switch should be checked to ensurethat it is on. If it is off, it should be repositionedto ON after 30 seconds have elapsed. If turnedon with no current flow, the PROP AUTOswitch must be turned OFF and the manualbackup system used to supply current to thepropellers.

For a reading below the green arc, use of thePROP AUTO switch may be continued eventhough one or more boots is probably not heat-ing. If propeller imbalance occurs, rpm mustbe increased briefly to aid in ice removal.

For a reading higher than the green arc, nor-mal automatic operation may be continuedunless the circuit-breaker switch trips. If theautomatic circuit breaker does not trip, auto-matic deicing may be continued. If propellerimbalance occurs, rpm must be increasedbriefly to aid in ice removal. If the circuitbreaker switch trips, the manual backup sys-tem must be used and the loadmeter moni-tored for excessive current flow. If the manualcircuit breaker(s) trip, icing conditions shouldbe avoided as soon as possible.

CAUTION

CAUTION

CAUTION




NOTESFor manual backup on airplanes BB-816, 825-990, 992 and subsequent;BL-30 and subsequent, the switch isheld to the ON position for approxi-mately 90 seconds. This backup sys-tem may be repeated as required andthe loadmeter should be monitoredfor a deflection of approximately 5%.

For manual backup on airplanesprior to BB-2 through 815, 817-824,991; BL-1 through 29, the PROPINNER/OUTER swi tch i s pos i -tioned first to the OUTER, then tothe INNER position for 30 secondsin each position and the loadmetermonitored for a deflection of ap-proximately 5%.




11-i

CHAPTER 11AIR-CONDITIONING SYSTEM

CONTENTS

Page

INTRODUCTION ................................................................................................................. 11-1

GENERAL............................................................................................................................. 11-1

System Description and Location .................................................................................. 11-1

Air-Conditioning System Controls ................................................................................ 11-8

LIMITATIONS .................................................................................................................... 11-11

QUESTIONS....................................................................................................................... 11-12




11-iii

ILLUSTRATIONS

Figure Title Page

11-1 Super King Air Air-Conditioning System, BB-1180 and After (With Aft Evaporator)............................................................................................. 11-2

11-2 Super King Air Air-Conditioning System Prior to BB-1180(With Aft Evaporator)............................................................................................. 11-3

11-3 AIR CND N1 LOW Advisory Light ...................................................................... 11-4

11-4 Floor and Ceiling Outlets ....................................................................................... 11-4

11-5 Cockpit “Eyeball” Outlets ...................................................................................... 11-5

11-6 Receiver-Dryer Sight Gage .................................................................................... 11-5

11-7 Air Control Knobs .................................................................................................. 11-6

11-8 DUCT OVERTEMP Caution Light ....................................................................... 11-6

11-9 ELECTRIC HEAT Switch...................................................................................... 11-7

11-10 RADIANT HEAT Switch and Panel ...................................................................... 11-7

11-11 ENVIRONMENTAL Group Switches and Knobs ................................................. 11-8

11-12 CABIN TEMP MODE Selector Switch ................................................................ 11-8

11-13 Air-Conditioning System Control Diagram ........................................................... 11-9

11-14 Ram-Air Scoop ..................................................................................................... 11-11




INTRODUCTIONThe Super King Air’s air-conditioning system (Figures 11-1 and 11-2) provides thecrew and passengers with cooling, heating and unpressurized ventilation. In addition tothe heating afforded by the air-conditioning system, electric heat (radiant heat prior toBB-1444, except 1439) is available as an option. The air-conditioning system may beoperated in the heating mode and the cooling mode either under automatic mode con-trol or manual mode control.

GENERAL

SYSTEM DESCRIPTION ANDLOCATION

CoolingCabin cooling is provided by a refrigerantgas, vapor-cycle refrigeration system. Thissystem consists of the following components:

• A belt-driven, engine-mounted com-pressor (right engine)

• Refrigerant plumbing

• N1 speed switch

• High- and low-pressure protect ionswitches

• A condenser coil

• A condenser blower

CHAPTER 11AIR-CONDITIONING SYSTEM







FLOOR DUCTTO AFT FLOOROUTLETS

TO CEILINGOUTLETS

FWD

SIDEVIEW

DOOR

DETAIL A

AIR-CONDITIONED AIRFROM AFT EVAPORATOR

NORMAL OUTFLOW VALVE

SAFETY/DUMP VALVEAFT PRESSUREBULKHEAD* NOTE

FORWARD AND AFT HEATERFOR BB-1439, 1444 AND AFTER

CEILINGOUTLET

HOT ENGINE BLEED AIR


RECIRCULATED CABIN AIR(AIR CONDITIONED WHENEVAPORATOR IS ON)

AMBIENT AIR

PRESSURE VESSEL

LEGENDFLOOR OUTLET

CEILING OUTLETS

AIR-TO-AIRHEAT

EXCHANGER

CABIN-HEATCONTROL VALVE

FLOOROUTLET

CEILINGOUTLET

AIR INLETSCOOP

FIREWALL

PNEUMATICBLEED-AIR

SHUTOFFVALVE

AMBIENT AIRMODULATING

VALVE

PNEUMATICTHERMOSTAT

CONDENSER BLOWER

OUTLET AIR

RECEIVER-DRYERMIXING PLENUM

WINDHSHIELD DEFROSTER(ON GLARESHIELD)

FWD PRESSURE BULKHEAD

CREWHEAT DUCT

INSTRUMENT PANEL

PILOT'S VENTAIR CONTROL

WINDSHIELD DEFROSTER

CONTROL

ENVIRONMENTAL BLEED-AIR SHUTOFF VALVE

ENVIRONMENTAL BLEED-AIR FLOW CONTROL UNITINCLUDING MODULATINGAND SHUTOFF VALVE

FLOOROUTLET

DOOR (COOLED AIR TOFLOOR OUTLETS)

CEILING OUTLET

AFT EVAPORATORAIR FILTER

AFT EVAPORATOR

CEILING OUTLET

FLOOR OUTLET

CEILING OUTLET


AIR-TO-AIRHEAT EXCHANGER

FIREWALL

AMBIENT AIRMODULATINGVALVE

ENVIRONMENTALBLEED-AIRSHUTOFF VALVE

PNEUMATICTHERMOSTAT

REFRIGERANTCOMPRESSOR

CONDENSER

RAM-AIR SCOOP

FRESH AIR VALVE(CLOSED WHEN PRESSURIZED)

VENT BLOWER

FWD EVAPORATORAIR FILTER

FWD EVAPORATOR

RETURN AIR VALVE

RETURN AIR FILTER

COPILOT'SVENT AIR CONTROL

CABIN AIR CONTROL

CEILING DUCT/FLOOR DUCT DIVIDER

COPILOT'SCEILING OUTLETDUCT OVERTEMPSENSOR

CABIN AIRCONTROL VALVE

REFRIGERANT LINESAIR INLET

FLAPPERVALVE

FORWARD HEATER *

AFT HEATER *

PNEUMATIC BLEED-AIRSHUTOFF VALVE


Figure 11-1. Super King Air Air-Conditioning System, BB-1180 and After(With Aft Evaporator)




FLOOR DUCTTO AFT FLOOROUTLETS

TO CEILINGOUTLETS

FWD

SIDEVIEW

DOOR

DETAIL A

AIR-CONDITIONED AIRFROM AFT EVAPORATOR

NORMAL OUTFLOW VALVE

SAFETY/DUMP VALVEAFT PRESSUREBULKHEAD

CEILINGOUTLET

HOT ENGINE BLEED AIR


RECIRCULATED CABIN AIR(AIR CONDITIONED WHENEVAPORATOR IS ON)

AMBIENT AIR

PRESSURE VESSEL

LEGENDFLOOR OUTLET

CEILING OUTLETS

AIR-TO-AIRHEAT

EXCHANGER


FLOOROUTLET

CEILINGOUTLET

AIR INLETSCOOP

FIREWALL

PNEUMATICBLEED-AIR

SHUTOFFVALVE

AMBIENT AIRMODULATING

VALVE

PNEUMATICTHERMOSTAT

CONDENSER BLOWER

OUTLET AIR

RECEIVER-DRYERMIXING PLENUM

WINDHSHIELD DEFROSTER(ON GLARESHIELD)

FWD PRESSURE BULKHEAD

CREWHEAT DUCT

INSTRUMENT PANEL

PILOT'S VENTAIR CONTROL

WINDSHIELD DEFROSTER

CONTROL

ENVIRONMENTAL BLEED-AIR SHUTOFF VALVE


FLOOROUTLET

DOOR (COOLED AIR TOFLOOR OUTLETS)

CEILING OUTLET

AFT EVAPORATORAIR FILTER

AFT EVAPORATOR

CEILING OUTLET

FLOOR OUTLET

FLAPPER VALVE


AIR-TO-AIRHEAT EXCHANGER

FIREWALL

AMBIENT AIRMODULATINGVALVE

ENVIRONMENTALBLEED-AIRSHUTOFF VALVE

PNEUMATICTHERMOSTAT

REFRIGERANTCOMPRESSOR

CONDENSER

RAM-AIR SCOOP

FRESH AIR VALVE(CLOSED WHEN PRESSURIZED)VENT BLOWER

FWD EVAPORATORAIR FILTER

FWD EVAPORATOR

RETURN AIR FILTERRETURN AIR VALVE

COPILOT'SVENT AIR CONTROL

CABIN AIR CONTROL

CEILING DUCT/FLOOR DUCT DIVIDER

COPILOT'SCEILING OUTLET

DUCT OVERTEMPSENSOR

CABIN AIRCONTROL VALVE

REFRIGERANT LINESAIR INLET

CEILINGOUTLET


PNEUMATIC BLEED-AIRSHUTOFF VALVE

Figure 11-2. Super King Air Air-Conditioning System Prior to BB-1180 (With Aft Evaporator)

• An evaporator with an optional aft evaporator.

• A receiver-dryer.

• An expansion valve or two expansionvalves if aft evaporation is installed.

• A bypass valve.

The plumbing from the compressor, which ismounted on the right engine, is routed throughthe right wing and then forward to the con-denser coil, receiver-dryer, expansion valve,bypass valve, and evaporator—all of which arelocated in the nose of the airplane.

The high- and low-pressure limit switches andthe N1 speed switch (engine speed) preventcompressor operation outside of establishedlimitation parameters. The N1 speed switchdisengages the compressor clutch when the en-gine speed is below 62% N1 and air condi-tioning is requested. When the N1 speed switchopens, and if air conditioning is being re-quested, the green AIR CND N1 LOW advi-sory annunciator (Figure 11-3) will illuminate.

The forward vent blower moves recirculatedcabin air through the forward evaporator, intothe mixing plenum, into the floor-outlet ducts,and ceiling eyeball outlets. Approximately75% of the recirculated air passes through thefloor outlets while approximately 25% of theair is routed through the ceiling outlets, by-passing the mixing plenum (Figure 11-4).

The forward vent blower, with the system inAUTOmatic normally runs at low speed.

If the cooling mode is operating, refrigerantcirculates through the forward evaporator,cooling the output air. All the air entering theceiling-outlet duct is cooler than the air en-tering through the f loor outlets if ei therBLEED AIR VALVE switch is in the OPEN po-sition. This air discharges through “eyeball”outlet nozzles (Figure 11-5) in the cockpitand cabin. Each nozzle is movable so theairstream may be directed as desired. Also, thevolume of air can be adjusted from full opento closed by twisting the nozzle. As the noz-zle is twisted, a damper opens or closes toregulate airflow.

Cool air also enters the floor-outlet duct, butin order to provide cabin pressurization, warmbleed air also enters this duct any time eitherBLEED AIR VALVE switch is in the OPEN po-sition. Therefore, pressurized air dischargedfrom the floor outlets is always warmer thanthe air discharged from the ceiling outlets, nomatter what temperature mode is used.




Figure 11-3. AIR CND N1 LOW AdvisoryLight

Figure 11-4. Floor and Ceiling Outlets




NOTEOn the Super King Air 200, prior to BB310 and all cargo door airplanes, alever on each floor outlet register (ex-cept the forward facing register in thebaggage compartment) can be movedvertically to regulate the airflow. OnBB 310, 343 and all subsequent pas-senger door models, this feature hasbeen deleted. A vane-axial blower inthe nose section draws ambient airthrough the condenser to cool the re-frigerant gas when the cooling modeis operating. On Serial Nos. BB-345and subsequent and BL-1 and subse-quent (and any earlier serials that havecomplied with Beechcraft ServiceInstructions No. 0968 by the instal-lation of Kit Number 101-5035-1 S or101-5035-3 S), this blower shuts offwhen the gear is retracted.

The receiver-dryer and sight gage (glass) arelocated high in the condenser compartment.

The crew can view these components by re-moving the upper-compartment access panel,located on top of the nose section left of cen-terline. This, however, is not a normal preflightaction. If there are bubbles seen through thesight glass (Figure 11-6), the refrigerant sys-tem is low on the refrigerant gas being used.If, after adding more refrigerant gas, bubblesare still appearing, then the system needs tobe evacuated and recharged.

An optional aft evaporator and blower isavailable for additional cooling. It is locatedbelow the center aisle cabin floor behind therear spar. The additional unit increases theairplane’s cooling capacity from 18,000 Btu(with the forward evaporator only) to 32,000Btu. Refrigerant flows through the aft evap-orator any time it flows through the forwardevaporator; however, the additional coolingis provided only when the aft blower is op-erating, recirculating cabin air through theaft evaporator, and routing it to the aft floorand ceiling outlets.

Figure 11-5. Cockpit “Eyeball” Outlets

RECEIVER-DRYER AND

SIGHT GAGE

Figure 11-6. Receiver-Dryer Sight Gage

HeatingBleed air from the compressor of each engineis delivered into the cabin for heating, as wellas pressurization. When the left landing gearsafety switch is in the ground position, theambient air valve in each flow control unit isclosed; therefore, only bleed air is delivered.When airborne, bleed air is mixed with out-side ambient air from the ambient air valve ineach flow control unit until a cold air tem-perature closes off the ambient flow. Thenonly bleed air is delivered.

In the cockpit, additional air can be providedby adjusting either the pilot’s damper, whichis controlled by the PILOT AIR knob (Figure11-7), or the copilot’s damper, which is con-trolled by the COPILOT AIR knob. Movementof these knobs affects cockpit temperature byadjusting the air volume (Figure 11-7). TheCABIN/COCKPIT AIR knob (simply CABINAIR prior to BB-1444, except 1439) controlsair volume to the cabin (Figure 11-7) and is lo-cated on the copilot’s left subpanel below andinboard of the control column. This knob con-trols the cabin air control valve. When thisknob is pulled out of its stop, a minimumamount of air passes through the valve to thecabin, thus increasing the volume of air avail-able to the pilot and copilot outlets and the

defroster. When the knob is pushed all the wayin, the valve opens, allowing the air in the ductto be directed into the cabin floor outlets.

The DEFROST AIR knob (Figure 11-7) con-trols a valve on the pilot/copilot heat ductwhich admits air to two ducts that deliver thewarm air to the defroster, located below thewindshields and at the top of the glareshield.

The rest of the air in the bleed-air duct mixeswith recirculated cabin air and is routed aftthrough the floor-outlet duct, which handles75% of the total airflow. If the airflow be-comes too low in the ducting, the amber DUCTOVERTEMP caution light (Figure 11-8) illu-minates, indicating that the duct temperaturehas reached approximately 300°F.




OFF

ICEEMERGENCY

EXTENSIONLEFT ENG

PILOTAIR

PULLON

STALLWARN

BRAKEDEICE


NORMAL

SURFACEDEICE

SINGLE PITOT

LANDING TAXI ICE NAV

OFF

OFF

LIGHTS

MANUAL

PILOT COPILOT

LEFT RIGHT

LEFT

LEFT RIGHT

OFF

OFF

AUTO MANUAL

HI

DEFROSTAIR

PULLON

OFF

COPILOTAIR

PULLON

INCR

ENVIRONMENTAL

HIGH

AUTO

LO

DECR

INCR

CABIN TEMP

CABIN TEMP MODE

OFF

MANUALTEMP

VENTBLOWER

MAINHEAT

AUTO

MAINCOOL CABIN

AIR

PULLDECR

ELECHEAT

OFFOFF

OFF

AFTBLOWER

ON

INSTR & ENVIR OFF

ENVROFF

OPENRIGHT

BLEED AIR VALVES


LEXTR

LDET

R

DEFROST AIR KNOB CABIN AIR KNOB COPILOT AIR KNOBPILOT AIR KNOB

Figure 11-7. Air Control Knobs

Figure 11-8. DUCT OVERTEMP CautionLight

Electric Heat (BB-1439, 1444and Subsequent)A supplemental electric heating system isavailable for cabin comfort. It is operated bya solenoid-held switch on the copilot’s leftsubpanel placarded ELEC HEAT–OFF (Figure11-9). This system can be used in conjunctionwith an external power unit for warming thecabin prior to starting the engines, and is usedin the manual heat or automatic temp controlmode only.

This system uses one forward heating elementlocated in a forward duct and one aft heating el-ement located in the aft evaporator plenum.Both the forward and the aft blower must be op-erating during electric heat operation. An ELECHEAT ON advisory annunciator is provided toindicate that the power relays are in the closedposition to apply electrical power to the heat-

ing elements. When the electric heat system isselected to OFF, the ELEC HEAT ON annun-ciator must be extinguished to indicate poweris removed from the heating elements before theblowers are switched to OFF.

NOTEThe electric heat system will drawapproximately 300 amps.

The system is available for ground operationonly. If the aircraft takes off with the electricheat ON, the squat switch will remove electricpower via the solenoid-operated electric heatswitch and the switch goes to the OFF posi-tion.

Radiant Heating (Prior to BB-1444except BB-1439)An optional electric radiant heating system isavailable for the Super King Air. This system isturned on or off by the RADIANT HEAT switch(Figure 11-10) located in the ENVIRONMENTALgroup on the copilot’s left subpanel. This systemuses overhead heating panels to warm the cabinprior to engine start, as well as to provide sup-plemental heat in flight (Figure 11-10). Duringground operations when using radiant heat, the useof an auxiliary power unit is highly encouraged.




Figure 11-9. ELECTRIC HEAT Switch

RADIANT HEAT SWITCH RADIANT HEAT PANELS

Figure 11-10. RADIANT HEAT Switch and Panel

NOTEThe radiant heating system shouldbe used with the manual tempera-ture control mode only.

AIR-CONDITIONING SYSTEMCONTROLSThe ENVIRONMENTAL control section(Figure 11-11) on the copilot’s left subpanelprovides automatic or manual control of theair-conditioning system. This section con-tains all the major controls of the environ-mental function, which are:

• BLEED AIR VALVE switches.

• Fo rward VENT BLOWER con t ro lswitch.

• AFT (evaporator) BLOWER ON/OFFswitch (if installed).

• ELECTRIC HEAT switch (if installed,BB-1439, 1444 and subsequent).

• RADIANT HEAT switch (if installed,prior to BB-1444, except 1439).

• MANUAL TEMPerature switch.

• CABIN TEMPerature level controlswitch.

• CABIN TEMP MODE selector switch.

Automatic Mode ControlWhen the CABIN TEMP MODE selectorswitch (Figure 11-12) is in the AUTO position,the air delivery system (Figure 11-13) oper-ates automatically to establish the temperaturerequested by the pilot. To reach the desiredtemperature setting, the automatic temperaturecontrol modulates the bypass valves and airconditioning compressor. For greater heating,bleed air is allowed to bypass the air-to-airheat exchangers in the wing center sections. Forgreater cooling, the bleed air is allowed to




PRIOR TO BB-1444, EXCEPT 1439 AFTER BB-1439, 1444 AND SUBSEQUENT

Figure 11-11. ENVIRONMENTAL Group Switches and Knobs

Figure 11-12. CABIN TEMP MODE SelectorSwitch

pass through the air-to-air heat exchangers toreduce its temperature. In either case, the re-sultant bleed air is mixed with recirculatedcabin air (which can be additionally cooled ifthe air conditioning compressor is activatedin the cooling mode) in the forward mixingplenum.

The CABIN TEMP level control (Figure11-11) provides regulation of the temper-ature level in the AUTO mode. The pilot canadjust the temperature in the aircraft byturning the CABIN TEMP level control asrequired. A temperature-sensing unit behindthe first set of passenger oxygen masks (orBB-54 through BB-310, in the cockpit ceil-ing; and prior to BB-54, in the lower leftside cabin wall), in conjunction with thetemperature level setting, initiates a heat orcool command to the temperature controller.

Manual Mode ControlWhen the CABIN TEMP MODE selectorswitch is in the MAN COOL or MAN HEATposition, regulation of the cabin temperatureis accomplished manually by momentarilyholding the MANUAL TEMP switch (Figure11-11) to either INCRease or DECRease po-sition as desired.

When released, this switch returns to the cen-ter (OFF) position. When held in either posi-tion, it results in modulation of the bypassvalves in the bleed-air lines. The pilot shouldallow one minute (30 seconds per valve) forboth valves to move fully open or fully closed.Only one valve moves at a time to vary theamount of bleed air routed through the air-to-air heat exchanger, thus causing a variance inbleed-air temperature. This bleed air mixeswith recirculated cabin air in the mixingplenum and is then routed to the floor regis-ters. Therefore, the cabin temperature varies




LEFT ENGINEBLEED AIR

RIGHT ENGINEBLEED AIR

AIR CONDITIONER

MANCOOL

MANHEAT

AUTO

OFF

CABIN TEMP MODE

MANUALTEMPINCR

DECR

AUTO TEMPCONTROLLER

DUCT CABIN SELECTOR

RH BYPASSVALVE MOTOR

LH BYPASSVALVE MOTOR

LH BYPASSVALVE MOTOR

SWITCH

COOL

COOL

MANUALHEATOR COOL

HEAT

HEAT

AUTO

MANUALCOOL

TO CABIN

AIR TO AIRHEATEXCHANGER

TO CABIN

AIR TO AIRHEATEXCHANGER

TEMPSENSORS

Figure 11-13. Air-Conditioning System Control Diagram

according to the position of the cabin-heatcontrol valves whether or not the refrigerantsystem is working.

NOTEThe air-conditioner compressor doesnot operate unless the bypass valvesare closed. To ensure that the valvesare closed, select MAN COOL thenhold the MANUAL TEMP switch inthe DECR position for one minute.

Airflow ControlFour additional manual controls on the sub-panels may be used to partially regulate cock-pit comfort when the cockpit partition door isclosed and the cabin comfort level is satis-factory (Figure 11-7). These controls are:

1. PILOT AIR CONTROL KNOB

2. DEFROST AIR CONTROL KNOB

3. CABIN AIR CONTROL KNOB

4. COPILOT AIR CONTROL KNOB

When these control knobs are fully pulled out,they provide maximum airflow to the cockpit;when fully pushed in, they provide minimumairflow. During flights in warm air, such asshort, low-altitude flights in the summer, allthe cabin ceiling outlets should be fully openfor maximum cooling. During high-altitudeflights, cool-night flights, and flights in coldweather, the ceiling outlets should be closedfor maximum cabin heating.

Bleed-Air ControlThe BLEED AIR VALVE switches control thebleed air entering the cabin (Figure 11-11). Formaximum cooling on the ground, place theswitches in the ENVIR OFF position.

Vent Blower ControlThe VENT BLOWER switch located in theENVIRONMENTAL group, controls the for-ward vent blower (Figure 11-11). This switchhas three positions, HI–LO–AUTO. When itis in the AUTO position, the blower operatesat a low speed if the CABIN TEMP MODE se-lector switch is in any position other thanOFF. When the VENT BLOWER switch is inAUTO and the mode selector is in the OFF po-sition, the blower ceases to operate. Any timethe blower switch is in the LO position, theblower operates at low speed, even if the modeselector is in the OFF position. Likewise, if theblower switch is in the HI position, the bloweroperates at high speed even if the mode selectoris in the OFF position.

If the optional aft evaporator unit is installedin the airplane, an aft blower is also installedunder the floor next to the evaporator in therear of the cabin. The aft blower, which drawsin cabin air, blows it across the evaporatorand to the aft floor and ceiling outlets; it op-erates at high speed only. The AFT BLOWERswitch (Figure 11-11), located in the ENVI-RONMENTAL group, controls the blower,which is independent of any other control.

NOTEIf the aft blower is turned on duringthe heating mode of operation, thedoor between the aft-blower duct andthe warm air (floor-outlet) duct opens.This stops the flow of bleed air to theaft floor registers and delivers recir-culated cabin air (which comes fromunder the floor and will be coolerthan cabin air) to the aft floor regis-ters and ceiling outlets (DETAIL A,Figures 11-1 and 11-2). For BB-1439,1444 and subsequent, both the ventblower and aft blower must be oper-ating if electric heat is on.




On Super King Air 200 airplanes, serials priorto BB-39, some airplanes were delivered witha two-speed aft blower which did not have a sep-arate AFT BLOWER switch, but was controlledby the forward VENT BLOWER switch and atemperature sensor. In this installation, aftblower operation is entirely automatic and can-not be controlled by the pilot (Figure 11-11).

Unpressurized VentilationFresh air is available during unpressurized flightwith the CABIN PRESS switch in the DUMPposition. This ambient (ram) air is obtained viathe fresh air door and the ram-air scoop in theairplane nose section (Figure 11-14). This dooris open only during unpressurized flight whenthe switch is in the DUMP position and there is0 psid. This allows the forward blower to drawram air into the cabin. This air is mixed with re-circulated cabin air in the plenum chamber andthen directed to both the floor registers andceiling outlets.

On early Super King Air 200 models, the vol-ume of air from the registers is regulated by mov-ing a sliding handle (lever) at the side of eachinboard-facing register. On BB 310, 345 andafter on the Super King Air B200 the air vol-ume is regulated by the CABIN AIR controlknob (Figure 11-7).

NOTEA flight conducted with the bleed-airswitches placed in any position otherthan OPEN will also result in un-pressurized flight, but the fresh airdoor will not be open.

LIMITATIONSThe following limits are imposed upon theair-conditioning system:

• Underpressure limit—2.5 psi (whichdisengages the air-conditioning clutch).

• Overpressure limit, forward evapora-tor—290 psi (which disengages the com-pressor clutch).

• Overpressure limit with aft evaporator—340 psi (which disengages the com-pressor clutch).

NOTEPrior to Serial No. 345, a 50°F OATswitch is located on top of the con-denser for auto only:

• OAT above 50°F, condenser motor runs.

• OAT below 50°F, condenser motor stops.

• After Serial No. 345, the nose gear limitswitch stops the condenser motor whenthe gear is up.

• Air-conditioning system rated at 18,000BTU with forward blower only.

• Air-conditioning system rated at 32,000BTU with aft blower and forward bloweroperating.

• AIR CND N1 LOW illuminates if rightengine is below 62% N1 speed and sys-tem is requesting air conditioning.

• Air-conditioning system current drawis 80 amps.




Figure 11-14. Ram-Air Scoop

12-i

CHAPTER 12PRESSURIZATION

CONTENTS

Page

INTRODUCTION................................................................................................................. 12-1

GENERAL ............................................................................................................................ 12-1

System Description and Location .................................................................................. 12-3

Operation ....................................................................................................................... 12-5

Preflight Operation ........................................................................................................ 12-8

In-Flight Operation ........................................................................................................ 12-8

Descent and Landing Operation .................................................................................... 12-8

LIMITATIONS ...................................................................................................................... 12-9

QUESTIONS....................................................................................................................... 12-10




12-iii

ILLUSTRATIONS

Figure Title Page

12-1 Pressurization Controls .......................................................................................... 12-2

12-2 Electronic Flow Control Unit (BB-1180 and Subsequent, BL-71 and Subsequent) ............................................. 12-3

12-3 Pneumatic Flow Control Unit (Prior to BB-1180, Prior to BL-71) ....................... 12-4

12-4 Outflow Valve ........................................................................................................ 12-6

12-5 Safety Valve ........................................................................................................... 12-6

12-6 Pressurization Controller........................................................................................ 12-7

12-7 Cabin Altimeter...................................................................................................... 12-7

12-8 CABIN CLIMB Indicator ...................................................................................... 12-7

12-9 CABIN PRESSURE Switch .................................................................................. 12-8

12-10 ALT WARN Annunciator....................................................................................... 12-8

TABLES

Table Title Page

12-1 Pressurization Controller Setting for Landing ....................................................... 12-9




INTRODUCTIONOn Super King Air 200s, BB-2 through BB-194, the pressurization system is designedto provide a normal working pressure differential of 6.0 ± 0.1 psi, which provides cabinpressure altitudes of approximately 3,900 feet at an altitude of 20,000 feet, 9,900 feetat 31,000 feet, and 11,700 feet at 35,000 feet. The normal working pressure differentialfor Super King Air 200s, Serial Nos. BB-195 up to the B200, is 6.1 psi.

On Super King Air B200 airplanes, the pressurization system is designed to provide anormal working pressure differential of 6.5 ± 0.1 psi, which provides cabin pressure al-titudes of approximately 2,800 feet at 20,000 feet, 8,600 feet at 31,000 feet, and 10,400feet at 35,000 feet.

GENERALPressurization is regulated through a pres-surization controller, monitored by a cabinaltimeter/psid indicator, and a rate-of-climbindicator. Pressurization can be dumped bya CABIN PRESSure switch. These compo-

nents are mounted near the throttle quad-rant. Additional components are a vacuumline drain and the outflow and safety valves(Figure 12-1).

CHAPTER 12PRESSURIZATION







FILTER

FLOW CONTROLPRESSURE

MOISTUREACCUMULATION

DRAIN

PLUG

STATIC

STATIC

OUT-FLOWVALVE

SAFETYVALVE

RATE ALTITUDE

RESTRICTOR

VACUUM SOURCEFROMPNEUMATICMANIFOLD

CONTROL SWITCHCABIN PRESSURE

DUMP SOLENOIDNC

CABIN PRESETSOLENOID

NO LGSAFETYSWITCH

LEGEND

CABIN AIR

VACUUM SOURCE

STATIC AIR

CONTROL PRESSURE

INTERNAL PRESSURE

Figure 12-1. Pressurization Controls




SYSTEM DESCRIPTION ANDLOCATIONBleed air from each engine is used to pressurizethe pressure vessel (cabin and cockpit areas).A flow control unit in each of the engine na-celles controls the volume of the bleed air andcombines ambient air with it to provide a suit-able air density for pressurization. The BLEEDAIR VALVE switch in the ENVIRONMEN-TAL group on the copilot’s left subpanel con-trols the flow control unit (Figure 12-2). Whenthis switch is either in the ENVIR OFF or theINSTR & ENVIR OFF position, the flow con-trol unit is closed. These switch positions willalso illuminate the green L or R BL AIR OFFadvisory annunciator light. When it is in theOPEN position, the mixture of engine bleedair and ambient air flows through the flowcontrol unit and through or around the air-to-air heat exchangers.

Electronic Flow Control Unit(BB-1180 and Subsequent, BL-71 and Subsequent)Electronic flow control units control the massflow of both ambient and bleed air into the

cabin (Figure 12-2). Each unit consists of anambient temperature sensor, an electronic con-troller, and an environmental air control valveassembly, interconnected by a wire harness.

The control valve assembly consists of:

• Mass flow transducer

• Ambient flow motor and modulatingvalve

• Check valve that prevents the bleed airfrom escaping through ambient air intake

• Bleed air flow transducer

• Bleed air flow motor and modulatingvalve (including bypass line)

• Air ejector• Flow control solenoid valve• Environmental shutoff valve

After engine start up when the flow control unitis energized, the bleed air modulating valvecloses. When it is fully closed, it actuates thebleed-air shaft switch, signaling the electroniccontroller to open the solenoid valve. Thisenables P3 bleed air to pressurize the envi-ronmental shutoff valve, causing it to open.

BLEED-AIRFLOW TRANSDUCER

POWER

SQUAT SWITCH

AMBIENTAIR

INLET

AMBIENTFLOW CONTROL

MOTOR

AMBIENTFLOW

TRANSDUCER(MASS FLOW SENSOR)

CHECK VALVEENGINE

BLEED AIR

ELECTRONICCONTROLLER

BLEED-AIRFLOW CONTROLMOTOR

SOLENOID (N.C.)

ENVIRONMENTALSHUTOFFVALVE (N.C.)

AIREJECTOR

BLEED-AIR(HIGH FLOW)BYPASS

TO DUCTDISTRIBUTIONSYSTEM

COCKPIT BLEED AIR VALVE SWITCH

FIRESEAL

AMBIENTTEMPERATURESENSOR

Figure 12-2. Electronic Flow Control Unit (BB-1180 and Subsequent,BL-71 and Subsequent)

The bleed-air shaft continues to open until thedesired bleed-air flow rate to the cabin isreached. (The flow rate is sensed by the bleed-air flow transducer and controlled by the elec-tronic controller per the input of the ambienttemperature sensor.)

As the airplane enters a cooler environment,ambient airflow is gradually reduced andbleed-air flow gradually increased to maintaina constant inflow and to provide sufficientheat for the cabin. At approximately 0°C am-bient temperature, ambient airflow is com-pletely closed off and the bleed-air valvebypass section is opened, as necessary, toallow more bleed air flow past the fixed flowpassage of the air ejector.

Flow Control Unit (Prior to BB-1180, Prior to BL-71)Each flow control unit (Figure 12-3) consistsof an ejector and an integral bleed-air modu-lating valve, firewall shutoff valve, ambientair modulating valve, and a check valve thatprevents the bleed air from escaping throughthe ambient air intake. The flow of bleed airthrough the flow control unit is controlled asa function of atmospheric pressure and tem-perature. Ambient airflow is controlled as afunction of temperature only. When the bleed-air valve switches on the copilot’s left subpanelare in the open position, a bleed-air shutoffelectric solenoid valve on each flow controlunit opens to allow the bleed air into the unit.




AMBIENTFLOW

N.O.AMBIENT AIRMODULATINGVALVE

N.O.SOLENOID VALVE

TOCABIN

AIR TOAIR HEATEXCHANGER

BYPASSVALVE AMBIENT

SENSEANEROID

BYPASSVALVE

N.C.SOLENOID

PRESSUREREGULATOR

EJECTORFLOW

CONTROLACTUATOR

CHECKVALVE

BLEEDAIR FLOW

EJECTOR

TO LH L.G.SAFETYSWITCH

PNEUMOSTAT(PNEUMATIC

THERMOSTAT)

FILTER

N.C.FIREWALLSHUT-OFF

VALVE

TO O

PE

N

TO CLOSETO OPEN

Figure 12-3. Pneumatic Flow Control Unit (Prior to BB-1180, Prior to BL-71)




As the bleed air enters the flow control unit,it passes through a filter before going to thereference pressure regulator. The regulatorwill reduce the pressure to a constant value (18to 20 psi). This reference pressure is then di-rected to the various components within theflow control unit that regulate the output to thecabin.

One reference pressure line is routed to thefirewall shutoff valve located downstream ofthe ejector. A restrictor is placed in the line im-mediately before the shutoff valve to providea controlled opening rate. At the same time,the reference pressure is directed to the am-bient air modulating valve, located upstreamof the ejector, and to the ejector flow controlactuator.

A pneumatic thermostat with a variable ori-fice is connected to the modulating valve. Thepneumatic thermostat (pneumostat) is locatedon the lower aft side of the fireseal forward ofthe firewall. The bi-metallic sensing discs ofthe thermostat are inserted into the cowling in-take. These discs sense ambient temperatureand regulate the size of the thermostat ori-fices. Warm air will open the orifice and coldair will restrict it until, at minus 30°F, the ori-fice will completely close. When the variableorifice is closed, the pressure buildup willcause the modulating valve to close off the am-bient air source. An ambient air shutoff valve,located in the line to the pneumatic thermo-stat is wired to the left landing gear safetyswitch. When the airplane is on the ground, thissolenoid valve is closed, thereby directing thepressure to the modulating valve, causing it toshut off the ambient air source. The exclu-sion of ambient air allows faster cabin warm-up during cold weather operation. An electriccircuit containing a time delay relay is wiredto the above mentioned solenoid valves toallow the left valve to operate several sec-onds before the right valve. This precludesthe simultaneous opening of the modulatingvalves and a sudden pressure surge into thecabin. A check valve, located downstreamfrom the modulation valve, prevents the lossof bleed air through the ambient air intake. Theejector flow control actuator is connected to

another variable orifice of the pneumatic ther-mostat and a variable orifice controlled by anisobaric aneroid. The pneumostat orifice isrestricted by decreasing ambient temperature,and the isobaric aneroid orifice is restricted bydecreasing ambient pressure. The restrictionof either orifice will cause a pressure buildupon the ejector flow control actuator, permit-ting more bleed air to enter the ejector.

OPERATIONThe flow control units regulate the rate of air-flow to the pressure vessel. The bleed air por-tion is variable from approximately 5 to 14pounds per minute depending upon ambienttemperature. On the ground, since ambientair is not available, cabin inflow is variable andlimited by ambient temperature. Inflight, am-bient air provides the balance of the constantairflow volume of 12 to 14 pounds per minute.

From here, the air, which is also being used forcooling and heating, flows into the pressurevessel, creating differential, and out throughthe outflow valve (Figure 12-4) located onthe aft pressure bulkhead. To the left of the out-flow valve (looking forward) is a safety valve(Figure 12-5) which provides pressure reliefif the outflow valve fails, depressurizes the air-plane whenever the CABIN PRESS switch ismoved into the DUMP position, and keeps theairplane unpressurized while it is on the groundwith the left landing-gear safety switch com-pressed. A negative-pressure relief function,which prevents outside atmospheric pressurefrom exceeding cabin pressure by more than0.1 psi during rapid descents with or withoutbleed air flow, is also incorporated into bothvalves.

When the BLEED AIR VALVE switches arein the OPEN position, the air mixture (bleedair and ambient air) from the flow controlunits enters the plane. When the plane is onthe ground, only bleed air enters the cabin be-cause the safety switch causes the flow con-trol units to close a valve that allows ambientair to mix with the bleed air. At liftoff, thesafety valve closes and, except for cold tem-peratures, ambient air begins to enter the flow




SCHRADERVALVE

STATIC AIR

MAXIMUMDIFFERENTIALDIAPHRAGM

SAFETY VALVE DUMP SOLENOID

UPPER DIAPHRAGM

NEGATIVE RELIEFDIAPHRAGM

REARPRESSUREBULKHEAD

CABINAIR

Figure 12-5. Safety Valve

SCHRADERVALVE

PLUG

STATIC AIR

MAXIMUMDIFFERENTIALDIAPHRAGM

TO CONTROLLERCONNECTION

UPPER(CONTROL)DIAPHRAGM

NEGATIVERELIEFDIAPHRAGM

REARPRESSUREBULKHEAD

Figure 12-4. Outflow Valve




control unit, then the pressure vessel. As theleft flow control unit’s ambient air valve opens,in approximately six to eight seconds, theright flow control unit’s ambient air valveopens. By increasing the airflow volume grad-ually (left first, then right), excessive pressurebumps are avoided during takeoff.

An adjustable cabin pressurization controller(Figure 12-6) mounted in the pedestal, com-mands modulation of the outflow valve. Adual-scale indicator dial, mounted in the cen-ter of the controller, indicates the cabin pres-sure altitude on the outer scale (CABIN ALT)and the maximum airplane altitude on theinner scale (ACFT ALT), at which the air-plane can fly without causing the cabin pres-su r e t o exceed max imum d i f f e r en t i a l .Airplanes equipped with the PT6A-41 enginesand maintaining a 6.0 ± 0.1 psi differential canprovide a nominal cabin pressure altitude of10,000 feet at an airplane altitude of 31,300feet. Airplanes equipped with PT6A-42 en-gines and maintaining a 6.5 ± 0.1 psi differ-ential can provide a nominal cabin pressurealtitude of 10,400 feet at an aircraft altitudeof 35,000 feet. The RATE control knob con-trols the rate at which the cabin pressure alti-tude changes from the current value to theselected value. The selected rate of changemay be from approximately 200 to 2,000 feetper minute (fpm).

The actual cabin pressure altitude (outer scale)and cabin differential (inner scale) is contin-uously monitored by the cabin al t imeter(Figure 12-7), located on the right side of thepanel above the throttle quadrant. To the leftof the cabin altimeter is the CABIN CLIMB(cabin vertical speed) indicator (Figure 12-8),which continuously monitors, in feet perminute, the rate of cabin climb and descent.

Figure 12-6. Pressurization Controller

Figure 12-7. Cabin Altimeter

Figure 12-8. CABIN CLIMB Indicator

The CABIN PRESSure switch (Figure 12-9),located to the left of the pressurization con-troller, in the DUMP (forward lever locked)position, opens the safety valve, allowing thecabin to depressurize and stay unpressurizeduntil the switch is placed in the PRESS (cen-ter) position. In the PRESS position, the safetyvalve closes and the pressurization controllertakes command of the outflow valve. In theTEST (aft, spring-loaded to the center) posi-tion, the safety valve is held closed, bypass-ing the landing gear safety switch to allowcabin pressurization tests on the ground.

PREFLIGHT OPERATIONPrior to takeoff, the cabin altitude selector knobis adjusted until the ACFT ALT (inner) scaleon the indicator dial reads an altitude approx-imately 500 feet or 1,000 feet above the plannedcruise pressure altitude. The RATE controlknob is adjusted as desired. When the indexmark is set between the 9 o’clock and 12 o’clockpositions, the most comfortable rate of climbis maintained. The CABIN PRESSure switchis placed in the PRESSure position.

IN-FLIGHT OPERATIONAs the airplane climbs, the cabin pressure alti-tude climbs at the selected rate of change untilthe cabin reaches the selected pressure altitude.

The system then maintains cabin pressure al-titude at the selected value. If the airplaneclimbs to an altitude higher than the value in-dexed on the ACFT ALT scale on the pressurecontroller, the cabin-to-ambient pressure dif-

ferential reaches the pressure relief setting ofthe outflow valve and the safety valve. Eitheror both valves then override the pressure con-troller in order to limit the cabin to ambient pres-sure differential to the normal working pressuredifferential previously stated. If the cabin pres-sure altitude should reach a value of 12,500 feet,a pressure-sensing switch on the forward pres-sure bulkhead closes, thus illuminating the redALT WARN annunciator light, (Figure 12-10),warning the pilot of operation requiring oxy-gen use. If the auto deployment oxygen systemis installed, a pressure-sensing switch in thecabin wall (copilot’s side) forward of the emer-gency exit also closes, deploying the passen-ger oxygen masks to face level. During cruiseoperation, if the flight plan requires an alti-tude change of 1,000 feet or more, the CABINALT dial should be readjusted.

DESCENT AND LANDINGOPERATIONDuring descent and in preparation for land-ing, the cabin altitude selector is set to in-dicate a cabin altitude of approximately 500feet above the landing field pressure alti-tude (Table 12-1). Also, the RATE controlknob is adjusted as required to provide acomfortable cabin altitude rate of descent.The airplane rate of descent is controlled sothe airplane altitude does not catch up withthe cabin pressure altitude until the cabinpressure altitude reaches the selected valueand stabilizes. As the airplane descends toand reaches the cabin pressure altitude, theoutflow valve remains open, keeping thevessel depressurized. As the airplane con-tinues to descend below the preselected cabinpressure altitude, the cabin remains depres-surized and follows the airplane rate of de-scent to touchdown.




Figure 12-9. CABIN PRESSURE Switch

Figure 12-10. ALT WARN Annunciator

LIMITATIONSThe following limitations have been imposedon the pressurization system:

• CABIN DIFFERENTIAL PRESSUREGAGE (B200)

Green Arc (Approved Operating Range)0 to 6.6 psi

Red Arc (Unapp roved Ope ra t i ng Range) 6.6 psi to end of scale

• CABIN DIFFERENTIAL PRESSUREGAGE (200; BB-195 and after)


Red Arc (Unapproved Operating Range)6.1 psi to end of scale

• CABIN DIFFERENTIAL PRESSUREGAGE (200; prior to BB-195)


Red Arc (Unapproved Operating Range)6.0 psi to end of scale

• MAXIMUM OPERATING PRESSURE-ALTITUDE LIMITS

Normal Operation .......... 35,000 feet

• MAXIMUM OPERATING PRESSURE-ALTITUDE LIMITS (Prior to BB-54,except 38, 39, 42, and 44)

Normal Operation .......... 31,000 feet




CLOSEST ADD TOALTIMETER SETTING AIRPORT ELEVATION

28.00 .......................... + 2,400

28.10 .......................... + 2,300

28.20 .......................... + 2,200

28.30 .......................... + 2,100

28.40 .......................... + 2,000

28.50 .......................... + 1,900

28.60 .......................... + 1,800

28.70 .......................... + 1,700

28.80 .......................... + 1,600

28.90 .......................... + 1,500

29.00 .......................... + 1,400

29.10 .......................... + 1,300

29.20 .......................... + 1,200

29.30 .......................... + 1,100

29.40 .......................... + 1,000

29.50 .......................... + 900

29.60 .......................... + 800

29.70 .......................... + 700

29.80 .......................... + 600

29.90 .......................... + 500

30.00 .......................... + 400

30.10 .......................... + 300

30.20 .......................... + 200

30.30 .......................... + 100

30.40 ............................ 0

30.50 ........................... - 100

30.60 ........................... - 200

30.70 ........................... - 300

30.80 ........................... - 400

30.90 ........................... - 500

Table 12-1. PRESSURIZATIONCONTROLLER SETTINGFOR LANDING

14-i

CHAPTER 14LANDING GEAR AND BRAKES

CONTENTS

Page

INTRODUCTION................................................................................................................. 14-1

LANDING GEAR (ELECTRIC) .......................................................................................... 14-2

General........................................................................................................................... 14-2

Gear Assemblies ............................................................................................................ 14-2

Wheel Well Door Mechanisms ...................................................................................... 14-2

Controls.......................................................................................................................... 14-4

Indicators ....................................................................................................................... 14-5

Warning System............................................................................................................. 14-6

Operation ....................................................................................................................... 14-7

LANDING GEAR (HYDRAULIC)...................................................................................... 14-9

General........................................................................................................................... 14-9

Gear Assemblies .......................................................................................................... 14-10

Wheel Well Door Mechanisms.................................................................................... 14-13

Controls ....................................................................................................................... 14-14

Indicators ..................................................................................................................... 14-15

Operation ..................................................................................................................... 14-17

NOSEWHEEL STEERING ................................................................................................ 14-21

General......................................................................................................................... 14-21

Operation ..................................................................................................................... 14-21




BRAKE SYSTEM............................................................................................................... 14-21

Operation ..................................................................................................................... 14-21

Care and Handling in Cold Weather............................................................................ 14-21

Main Gear Safety Switches ......................................................................................... 14-22

LIMITATIONS.................................................................................................................... 14-23

Air Speed Limitations.................................................................................................. 14-23

QUESTIONS....................................................................................................................... 14-25




14-iii

ILLUSTRATIONS

Figure Title Page

14-1 Nose Gear Assembly.............................................................................................. 14-2

14-2 Main Gear Assembly.............................................................................................. 14-3

14-3 Main Gear Door (Standard Gear)........................................................................... 14-4

14-4 Main Gear Door (High Flotation Gear) ................................................................. 14-4

14-5 Landing Gear Switch Handle and Indicator Lights................................................ 14-5

14-6 Normal Indications Gear Down ............................................................................. 14-5

14-7 Nose Gear Not Fully Extended .............................................................................. 14-6

14-8 Normal Indications, Gear Up ................................................................................. 14-6

14-9 One or More Gear Not Fully Retracted.................................................................. 14-6

14-10 Normal Landing Gear Operation ........................................................................... 14-8

14-11 Hydraulic Power Pack............................................................................................ 14-9

14-12 Components Locations......................................................................................... 14-10

14-13 Hydraulic Landing Gear System.......................................................................... 14-11

14-14 Nose Gear Assembly............................................................................................ 14-11

14-15 Internal Nose Gear Lock...................................................................................... 14-12

14-16 Main Gear Assembly ........................................................................................... 14-13

14-17 Main Gear Door Mechanism (Standard Gear).................................................... 14-14

14-18 Main Gear Door Mechanism (High-Flotation Gear) ........................................... 14-14

14-19 Landing Gear Control Handle and Indicator Lights ............................................ 14-15

14-20 Normal Indications Gear Down........................................................................... 14-15

14-21 Nose Gear Not Fully Extended ............................................................................ 14-16

14-22 Normal Indications, Gear Up............................................................................... 14-16

14-23 One or More Gear Not Fully Retracted ............................................................... 14-16




14-24 Normal Retraction................................................................................................ 14-18

14-25 Normal Extension ................................................................................................ 14-19

14-26 Alternate Extension.............................................................................................. 14-20

14-27 Brake System Schematic (Serial Nos. BB-666 and Subsequent) ........................ 14-22

14-28 Brake System Schematic (Serial Nos. BB-453 through BB-665) ....................... 14-23

14-29 Brake System Schematic (Serial Nos. BB-2 through BB-452) ........................... 14-24




INTRODUCTIONThe tricycle landing gear on the Super King Air 200 is actuated either by an electric motoror an electrically-driven hydraulic pump. The gear is controlled with a landing gear con-trol switch handle on the pilot’s right subpanel. On the electrically-actuated gear, motortorque is mechanically transmitted for gear extension and retraction. On the hydraulicgear, three hydraulic actuators provide motive power for gear operation.

Individual gear position lights provide gear position indication and two red lights in thegear control handle. In addition, a warning horn sounds if all three gears are not down andlocked when flap position and/or power lever settings are in the landing configuration.

The hydraulic wheel brake system is pressurized by master cylinders actuated by the pilot’sor copilot’s rudder pedals. Optional bleed-air deicing of the brakes is provided for coldweather operation.

Nosewheel steering is mechanical, actuated by the rudder pedals. Braking and differ-ential thrust can be used to supplement steering.

CHAPTER 14LANDING GEAR AND BRAKES




LANDING GEAR(ELECTRIC)

GENERALThe landing gear is actuated by a 28-VDCmotor powered by the right generator bus.The motor operates torque tubes and a chaindrive to transmit power to a mechanical actu-ator at each gear. A circuit breaker or currentlimiter and a spring-loaded friction clutch,prevents damage to the motor that may resultfrom a mechanical malfunction.

GEAR ASSEMBLIES

DescriptionThe landing gear assemblies (main and nose)consist of shock struts, torque knees (scissors),drag braces, actuators, wheels and tires, brakeassemblies, and a shimmy damper. Brake as-semblies are located on the main gear assem-blies; the shimmy damper is mounted on thenose gear assembly (Figures 14-1 and 14-2).

OperationThe upper end of the drag braces and twopoints on the shock struts are attached to theairplane structure. When the gear is extended,the drag braces are rigid components of thegear assemblies.

Airplane weight is borne by the air charge in theshock struts. At touchdown, the lower portionof each strut is forced into the upper cylinder;this moves fluid through an orifice, further com-pressing the air charge and thus absorbing land-ing shock.

A torque knee connects the upper and lowerportions of the shock struts. It allows strut com-pression and extension but resists rotationalforces, thereby keeping the wheels aligned withthe longitudinal axis of the airplane. On thenose gear assembly, the torque knee also trans-mits steering motion to the nosewheel, andnosewheel shimmy motion to the shimmydamper.

The shimmy damper, mounted on the rightside of the nose gear strut, is a balanced hy-draulic cylinder that bleeds fluid through anorifice to dampen nosewheel shimmy.

The jackscrew actuators retract and extendthe gear and provide a gear uplock due to fric-tion.

WHEEL WELL DOORMECHANISMSLanding gear doors are mechanically actu-ated by gear movement during extension andretraction. On airplanes configured with thestandard main gear, rollers on the shock strutcontact cams in the wheel well during re-traction (Figure 14-3).

Cam movement is transmitted through link-age to close the doors. During extension,roller action reverses cam movement to openthe doors. When the rollers have left thecams, springs drive the linkage overcenter tohold the doors open.

TORQUEKNEE

SHOCKSTRUT

DRAGBRACE

ROLLER(NOSEWHEELDOOR)

PIVOTPOINTSHIMMY

DAMPER

Figure 14-1. Nose Gear Assembly




SHOCKSTRUT

DRAGBRACE

TORQUEKNEE

BRAKEASSEMBLY

PIVOTPOINT

Figure 14-2. Main Gear Assembly




On airplanes configured with the high-flota-tion gear, the main gear wheels are larger andthe shock strut shorter than on the standardgear. Since the wheels will not retract com-pletely into the wheel well, a cutout in thedoors allows part of the wheel to protrude intothe airstream by approximately five inches. Onairplanes so configured, the main gear doorsare mechanically linked to the shock strut andare opened and closed as the gear extends orretracts (Figure 14-4).

Nose gear doors on airplanes with standard orhigh-flotation gear are mechanically actuatedin the manner previously described for stan-dard main gear doors.

CONTROLSThe landing gear is controlled by the LDGGEAR CONT switch handle on the pilot’sright subpanel. Gear position is indicated by

CAM SPRING

Figure 14-3. Main Gear Door (Standard Gear)

DOOR ACTUATING LINK

Figure 14-4. Main Gear Door (HighFlotation Gear)




three green gear position lights adjacent tothe switch handle, and two red lights to illu-minate the handle (Figure 14-5).

The switch handle is detented in both the UPand DN positions. A solenoid-operated down-lock latch (commonly referred to as the “J”hook) engages the handle when the airplaneis on the ground, preventing inadvertent move-ment of the handle to the UP position. Whenairborne, the safety switch on the right maingear completes circuitry to disengage the han-dle latch, and the handle can be positioned toUP. A DN LOCK REL button to the left of thehandle, when pressed, releases the downlocklatch whether the airplane is on the ground orin flight (Figure 14-5).

INDICATORSWhen the gear down cycle begins, the redlights illuminate the switch handle. As eachgear locks down, the corresponding greenlight comes on. When all three gear are downand locked, all three green lights are on andthe handle illumination ceases (Figure 14-6).

If any gear does not lock down during exten-sion, its corresponding green light will not beon and the red handle will remain illuminated(Figure 14-7).

DOWNLOCK REL

WARN HORN

HD LTTEST

LDG GEAR CONTROL

SILENCE

UP

DN

LANDINGGEAR

RELAY

2

RED LIGHT

CAP

DOWN LOCK

RELEASE

SWITCH HANDLE

GREEN POSITION LIGHTS

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

Figure 14-5. Landing Gear Switch Handle and Indicator Lights




WARN HORN

HD LTTEST

LDG GEAR CONTROL

DOWNLOCK REL

SILENCE

UP

DN

LANDINGGEAR

RELAY

5

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

Figure 14-6. Normal Indications Gear Down

When the gear up cycle begins, the handlewill illuminate and the three green positionlights go out. The handle remains illuminateduntil all gear are fully retracted, then goes out(Figure 14-8).

If any gear fails to retract completely, the redlights in the handle remain on (Figure 14-9).

Pushing on the individual light or the lighthousing tests the green position indicator lights.Test the handle illumination lights by pressingthe HDL LT TEST switch (Figure 14-9).

If the DN LOCK REL is overridden and thelanding gear switch handle is placed in theUP position with the airplane on the ground,the handle will illuminate and the warninghorn will sound, provided DC power is avail-able to the airplane.

WARNING SYSTEMThe landing gear warning system consists ofthe red lights that illuminate the LDG GEARCONT switch handle, and a warning horn thatsounds when the gear is not down and lockedduring certain flight regimes.

Super King Air 200, BB-2through BB-452With the flaps in the UP position and either orboth power levers retarded below a certainpower level, the landing gear switch handlewill illuminate. Also, the warning horn willsound intermittently (on Serial Nos. BB-324through BB-452, but only if the airspeed isbelow 140 knots). The horn can be silenced bypressing the WARN HORN SILENCE buttonadjacent to the switch handle; the lights in theswitch handle cannot be cancelled. The land-ing gear warning system will be rearmed if thepower lever(s) are advanced sufficiently.

DOWNLOCK REL

WARN HORN

HD LTTEST

LDG GEAR CONTROL

SILENCE

UP

DN

LANDINGGEAR

RELAY

5

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

Figure 14-8. Normal Indications, Gear Up

WARN HORN

HD LTTEST

LDG GEAR CONTROL

DOWNLOCK REL

SILENCE

UP

DN

LANDINGGEAR

RELAY

5

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

Figure 14-7. Nose Gear Not Fully Extended

DOWNLOCK REL

WARN HORN

HD LTTEST

LDG GEAR CONTROL

SILENCE

UP

DN

LANDINGGEAR

RELAY

5

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

P

R E S S T O T E

DIM

Figure 14-9. One or More Gear Not Fully Retracted




Super King Air 200, BB-453 andSubsequent, and BL-1 andSubsequentSuper King Air B200/B200CWith the flaps in the UP or APPROACH po-sition and either or both power levers retardedbelow a certain power level, the warning hornwill sound intermittently and the switch han-dle lights will illuminate. The horn can be si-l enced by p r e s s ing t he WARN HORNSILENCE button; the lights in the switch han-dle cannot be cancelled. The warning systemwill be rearmed if the power lever(s) are ad-vanced sufficiently.

Super King Air 200, Prior to BB-324With the flaps in the APPROACH positionand either or both power levers retarded belowa certain power level, the warning horn andswitch handle lights will be activated and nei-ther can be cancelled.

Super King Air 200, BB-324 ThroughBB-452With the flaps in the APPROACH position orbeyond, the switch handle lights will illuminateand, if the airspeed is below 140 knots, thewarning horn will sound intermittently. Neitherthe horn nor the lights can be cancelled.

Super King Air 200, Prior to BB-324, BB-453 and Subsequent,and BL-1 and SubsequentSuper King Air B200/B200CWith the flaps beyond the APPROACH posi-tion, the warning horn and the switch handlelights will be activated regardless of the powersettings, and neither can be cancelled.

OPERATION

NormalPull the LDG GEAR CONT switch handle outof detent and position it to UP or DN, as ap-plicable. This applies DC power from the rightgenerator bus to the applicable field windingof the landing gear motor (Figure 14- 10).

As the motor operates, torque tubes and theduplex chain arrangement from the motorgearbox drive the main and nose gear actu-ators to extend or retract the gear. In addi-tion, a 200-amp remote circuit breaker (SerialNos. BB-2 through BB-185) or a 150-amperecurrent limiter (Serial Nos. BB-186 throughBB-1192) protects the motor from overload.

When full extension or retraction has beenachieved, a dynamic brake relay, controlledby up or down limit switches, simultane-ously breaks the motor circuit and completesa circuit through the armature and the unusedfield winding to stop the motor.

Friction in the jackscrew assembly in each ac-tuator holds the gear in the retracted position.

The nose gear is locked down by an over-center condition of the drag brace. The maingears are locked down by a notched hookand plate attachment on the drag braces.

EmergencyReduce speed to 130 knots, place the LDGGEAR CONT switch handle to the DN posi-tion, and pull the LANDING GEAR RELAYcircuit breaker (Figure 14-10).




Figure 14-10. Normal Landing Gear Operation




CROSS-SHAFT

SPRING-LOOADDEDIDLERRS

GEAR BOX

MOTOOTOOROR

A

DETAIL A

LANDING GEARMANUAL EXTENSION SYSTEM

Pull up on the emergency engage handle lo-cated on the floor aft or to the left of thepedestal, and turn it clockwise to the stop.This disconnects electrical power from themotor and engages the emergency drive sys-tem. Pump the extension handle until the threegreen gear position indicator lights come on.Additional pumping could bind the drivemechanism and prevent subsequent retrac-tion; however, if the green indicator lights donot come on, continue pumping until a defi-nite resistance is felt.

After an emergency landing gear ex-tension has been made, do not stowthe extension handle or move anylanding gear controls or reset anyswitches or circuit breakers until theairplane is on jacks. These precau-tions are necessary because the fail-ure may have been in the gear-upcircuit, in which case, the gear mightretract on the ground. The gear can-not be retracted manually.

If a practice emergency extension is made, thegear can be retracted electrically. Rotate theemergency engage handle counterclockwiseand push it down. Stow the extension handleand reset the LANDING GEAR RELAY circuitbreakers. Place the LDG GEAR CONT switchhandle in the UP position to retract the gear.

LANDING GEAR(HYDRAULIC)

GENERALOn airplanes Serial Nos. BB-1193, BL-73,and subsequent, the landing gear is actuatedby a hydraulic power pack (Figure 14-11).The pack consists mainly of a 28-VDC motor-driven hydraulic pump, a hydraulic reser-voir pressurized by engine bleed air, filters,a solenoid-operated selector valve, and an up-lock pressure switch. Adjacent to the pack isa service valve used for hand pump actuationof the gear during ground maintenance op-erations. Figure 14-12 shows the power packand components locations.

WARNING




SYSTEM FILTER

TO NORMAL EXTENDSIDE OF SYSTEM

FROM THE EMERGENCYEXTEND SIDE OF SYSTEM

TO HANDPUMP

FILTER

FROMHANDPUMP

TO RETRACT SIDEOF SYSTEM

SERVICE VALVE

28-VDC PUMPMOTOR

RESERVOIR

FLUID LEVEL SENSORSELECTORVALVE

SOLENOID

UPLOCK PRESSURESWITCH

TO NORMALRETRACT SIDEOF SYSTEM

LEGEND

GEAR EXTEND PRESSURE

GEAR RETRACT PRESSURE

HAND PUMP PRESSURE

SUCTION

RETURN

Figure 14-11. Hydraulic Power Pack

The power pack reservoir, serviced with MIL-H-5606 hydraulic fluid, is divided into two sec-tions. One section supplies the electrically-driven hydraulic pump, and the other sectionsupplies the hand pump. A fill reservoir justinboard of the left nacelle and forward of themain spar (Figure 14-12) features a cap anddipstick assembly for maintaining system fluidlevel.

When reservoir fluid level is low, a sensor onthe reservoir completes a circuit to illuminatean amber HYD FLUID LOW annunciator.Pressing the HYD FLUID SENSOR TESTbut ton on the p i lo t ’s subpanel tes t s theannunciator.

The landing gear is extended and retractedby the power pack in conjunction with threehydrau l i c ac tua to r s , one fo r each gea r(Figure 14-13).

GEAR ASSEMBLIES

DescriptionThe landing gear assemblies (main and nose)consist of shock struts, torque knee (scis-sors), drag braces, actuators, wheels andt i r e s , b r ake a s sembl i e s , and a sh immydamper. Brake assemblies are located on themain gear assemblies; the shimmy damper ismounted on the nose gear assembly (Figures14-14 and 14-15).

OperationThe upper end of the drag braces and twopoints on the shock struts are attached to theairplane structure. When the gear is extended,the drag braces are rigid components of thegear assemblies.




NOSE GEARACTUATOR

SERVICEVALVE

OVERBOARDBLEED AIR

VENT

ACCUMULATOR

MAIN GEARACTUATOR

BLEED AIRREGULATOR

POWER PACKASSEMBLY

MAIN GEARACTUATOR

CHECKVALVE

FILL RESERVOIR

Figure 14-12. Components Locations

Airplane weight is borne by the air charge inthe shock struts. At touchdown, the lower por-tion of each strut is forced into the upper cylin-der; this moves fluid through an orifice, furthercompressing the air charge and thus absorb-ing landing shock. Orifice action also reducesbounce during landing.

At takeoff, the lower portion of the strut ex-tends until an internal stop engages.

A torque knee connects the upper and lowerportion of the shock struts. It allows strut com-pression and extension but resists rotationalforces, thereby keeping the wheels alignedwith the longitudinal axis of the airplane. Onthe nose gear assembly, the torque knee alsotransmits steering motion to the nosewheel,and nosewheel shimmy motion to the shimmydamper.

The shimmy damper, mounted on the rightside of the nose gear strut, is a balanced hy-draulic cylinder that bleeds fluid through anorifice to dampen nosewheel shimmy.




NOSEGEAR

ACTUATOR

HAND PUMPLH MAIN GEAR

ACTUATOR

RH MAIN GEARACTUATOR

HYDRAULICPOWER

PACK

PLUMBING NETWORKFROM POWER PACK

Figure 14-13. Hydraulic Landing Gear System

TORQUEKNEE

SHOCKSTRUT

DRAGBRACE

ROLLER(NOSEWHEELDOOR)

PIVOTPOINTSHIMMY

DAMPER

Figure 14-14. Nose Gear Assembly




PISTON

INLETPORT

ACTUATORDOWNLOCKSWITCH(UNLOCKED)

LOCKSPRING

BALLLOCK

LOCKCOLLAR

BALLLOCK

LOCKCOLLAR

LOCKSPRING

INLETPORT

PISTON

UNLOCKED

LOCKED

ACTUATORDOWNLOCKSWITCH(LOCKED)

Figure 14-15. Internal Nose Gear Lock

A hydraulic actuator attached to the foldingdrag brace of each gear assembly providesmotive force for gear actuation. Nose geardownlocking is provided by an internal lockmechanism (Figure 14-15) in the hydraulicactuator and by the overcenter condition of thedrag brace.

The main gears are mechanically locked downby a notched hook and plate attachment on themain gear drag braces (Figure 14-16).

WHEEL WELL DOORMECHANISMSGear movement during extension and re-traction mechanically actuates landing geardoors. On airplanes configured with the stan-dard main gear, rollers on the shock strutcontact cams in the wheel well during re-traction (Figure 14-17).

Cam movement is transmitted through linkageto close the doors. During extension, roller ac-tion reverses cam movement to open the doors.When the rollers have left the cams, springsdrive the linkage overcenter to hold the doorsopen.

On airplanes configured with the high-flota-tion gear, the main gear wheels are larger andthe shock strut shorter than on the standardgear.Since the wheels will not retract com-pletely into the wheel well, a cutout in thedoors allows part of the wheel to protrude intothe airstream by approximately five inches. Onairplanes so configured, the main gear doorsare mechanically linked to the shock strut andare opened and closed as the gear extends orretracts (Figure 14-18).

Nose gear doors on airplanes with standard orhigh-flotation gear are mechanically actuatedin the manner previously described for stan-dard main gear doors.




Figure 14-16. Main Gear Assembly

CONTROLS The LDG GEAR CONT switch handle on thepilot’s right subpanel controls the landinggear. Gear position is indicated by three greengear position lights adjacent to the switch han-dle, and two red lights to illuminate the han-dle (Figure 14-19).

The switch handle is detented in both the UPand DN positions. A solenoid-operated down-lock latch (commonly referred to as the “J”hook) engages the handle when the airplaneis on the ground, preventing inadvertent move-ment of the handle to the UP position. Whenairborne, the safety switch on the right maingear completes circuitry to disengage the han-dle latch, and the handle can be positioned toUP. A DN LOCK REL button to the left of thehandle, when pressed releases the downlock




CAM SPRING

Figure 14-17. Main Gear Door Mechanism (Standard Gear)

DOOR ACTUATING LINK

Figure 14-18. Main Gear Door Mechanism(High-Flotation Gear)

latch whether the airplane is on the ground orin flight (Figure 14-19). As an additionalsafety factor, control circuitry to the landinggear selector valve is complete only when themain gear safety switches sense an airbornecondition.

INDICATORSLanding gear position is indicated by an as-sembly of three green lights in a single unit tothe right of the LDG GEAR CONT switchhandle. Two red parallel-wired lights in thehandle illuminate to indicate that the gear isunlocked or in transit.

When the gear down cycle begins, the redlights illuminate the switch handle. As eachgear locks down, the corresponding greenlight comes on. When all three gear are downand locked, all three green lights are on, andthe handle illumination ceases (Figure 14-20).

If any gear does not lock down during exten-sion, its corresponding green light will not beon, and the red handle lights will remain on(Figure 14-21).

When the gear up cycle begins, the handle il-luminates and the three green position lightsgo out. The handle remains illuminated untilall gear are fully retracted, then goes out(Figure 14-22).




DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

RED LIGHT

CAP

DOWN LOCK

RELEASE

Figure 14-19. Landing Gear Control Handle and Indicator Lights

DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

Figure 14-20. Normal Indications Gear Down

If any gear fails to retract completely, the han-dle continues to be illuminated (Figure 14-23).

Pushing on the light capsule tests the green po-sition indicator lights. Test the handle illu-mination by pressing the HDL LT TEST switch(Figure 14-23).

Warning SystemThe landing gear warning system consists ofthe red lights that illuminate the LDG GEARCONT switch handle and a warning horn thatsounds when the gear is not down and lockedduring certain flight regimes.

With the flaps in the UP or APPROACH po-sition and either or both power levers retardedbelow approximately 85% N1, the warninghorn will sound intermittently and the switchhandle lights will illuminate. The horn can besilenced by pressing the WARN HORN SI-LENCE button; the lights in the switch han-dle cannot be cancelled. The warning systemwill be rearmed if the power lever(s) are ad-vanced sufficiently. With the flaps beyond theAPPROACH position, the warning horn andthe switch handle lights will be activated re-gardless of the power settings, and neithercan be cancelled.




DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

Figure 14-22. Normal Indications, Gear Up

DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

Figure 14-21. Nose Gear Not Fully Extended

DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

LANDINGGEAR

RELAY

2

OFF

HANDLE LIGHTS TEST SWITCH

Figure 14-23. One or More Gear Not Fully Retracted

OPERATION

Normal RetractionWith the safety switches sensing an airbornecondition, moving the LDG GEAR CONTswitch handle UP completes circuits to thepump motor relay and the up solenoid of thegear selector valve (Figure 14-24).

Power to the pump motor relay pulls in 28VDC to the hydraulic pump motor in the powerpack. The gear selector valve is energized tothe gear up position, directing fluid pressureto the retract side of all three gear actuators.When retraction is complete (approximatelysix seconds), the gear actuators bottom out, andpressure increases rapidly. At 2,775 psi, theuplock pressure switch opens, breaking thecircuit to the pump motor relay, and the pumpmotor deenergizes.

Since there are no gear uplock mechanisms,pressure in the retract side holds the gear re-tracted. When system leakage drops the pres-sure to 2,475 psi, the uplock pressure switchcloses to reestablish the power circuit to thepump. Automatic cycling of the pump main-tains pressure to keep the gear up and locked.

Normal ExtensionPlacing the LDG GEAR CONT switch handlesin the DN position completes a circuit to thedown solenoid of the gear selector valve andthrough any of three gear downlock switchesto the pump motor relay (Figure 14-25). Theenergized relay pulls in 28 VDC for operationof the hydraulic pump motor in the power pack.

The gear selector valve is energized to thedown position, routing pressure to the extendside of all three gear actuators. As each maingear is fully extended, mechanical downlockmechanisms in the drag braces lock the gearin the extended position. A mechanical lockwithin the nose gear actuator locks the nosegear down. As each gear locks down, its down-lock switch is actuated. When the last gearlocks down, the circuit to the pump motorrelay is opened, stopping the pump. The pumpmotor does not cycle after gear extension.

The gear selector valve is spring-loaded tothe down position for fail-safe operation in theevent of electrical power loss.

Alternate ExtensionIn the event of electrical power loss or hy-draulic power pack malfunction, a hydraulichand pump is provided for an alternate gear ex-tension (Figure 14-26). The hand pump, lo-cated on the floor between the pilot’s rightfoot and the pedestal, is labeled LANDINGGEAR EMERGENCY EXTENSION.

To use the alternate extension system, pullthe LANDING GEAR RELAY circuit breakeron the gear control panel, and position theLDG GEAR CONT switch handle DN.

Remove the hand pump handle from the se-curing clip and actuate the hand pump until thethree green gear position lights (NOSE–L–R)illuminate. Place the pump handle in the downposition, and secure in the retaining clip.

If the green gear position lights do notilluminate, continue pumping untilheavy resistance is felt to ensure thegear is down and locked. Then leavethe handle at the top of the stroke.

NOTEThe landing gear cannot be damagedby continued operation of the handpump.

After an alternate gear extension hasbeen made and the pump handleplaced in the securing clip, do notmove any other landing gear con-trols or reset any switches or circuitbreakers until the airplane is on jacksand the cause of the malfunction hasbeen determined and corrected.

WARNING

WARNING




14-18FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

CURRENTLIMITING

RESISTOR

28 VDC 2A

DOWN

UP

SERVICEVALVE

SWITCHES

RIGHT SAFETYSWITCH

TIME DELAY

IN OUTLOGICRELAY

RIGHTMAIN

NOSEACT

LEFTMAIN

DOWNLOCKSWITCHES

LEFT SAFETYSWITCHES

SERVICEVALVE

HAND PUMP

PUMPMOTORRELAY

60A

5A

CHECK VALVE

OVERBOARDVENT

ORIFICE

FILLRESERVOIR

FILLPORT

HANDPUMPSUCTIONPORT

HANDPUMPPRESSUREPORT

SECONDARYRESERVOIR

PRESSURESWITCH

22

HANDPUMPDUMPVALVE

SOLENOID

LEFT MAINACTUATOR

1

1

2

PRESSURECHECKVALVE

1 1

ACCUMULATOR

NOSEACTUATOR

RIGHT MAINACTUATOR

REGULATED ENGNEBLEED AIR (18 TO 20 PSI)

FILTER

FILTER

THERMAL RELIEF VALVE

VENT VALVE

SOLENOIDUP

AUXILIARYPRESSUREPORT(PLUGGED)

GEARDOWNPORT

GEAR UPPORT

PUMP MOTOR

PRIMARYRESERVOIR

RETURNFILTER

VENT PORTAUXILIARY RETURNPORT (PLUGGED) POWER PACK ASSEMBLY

PUMP

PUMPCHECKVALVE

SYSTEMRELIEFVALVE

PRESSURE FLUID

RETURN FLUID

NOTE:

THE INTERNAL SHUTTLE VALVE IS SPRING LOADED TO A POSITION WHICH ALLOWS FLUID IN THE ACTUATOR TO FLOW OUT THE NORMAL EXTENDED PORT.

PRESSURE SWITCH CIRCUIT OPENS ON INCREASING PRESSURE AT 2,275 ± 55 PSIG AND CLOSES ON DECREASING PRESSURE AT A DIFFERENTIAL OF 300–400 PSIG.

LEGEND

CONTROLSWITCH

DOWN

FILTERRELIEFVALVE

28 VDC

ELECTRIC POWER

SELECTOR VALVE

Figure 14-24. Normal Retraction

14-19FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

CURRENTLIMITING

RESISTOR

28 VDC 2A

DOWN

UP

SERVICEVALVE

SWITCHES

RIGHT SAFETYSWITCH

TIME DELAY

IN OUTLOGICRELAY

RIGHTMAIN

NOSEACT

LEFTMAIN

DOWNLOCKSWITCHES

LEFT SAFETYSWITCHES

SERVICEVALVE

HAND PUMP

PUMPMOTORRELAY

60A

5A

CHECK VALVE

OVERBOARDVENT

ORIFICE

FILLRESERVOIR

FILLPORT

HANDPUMPSUCTIONPORT


SECONDARYRESERVOIR

PRESSURESWITCH

2 2

HANDPUMPDUMPVALVE

SOLENOID

LEFT MAINACTUATOR

1

1

2

PRESSURECHECKVALVE

1 1

ACCUMULATOR

NOSEACTUATOR

RIGHT MAINACTUATOR

FILTER

FILTER


VENT VALVE

SOLENOIDUP


GEARDOWNPORT

GEAR UPPORT

PUMP MOTOR

PRIMARYRESERVOIRRETURN

FILTER

VENT PORTAUXILIARY RETURNPORT (PLUGGED) POWER PACK ASSEMBLY

PUMP

PUMPCHECKVALVE

SYSTEMRELIEFVALVE

PRESSURE FLUID

RETURN FLUID

NOTE:

THE INTERNAL SHUTTLE VALVE IS SPRING LOADED TO A POSITION WHICH ALLOWS FLUID FROM GEAR DOWN PORT OF POWER PACK TO FLOW INTO ACTUATOR.

FLUID PRESSURE FROM PUMP UNLOCKS VALVE.

LEGEND

CONTROLSWITCH

DOWN

FILTERRELIEFVALVE

28 VDC

REGULATED ENGNEBLEED AIR (18 TO 20 PSI)

ELECTRIC POWER

SELECTOR VALVE

Figure 14-25. Normal Extension

14-20FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

CURRENTLIMITING

RESISTOR

28 VDC 2A

DOWN

UP

SERVICEVALVE

SWITCHES

RIGHT SAFETYSWITCH

TIME DELAY

IN OUTLOGICRELAY

RIGHTMAIN

NOSEACT

LEFTMAIN

DOWNLOCKSWITCHES

LEFT SAFETYSWITCHES

SELECTOR VALVE

1 1 1

REGULATED ENGINEBLEED AIR (18 TO 20 PSI)

CHECK VALVE

ORIFICE

FILLRESERVOIR

OVERBOARDVENT

28VDC60A

5A

HAND PUMP

HANDPUMPSUCTIONPORT

SERVICEVALVE

LEFT MAINACTUATOR

NOSEACTUATOR

RIGHT MAINACTUATOR

GEAR UPPORT

ACCUMULATOR


VENT VALVE

FILTER

VENT PORTAUXILIARY RETURNPORT (PLUGGED)

POWER PACK ASSEMBLY

PRIMARYRESERVOIR

PUMPSYSTEMRELIEFVALVE


GEARDOWN SUPPORT


PRESSURESWITCH

22

HANDPUMPDUMPVALVE PRESSURE

CHECKVALVE

DOWNSOLENOID

RETURNFILTER

FILLPORT

HAND PUMP PRESSURE FLUID

RETURN FLUID

HAND PUMP SUCTION

CONDITIONS:1. LANDING GEAR CONTROL HANDLE IN "DOWN" POSITION2. 2-AMPERE CONTROL CIRCUIT BREAKER PULLED

NOTES:

PRESSURE FLUID FROM HAND PUMP SHUTTLES INTERNAL SHUTTLE VALVE TO ALLOW FLUID TO FLOW INTO ACTUATOR.

HAND PUMP PRESSURE FLUID UNSEATS VALVE.

UPSOLENOID

SECONDARYRESERVOIR

PUMPMOTORRELAY

PUMPCHECKVALVE

PUMP MOTOR

FILTERRELIEFVALVE

CONTROLSWITCH

LEGEND

1

2

Figure 14-26. Alternate Extension

The landing gear cannot be retracted with thealternate extension system.

After a practice alternate extension, the gearmay be retracted hydraulically by resettingthe LANDING GEAR RELAY circuit breakerand moving the LDG GEAR CONT switchhandle to UP.

NOSEWHEEL STEERING

GENERALDirect linkage from the rudder pedals to an armnear the top of the shock strut mechanicallyactuates nosewheel steering. The steeringangle is from 14° left of center to 12° right ofcenter, but can be considerably increased whenaugmented by differential braking and/or dif-ferential thrust.

OPERATIONSince motion of the rudder pedals is trans-mitted by cables and linkage to the rudder,deflection of the rudder occurs when force isapplied to any of the pedals. With the nose-wheel stationary on the ground or with theself-centering nose gear retracted, rudder pedalmovement compresses a spring-loaded linkin the system but it is not sufficient to steer thenosewheel. If the nosewheel is on the groundand rolling, less force is required for steering;therefore, pedal deflection results in steeringthe nosewheel.

BRAKE SYSTEM

OPERATIONEither the pilot or copilot can apply the brakes.Toe pressure applied to either set of rudderpedals actuates two master cylinders to gen-erate braking pressure (Figures 14-27, 14-28, and 14-29).

Pressure from the master cylinders is appliedto the brake assemblies. Each master cylindersupplies pressure to its set of brake assemblies;therefore, differential braking is available.

Prior to BB-666, the initial pressure from a setof pedals will position a shuttle valve in thebraking system. Brake operation from the op-posite side can then only be accomplished bymoving the shuttle valve.

An optional brake deicing system using bleedair is provided for cold weather operation.This feature is covered in Chapter 10, ICEAND RAIN PROTECTION.

The pilot can set the parking brakes by ap-plying the brakes, then pulling out on thePARKING BRAKE handle on the pilot’s leftor right subpanel. The brakes can be releasedby applying toe pressure on the pedals, thenpushing in the PARKING BRAKE handle.Prior to BB-453, only the pilot can set theparking brake (Figure 14-29).

On some airplanes the PARKING BRAKEhandle is located on the pilot’s right subpanel,below the LDG GEAR CONT switch handle.On these airplanes, either the pilot or the copi-lot can set the parking brakes.

CARE AND HANDLING INCOLD WEATHER

PreflightCheck the brakes and the tire-to-ground con-tact for freeze lockup. Anti-ice solutions maybe used on the brakes and tires if freezeup oc-curs. No anti-ice solution, which contains a lu-bricant, such as oil, should be used on thebrakes. It will decrease the effectiveness of thebrake friction areas.

TaxiingWhen possible, taxiing in deep snow or slushshould be avoided. Under these conditions the




snow and slush can be forced into the brakeassemblies. Keep flaps retracted during taxi-ing to avoid throwing snow and slush into theflap mechanism and to minimize damage toflap surfaces.

Do not taxi with a flat shock strut.

MAIN GEAR SAFETYSWITCHESThe main gear safety switches control somelanding gear functions in addition to func-tions in other systems, as follows.

Left Gear Safety Switch• Safety valve

• Preset solenoid

• Dump solenoid

• Door seal solenoid

• Ambient air modulating valves

• Lift computer (stall warning)

• Stall warning heat control

• Landing gear solenoid (hydraulic gear)

CAUTION




OVERBOARD DRAIN

RESERVOIR

COPILOT’SMASTER

CYLINDER

RIGHT WHEEL BRAKELEFT WHEEL BRAKE

PILOT’SMASTER

CYLINDER

PARKING BRAKE

Figure 14-27. Brake System Schematic (Serial Nos. BB-666 and Subsequent)

Right Gear Safety Switch• Landing gear handle lock solenoid

• Landing gear motor

• Landing gear emergency control

• Flight hourmeter

LIMITATIONS

AIRSPEED LIMITATIONSMaximum Landing Gear Operating Speed

VLO

• Do not extend landing gear above 182KCAS/181 KIAS.

• Do not retract landing gear above 164KCAS/163 KIAS.

Maximum Landing Gear Extended Speed

VLE

• Do not exceed 182 KCAS/181 KIASwith landing gear extended.




OVERBOARD DRAIN

RESERVOIR

LEFT WHEEL BRAKE

PILOT’SMASTER

CYLINDERPARKING BRAKE

SHUTTLEVALVE

SHUTTLEVALVE

PARKVALVE

PARKVALVE

RIGHT WHEEL BRAKE

COPILOT’SMASTERCYLINDER

Figure 14-28. Brake System Schematic (Serial Nos. BB-453 through BB-665)




Figure 14-29. Brake System Schematic (Serial Nos. BB-2 through BB-452)

15-i

CHAPTER 15FLIGHT CONTROLS

CONTENTS

Page

INTRODUCTION................................................................................................................. 15-1

GENERAL ............................................................................................................................ 15-1

FLIGHT CONTROL LOCKS............................................................................................... 15-2

ROLL..................................................................................................................................... 15-3

Operation ....................................................................................................................... 15-3

PITCH.................................................................................................................................... 15-3

Operation ....................................................................................................................... 15-3

YAW ...................................................................................................................................... 15-3

Operation ....................................................................................................................... 15-3

Rudder Boost ................................................................................................................. 15-3

Yaw Dampening............................................................................................................. 15-5

TRIM SYSTEMS .................................................................................................................. 15-5

Operation ....................................................................................................................... 15-5

Elevator Electric Trim.................................................................................................... 15-6

FLAPS ................................................................................................................................... 15-6

Operation ....................................................................................................................... 15-8

Split Flap Protection .................................................................................................... 15-10

STALL WARNING............................................................................................................. 15-10

Operation ..................................................................................................................... 15-10




LIMITATIONS.................................................................................................................... 15-12

Airspeed Limitations ................................................................................................... 15-12

Maneuver Limits.......................................................................................................... 15-12

Flight Load Factor Limits at 12,500 Pounds ............................................................... 15-12

Maximum Operating Pressure-Altitude Limits ........................................................... 15-12

QUESTIONS....................................................................................................................... 15-13




15-iii

ILLUSTRATIONS

Figure Title Page

15-1 Flight Controls and Trim Tabs ............................................................................... 15-2

15-2 Flight Control Locks .............................................................................................. 15-3

15-3 Rudder Boost and Yaw Damp Switches ................................................................ 15-4

15-4 Rudder Boost Diagram........................................................................................... 15-5

15-5 Autopilot and Yaw Damp Switches ....................................................................... 15-5

15-6 Trim System Control .............................................................................................. 15-6

15-7 Elevator Electric Trim Controls ............................................................................. 15-7

15-8 Flap System Diagram............................................................................................. 15-8

15-9 Flap Control and Indication ................................................................................... 15-9

15-10 Stall Warning System........................................................................................... 15-11




INTRODUCTIONThe Super King Air is equipped with manually-actuated primary flight controls, oper-ated through cables, bellcranks, and pushrods. The ailerons and rudder are conven-tional; the horizontal stabilizer and elevators are mounted at the extreme top of the verticalstabilizer, conforming to the T-tail configuration. A pneumatic rudder boost system as-sists in directional control in the event of engine failure or a difference in engine bleedair pressure.

All surfaces are manually trimmed from the cockpit; however, optional elevator elec-tric trim is available. Two trailing-edge flaps on each wing are actuated by an electricmotor driving flexible drive shafts through a gearbox. A safety mechanism provides splitflap protection. A stall warning system provides aural warning of an imminent stall.

GENERALThe flight controls consist of ailerons, ele-vators, rudder, and flaps. Excluding flapsand the right aileron, all surfaces incorporate

trim tabs (the right aileron has a ground ad-justable trim tab) (Figure 15-1).

20

20 20

105

510 10

5

5

LO

C

GS

CHAPTER 15FLIGHT CONTROLS




FLIGHT CONTROLLOCKSThe flight control locks consist of a chain, twopins, and a U-shaped clamp (Figure 15-2).

The pin in the control column prevents con-trol wheel rotation and fore-and-aft move-ment of the control column, locking theailerons and elevators. On BB-82 and subse-quent, BL-1 and subsequent, and all B200 air-planes, the control column must be full forwardand a control wheel rotated 15° left beforethe pin can be installed.

The L-shaped pin inserted through the hole inthe floor aft of the pilot’s rudder pedals locks

the rudder. The pedals must be centered beforethis pin can be installed.

The U-shaped clamp around the power leversserves as a warning not to start engines withthe control locks installed.

NOTEThe rudder control lock must be re-moved prior to towing the airplaneto prevent damage to the steeringlinkage.

External flight control surface locks areoptional.




ELEVATORS

TRIM TABS

RUDDER

TRIM TAB

AILERON

TRIM TAB

FLAPS

FLAPS

GROUND ADJUSTABLE TAB

AILERON

Figure 15-1. Flight Controls and Trim Tabs




ROLL

OPERATIONRoll control around the longitudinal axis ismaintained by conventional ailerons mountedon the trailing edge of each wing, outboard ofthe flaps. Rotation of either interconnectedcontrol wheel on the control column mech-anically positions the ailerons. Aileron travelis approximately 25° up and 17° down, lim-ited by adjustable stops.

PITCH

OPERATIONPitch control around the lateral axis is providedby elevators attached to the aft edge of the hor-izontal stabilizer. Since the control columnsare linked together, fore-or-aft movement ofeither column transmits motion through cables,

bellcranks, and pushrod linkage to move theelevators. Elevator travel is approximately20° up and 14° down, and is limited by ad-justable stops.

YAW

OPERATIONYaw control around the vertical axis is main-tained by the rudder, which extends along theentire aft edge of the vertical stabilizer. It isactuated, through cables and bellcranks, by ei-ther set of mechanically-connected rudderpedals. Rudder travel is approximately 15°either side of neutral, and is limited by ad-justable stops. Yaw damping and rudder boostare also activated through the rudder.

RUDDER BOOSTA rudder boost system is provided as an aidin maintaining directional control in the eventof engine failure or a large variation of powerbetween the engines. Two pneumatic boostservos are incorporated into the rudder cablesystem to provide force for rudder boosting,when required.

OperationThe rudder boost system is armed by placingthe RUDDER BOOST switch to the ON posi-tion, and both the left and right BLEED AIRVALVE switches in either the OPEN or ENVIROFF positions (Figure 15-3).

A differential pressure switch in the system( c o m m o n l y r e f e r r e d t o a s t h e D e l t a Pswitch) senses bleed-air pressure from eachengine. If a substantial pressure differen-tial exists (60 ± 5 psi), a circuit is completedto open a solenoid operated valve that di-rects regulated bleed-air pressure to theapplicable rudder boost servo, boosting therudder to compensate for asymmetr ica lthrust (Figure 15-4). Placing either of theBLEED AIR VALVE switches to the INSTR& ENVIR OFF position will cause the sys-tem to disengage.

Figure 15-2. Flight Control Locks




WARNING DEPRESSURIZE CABINBEFORE LANDING

CABINALT

1000FT

CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

PRESSTEST OFF OFF

YAWENG

APENG

SR

I/20

DN

L

UP

R

YAW ALT

YAW

HDG

COLLINS

HDG NAV APPR B/C CLIMB

ALT ALT SEL VS IAS DSC

INSTR & ENVIR OFF

ENVROFF

OPENLEFT RIGHT

BLEED AIR VALVES

DN

LIFT

LIFTIDLE

GDFINE

PITCH

TRI

M

UP

PROP

ITION

LOWIDLE

CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

PRESS

COCKPIT VOICE RECORDER

HEADSET

TEST ERASE

600 OHMS

Collins

POWER

TEST OFF OFF

PARKING BRAKEOFF

ENGINE ANTI-ICE

ON

MAIN

OFF

ACTUATORSTANDBY

LEFT RIGHT

COLLINS


OFF

LIGHTS

LEFT RIGHT

IN

PULL

Figure 15-3. Rudder Boost and Yaw Damp Switches




The system is tested during engine runup byretarding one engine to idle and advancingpower on the other engine until the rudderpedal on the same side as the high rpm enginemoves forward. Reverse the procedure to checkthe opposite side.

YAW DAMPING

OperationOn all airplanes, a yaw damping system isprovided. It can be activated with a switchlocated on the pedestal or autopilot panel(Figure 15-3 and Figure 15-5). On some in-stallations it will automatically activate withautopilot engagement.

The system is required to be operational above17,000 feet.

TRIM SYSTEMS

OPERATIONTrim in all three axes is maintained by trim tabson the primary flight control surfaces. A tab

18 PSIPRESSURE

REGULATOR

FILTER15 PSI

PRESSUREREGULATOR

N.C. N.C.


ENVIROFF

INSTR & ENVIR OFF

RUDDERBOOST

OFF

DUAL FED NO. 2VDC

VDC

VDC

LEFT CONTROL(DUAL FED NO. 1)

RIGHT CONTROL(DUAL FED NO. 2)

PNEUMATICPRESSURE

18 PSI

TO RIGHT BLEED AIRWARNING SYSTEM

TO LEFT BLEED AIRWARNING SYTEM

CHECK VALVECHECK VALVE

60 PSID SWITCH

LEFTRUDDERSERVO

RIGHTRUDDER

SERVO

PNEUMATICSHUTOFF VALVE

PNEUMATICSHUTOFF VALVE

Figure 15-4. Rudder Boost Diagram

YAWENG

APENG

SR

I/20

DN

L

UP

R

YAW ALT

YAW

HDG

COLLINS



Figure 15-5. Autopilot and Yaw DampSwitches

is located on the trailing edge of the rudder,each elevator, and the left aileron. Movingtrim wheels (Figure 15-6) mechanically trans-mits motion to screwjack actuators that posi-tion the tabs.

ELEVATOR ELECTRIC TRIMMost airplanes have an optional electric ele-vator trim system installed. An electric motorin the fuselage aft section actuates the eleva-tor trim tabs through a system of cables.

OperationThe ELEV TRIM switch must be placed in theON position to arm the system (Figure 15-7).

Electr ical power to the system is routedthrough the PITCH TRIM circuit breaker. DualPITCH TRIM thumb switches on the outboardside of either control wheel must be moved si-multaneously to achieve pitch trim. Eitherswitch alone will not actuate the trim motor.As an option in some airplanes, trim inputs bythe pilot override those made by the copilot.The PITCH TRIM switches are spring loaded

to the center (off) position when released. Themanual elevator trim wheel can be used fortrimming, even when the electrical trim sys-tem is switched on.

A bilevel, push-button, momentary-on trimdisconnect switch on each control wheel canbe used to disconnect the system. To initiatea trim disconnect, depress either of theseswitches to the second level. The green ELECTRIM OFF light on the advisory panel comeson when disconnect is selected. To reset thesystem for subsequent operation, cycle theELEV TAB CONTROL switch to OFF, thenback to ON.

FLAPSTwo flaps on each wing are driven by an elec-tric motor through a gearbox and four flexi-ble drive shafts connected to screwjacks ateach flap. The motor incorporates a dynamicbraking system through the use of two sets offield windings. Lowering the flaps results innose pitch-up, lowered stall speed, and re-




WARNING DEPRESSURIZE CABINBEFORE LANDING

CABINALT

1000FT

CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

PRESSTEST OFF OFF

DN

LIFT

LIFTIDLE

GDFINE

PITCH

TRIM

UP

PROP

ITION

LOWIDLE

CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

PRESS

COCKPIT VOICE RECORDER

HEADSET

TEST ERASE

600 OHMS

Collins

POWER

TEST OFF OFF

ELEVATORTRIM WHEEL

AILERONTRIM WHEEL

RUDDERTRIM WHEEL

ELEVATOR TABCONTROL SWITCH

Figure 15-6. Trim System Control




DEICE ON LDG/TAX I LIGHT

ELEC TRIM O FF

BATTERY C HARGE

PASS O

AIR COND

EXTFLIGHT

5

ALT

ALERT

5

COPLT

5

TURN &SLIP

3

PILOTPILOT

ALTMAIR DATA

1

COPLT

ENCD ALTM

5

PITCH

TRIM

5

RUDDER

BOOST

5

OUTSIDE

AIRTEMP

PITCH TRIM CIRCUIT BREAKER ELECT TRIM OFF LIGHT

CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

PRESSTEST OFF OFF

ELEVATOR TRIM SWITCHNOSE DN

NOSE UP

PI

TC

H TR

IM

TRIMDISCONNECTSWITCH

Figure 15-7. Elevator Electric Trim Controls

duced airspeed. The flaps limit switches andflaps position transmitter are located under theright inboard flap.

OPERATIONFlap movement is initiated by positioning theFLAP handle (Figure 15-8).

Placing the FLAP handle from the UP to theAPPROACH (40%) position connects No. 3d u a l - f e d b u s p o w e r t h r o u g h t h e F L A PMOTOR circuit breaker to the flap motor(Figure 15-9). The flaps are driven to the40% (14° ± 1°) position, as indicated on theflap position indicator. For BB-1439, 1444and subsequent, the flaps cannot be stoppedat any intermediate point during this travel.

Placing the handle to the DOWN position andleaving it there results in full 100%, (35° +1°,–2°) flap extension. For BB-1439, 1444 andsubsequent only the UP, APPROACH (or take-off), and DN positions are selectable. However,they are follow-up flaps which allows the flapsto extend or retract to achieve the selectedflap handle position. The flaps cannot bestopped in any intermediate position.

Prior to BB-1439, 1444 and subsequent, if anyposition between 40% and 100% is desired (forexample, 60%), place the handle to DOWNuntil the desired position is attained, then re-turn it to the APPROACH position. The flapswill stop at 60%. In like manner, the flaps maybe raised to any position between DOWN andAPPROACH by placing the handle in the UPposition beyond the APPROACH detent until




POSITIONTRANSMITTER

FLAPDYNAMICBRAKERELAY

UP

APP

DOWN

STALLWARNING

BIASRELAYS

FLAPMOTOR

LIFTCOMPUTER

SPLITFLAPPROTECTION

LH

RH

FUSES ORCAM SWITCHES

FLAP CONTROLC/B

FLAP MOTORC/B

DUAL FEDBUS NO. 3

UP

APPROACH

DOWN

FLAP

TAKEOFFAND

APPROACH

DOWN

UP

FLAPS20

60

80

LIMIT SWITCHES

POSITIONINDICATOR

Figure 15-8. Flap System Diagram




FLAPS

DOWN

TAKEOFFAND

APPROACH

UP20

60

80

OPEN

CLOSED

FIREWALLSHUTOFF

VALVE

LEFT

RIGHT

FUEL SYSTEM CLOSED

ITT PROPTACH

TURBINETACH

FUELFLOW

OILPRESS

OILTEMP

5 5 5 5 5 5 5

5 5 5 5 5 5 5

TORQUE

FIREWALLVALVE

AUXTRANS

FER

AUXTRANS

FER

QTYIND

PRESSWARN

5 10 5 5 5 5

STANDBYPUMP

STANDBYPUMP

CROSSFEED

QTYIND

FIREWALLVALVE

5 5 5 10 5

PRESSWARN

RIGHTLEFT

OPEN

ENGINE INSTRUMENTS

FIREWALLSHUTOFF

VALVE

NO 4

BUSFEEDERS

BUSFEEDERS

NO 3

50

50

RIGHT

PROPDEICE

LEFT

25

25

CONTROL

MOTORPROPDEICE

FLAP

20

5

GOV

CONTROL

PROP

5

5

RIGHT

IGNITORPOWER

LEFT

5

5

RIGHT

STARTCONTROL

LEFT

5

5

NO 4

NO 3

50

50

OPEN

CLOSED


FUEL SYSTEM

CLOSED

GOV

5

5 10 5 5 5 5 5 5 5 10 5

FIREWALLVALVE

AUXTRANS

FER

AUXTRANS

FER

QTYIND

PRESSWARN

STANDBYPUMP

STANDBYPUMP

CROSSFEED

QTYIND

FIREWALLVALVE

PRESSWARN

OPEN


NO 3

BUS FEEDERS

50 50 50 50

RIGHTLEFT

NO 4

LEFT

5

RIGHT LEFT RIGHT LEFT RIGHT

25 25

CONTROL PROP PROP

PROP DEICE FLAP

20 5

MOTOR CONTROL

5 5

POWER

5 5

CONTROLPROP IGNITOR START

BB-1439, BB-1444 THROUGH BB-1485, EXCEPT BB-1463 AND BB-1484; BL-139 AND BL-140

BB-1484, 1486 AND AFTERFLAPCIRCUT

BREAKERS

Figure 15-9. Flap Control and Indication

the desired position is reached, then returningit to the APPROACH detent. Moving the han-dle from DOWN to APPROACH will not retractthe flaps. When the handle is moved from theAPPROACH position to the UP position, theflaps retract completely and cannot be stoppedin between the APPROACH and UP positions.

A safety mechanism is provided to discon-nect power to the electric flap motor in theevent of a malfunction, which would causeany flap to be 3° to 6° out of phase with theother flaps.

The flap-motor power circuit is protected by a20-ampere flap-motor circuit breakerplacarded FLAP MOTOR, located on the leftcircuit-breaker panel below the fuel controlpanel. A 5-ampere circuit breaker for thecontrol circuit (placarded FLAP CONTROL) isalso located on this panel.

Super King Air 200, BB-453 andSubsequent, BL-1 and Subsequent andSuper King Air B200/B200C

With the flaps in the UP or the APPROACHposition and either or both power levers re-tarded below a certain power level, the warn-ing horn will sound intermittently and thelanding gear switch handle lights will illumi-nate. The horn can be silenced by pressingthe WARN HORN SILENCE button; the lightsin the switch handle cannot be cancelled. Thelanding gear warning system will be rearmedif the power lever(s) are advanced sufficiently.

SPLIT FLAP PROTECTIONA split flap sensing system (Figure 15-8) pro-vides protection in the event any flap panel isout of phase with the other panel. Airplane se-r i a l s BB-425 and subsequen t u t i l i z e acam/switch arrangement. BB-2 through BB-424 are equipped with a fuse block protec-tion system.

The fuse or switch is rigged in such a waythat if either flap on that side splits 3° to 6°during travel up or down, the circuit is inter-rupted and the motor stops. Once the motorstops due to a split flap condition, the flaps can-not be moved until the failure is corrected.

Protection is provided between each pair offlaps on that side of the airplane. There is nosplit flap protection between the left pair offlaps from the right.

STALL WARNINGOPERATIONThe stall warning system senses angle of at-tack through a lift transducer actuated by avane mounted on the leading edge of the leftwing (Figure 15-10).

Angle of attack from the lift transducer andflap position signals are processed by the liftcomputer to sound the stall warning hornmounted on the copilot’s side of the cockpit.The horn sounds when the following condi-tions are present:

1. Airspeed is 5 to 13 knots above stall,flaps are fully retracted.

2. Airspeed is 5 to 12 knots above stall,flaps are in the APPROACH (40%)position.

3. Airspeed is 8 to 14 knots above stall,flaps are fully extended.

The system can be tested prior to flight byplacing the STALL WARN TEST switch, lo-cated on the copilot’s left subpanel, in theTEST position. This simulates a stall condi-tion and sounds the warning horn.

The heating elements protect thelift transducer vane and faceplatefrom ice. However, a buildup of iceon the wing may change or disruptthe airflow and prevent the systemfrom accurately indicating an im-minent stall . Remember that thestall speed increases whenever iceaccumulates on any airplane.

WARNING







LIFT TRANSDUCER VANE COPILOT'S LEFT SUBPANEL

Figure 15-10. Stall Warning System

LIMITATIONSFor complete limitations information, referto the LIMITATIONS section of the Pilot’sOperating Manual.

AIRSPEED LIMITATIONSManeuvering Speed

VA (12,500 pounds)

• Do not make full or abrupt control move-ments above 182 KCAS/181 KIAS.

Maximum Flap Extension/Extended Speed

VFEAPPROACH Position—40%

• Do not extend flaps or operate with 40%flaps above 200 KCAS/KIAS.

Full DOWN Position—100%

• Do not extend flaps or operate with100% flaps above 144 KCAS/146 KIAS(King Air 200) or 155 KCAS/157 KIAS(King Air B200).

Maximum Landing Gear Operating Speed

VLO• Do not extend landing gear above 182

KCAS/181 KIAS.

• Do not retract landing gear above 164KCAS/163 KIAS.

Maximum Landing Gear Extended Speed

VLE

• Do not exceed 182 KCAS/181 KIASwith landing gear extended.

Air Minimum Control Speed

VMCA

• The lowest airspeed at which the air-plane is directionally controllable whenone engine suddenly becomes inopera-tive and the other engine is at takeoffpower is 91 KCAS/86 KIAS.

Maximum Operating Speed

VM0

MM0

• Do not exceed 260 KCAS/259 KIAS(.52 Mach) in any operation.

NOTESuper King Air B200/B200C, SuperKing Air 200 Serial No. BB-199 andsubsequent, BL-1 and subsequent,and any earlier airplanes modifiedby Beechcraft Kit Number 101-5033-1 in compliance with BeechcraftService Instruction Number 0894.

Maximum Operating SpeedVM0

MM0

• Do not exceed 270 KCAS/269 KIAS(.48 Mach) in any operation.

NOTESuper King Air 200 Serial No. BB-2 through BB-198, except airplanesmodified by Beechcraft Kit Number101-5033-1 in compl iance wi thBeechc ra f t Se rv ice Ins t ruc t ionNumber 0894.

MANEUVER LIMITSThe Super King Air 200 and B200 are NormalCategory Aircraft. Acrobatic maneuvers, in-cluding spins, are prohibited.

FLIGHT LOAD FACTOR LIMITSAT 12,500 POUNDSFlaps Up

• Do not exceed 3.17 positive Gs, or 1.27negative Gs.

Flaps Down

• Do not exceed 2.00 positive Gs, or 0negative Gs (B200); (1.27 negative for200).

MAXIMUM OPERATINGPRESSURE-ALTITUDE LIMITSDo not exceed 17,000 feet with yaw damperinoperative.




16-i

CHAPTER 16AVIONICS

CONTENTS

Page

INTRODUCTION................................................................................................................. 16-1

COLLINS PROLINE II......................................................................................................... 16-1

Audio System................................................................................................................. 16-1

Communication Radios.................................................................................................. 16-3

ADF Equipment........................................................................................................... 16-13

Transponder Equipment............................................................................................... 16-17

FLIGHT INSTRUMENTS.................................................................................................. 16-20

Pitot and Static System................................................................................................ 16-20

Outside Air Temperature Gage.................................................................................... 16-23

AUTOFLIGHT SYSTEM................................................................................................... 16-24

Yaw Damper ................................................................................................................ 16-24

STALL WARNING SYSTEM ............................................................................................ 16-25

COMMUNICATION SYSTEM.......................................................................................... 16-26

Static Discharging Description .................................................................................... 16-26

LIMITATIONS.................................................................................................................... 16-27

Airspeed Indicator ....................................................................................................... 16-27

Outside Air Temperature Gage.................................................................................... 16-27

Autopilot ...................................................................................................................... 16-27

QUESTIONS....................................................................................................................... 16-28




16-iii

ILLUSTRATIONS

Figure Title Page

16-1 B200 Audio Panel Controls ................................................................................... 16-2

16-2 VHF-22 COMM Radio Controls/Displays ............................................................ 16-5

16-3 VHF-22 COMM Radio Self Test Displays ............................................................ 16-6

16-4 VIR-32 NAV Receiver Controls/Displays ............................................................. 16-8

16-5 DME-42 Systems IND-42 Displays....................................................................... 16-9

16-6 Typical Pro Line II Dual DME Installation.......................................................... 16-11

16-7 IND-42 Self Test Displays ................................................................................... 16-13

16-8 VIR-32 NAV Receiver Self-Test Displays........................................................... 16-13

16-9 ADF-60A ADF Receiver Controls/Displays ....................................................... 16-14

16-10 ADF-60A ADF Receiver Self-Test Displays....................................................... 16-17

16-11 TDR-94 Transponder Controls/Displays ............................................................. 16-17

16-12 TDR-94 Transponder Self-Test Displays............................................................. 16-20

16-13 B200 Antenna Locations...................................................................................... 16-21

16-14 Pitot and Static System Diagram ......................................................................... 16-22

16-15 Pitot Mast Location.............................................................................................. 16-23

16-16 Static Ports Location ............................................................................................ 16-23

16-17 Pilot’s Static Air Source Valve Switch................................................................. 16-23

16-18 Typical OAT Gage and Probe .............................................................................. 16-24

16-19 YAW Damp Switch.............................................................................................. 16-25

16-20 Stall Warning Transducer Vane............................................................................ 16-25

16-21 STALL WARN TEST Switch (Copilot’s Left Subpanel) .................................... 16-25

16-22 STALL WARN Heat Switch (Pilot’s Right Subpanel) ........................................ 16-26

16-23 Static Wicks ......................................................................................................... 16-26




16-v

TABLES

Table Title Page

16-1 CTL-22 COMM Control, Controls and Indications ................................................ 16-4

16-2 CTL-32 NAV/DME Control, Controls and Indications .......................................... 16-7

16-3 IND-42A/C DME Indicator, Controls and Indications ......................................... 16-10

16-4 CTL-62 ADF Control, Controls and Indications .................................................. 16-15

16-5 CTL-92 ATC Control, Controls and Indications................................................... 16-18

16-6 Airspeed Indicator Markings................................................................................. 16-27




INTRODUCTIONThe Super King Air utilizes an avionics package which consists of, but is not limitedto, the navigation system, the weather radar system, the autoflight system, the stallwarning system, and the communication system.

COLLINS PROLINE II

AUDIO SYSTEM

GeneralThe audio system consists of an audio controlpanel, two flight compartment speakers withjacks for pilot and copilot headphones andmicrophones, dual audio amplifiers, a pas-senger speaker amplifier, and an aural warn-ing tone generator.

The audio control panel provides control overboth transmission and reception of all com-munication and navigation equipment installedin the airplane. ON–OFF switches, source se-lector switches and volume controls are pro-vided for pilot and copilot control of eachindividual audio system (Figure 16-1).

CHAPTER 16AVIONICS




16-2FO

R T

RA

ININ

G P

UR

PO

SES

ON

LY

SU

PER

KIN

G A

IR 2

00

/B

20

0 P

ILOT TR

AIN

ING

MAN

UAL

Flig

htS

afetyin

ternatio

nal

VOL VOL VOL

VOL VOL

AUTOCOMM

COMM

OFF

AUTOCOMM

OFFPILOT AUDIO OFF

NAV

AUDIOSPKR

SIDE-TONE

INTPHSENS

OFF

AUDIOSPKR

OFFNORM

HI

LO

DME

RANGE

VOICE PAGING INTPHAUDIOEMER

OFF

HOTINTPH

MKR BCN DME1 12 2

1 2

2 2 ADF1 1COMM

COPILOT AUDIO OFF

NAV DME1 12 2 2 2 ADF1 1

BOTH RANGE

VOICE BOTH

MKR BCN1 & 2

VO L

COMM 1

COMM 2

CABIN

VO L

COMM 1

COMM 2

CABIN

PUSHON/OFF

GNDCOMMPWR

ANNPUSH BRT

SIDE-TONE

INTPHSENS

ENCDALTM

1

2

AVIONICS BY BEECHCRAFT

MKR BCN

DIM

COPILOTSMICROPHONESELECTORSWITCH

PILOTSMICROPHONE

SELECTORSWITCH

COPILOT AUDIO OUTPUTSOURCE SELECT SWITCHES

PILOT AUDIO OUTPUTSOURCE SELECT SWITCHES

PASSENGERADDRESSVOLUME

CONTROL

COPILOT NAV/ADFVOICE/MORSEFILTER SWITCH

PILOT NAV/ADFVOICE/MORSEFILTER SWITCH

COPILOTSIDETONE

VOLUME POTPILOT

SIDETONEVOLUME POT

COPILOTINTERPHONE

THRESHOLD POT

PILOTINTERPHONE

THRESHOLD POT

COPILOTSPEAKERSWITCH

PILOTSPEAKER

SWITCH

*COPILOT AUTOCOMM SWITCH

*PILOT AUTOCOMM SWITCH

*AUTOMATICALLY TURNSON AUDIO OUTPUT FROMSELECTED TRANSMITTER

*AUTOMATICALLY TURNSON AUDIO OUTPUT FROMSELECTED TRANSMITTER

INTERPHONEVOLUME

CONTROL

COPILOTMASTERVOLUMECONTROL

PILOTMASTERVOLUME

CONTROL

MARKER BEACONVOLUME CONTROL

MARKER BEACONSENSE SWITCH

HOT INTERPHONEON/OFF SWITCH

DME VOLUMECONTROLS

*AUDIO AMPSBYPASS SWITCH

*WHEN IN EMER POSITION THE FOLLOWING AUDIO ONLY BYPASSES THE AMP AND ALL ARE MIXED TOGETHER:COMM 1, COMM,2 SIDETONE 1, SIDETONE 2, AND AURAL WARNING INPUTS

GROUND COMMPOWER SWITCH

Figure 16-1. B200 Audio Panel Controls




OperationThe two audio control amplifiers operate in-dependently for the pilot and copilot systems.If the pilot moves a source select switch up,the pilot’s headset will receive audio fromthat source. If the “Audio Speaker” switch isselected up, then the audio will also be broad-cast over the pilot’s speaker. If this audio con-tains both a Morse code (Range) and a voicemessage, these can be listened to simultane-ously or individually with the switch labeledVOICE–BOTH–RANGE. All of the above se-lections are identical on the copilot’s side.

A microphone selector switch allows trans-mission through either Communication RadioNo. 1 (COMM 1), Radio No. 2 (COMM 2), orto the “Cabin” speakers. For the COMM 1 andCOMM 2 selection an AUTO COMM switchwill automatically send the selected COMMradio’s received audio to the headset (andspeaker, if selected). This eliminates havingto select the COMM radio on the source selectpanel unless audio from the opposite radio isdesired (e.g., microphone selector switch onCOMM 1 for enroute communication but ATISon COMM 2).

Each COMM radio has an individual volumecontrol knob. However, a knob is provided oneach microphone selector switch for mastervolume control (this does not affect incomingwarning tones).

A GND COMM PWR button is connected tothe hot battery bus of the aircraft and allowsactivation of headsets, speakers and handheldmicrophones and COMM 1 prior to turning onthe main aircraft battery switch. This allowsfrequencies such as clearance delivery andATIS to be retrieved without excessive batteryuse. Once communication is complete, turn offthe system by pushing the button again.(Although the ground communication powerwill disconnect when turning the battery switchon, normal procedure is to turn it off throughthe GND COMM PWR button.)

The passenger speaker amplifier provides chimetones in conjunction with the fasten-seatbeltand no smoking selection and paging audio

through the use of the microphone selectorswitch mentioned above. A knob labeled PAG-ING allows volume control of cabin audio.

The aural warning tone generator generatesaural warning tones to the cockpit headphonesand speaker for stall warning, landing gearwarning, autopilot disconnect warning, and al-titude alert warning (other optional equip-ment may be included). The aural warningtone generator operates on command fromfault detection equipment and does not havea volume control.

An AUDIO EMER switch is provided shouldthere be a failure of both amplifier systems(Figure 16-1). When selected to EMER, theamplifier selector panel is no longer func-tional. COMM 1, COMM 2, SIDETONE 1,SIDETONE 2, and aural warning tones are allsimultaneously output to the headsets only,wi thout ind iv idua l se lec t ion capabi l i ty(Sidetone = transmitted audio sent to the head-set to moni tor qual i ty of t ransmiss ion) .Therefore each radio volume control shouldbe turned down to listen to the appropriateunit for that point in time. A Morse code iden-tifier check will not be possible on the re-maining NAV radios during the EMER mode.

COMMUNICATION RADIOS

GeneralTwo communications transceivers (VHF-22A) are installed and are each controlledby a CTL-22 control. This control is shownin Figure 16-2 and detailed operation/de-scription are shown in Table 16-1.

Operation When the COMM radio is first turned on, itsounds a brief tone while the microprocessorchecks its own memory. If there is a memorydefect, the tone is continuous to indicate thatthe transceiver can neither receive nor trans-mit. After the memory check, the displayshows the same active and preset frequenciesthat were present when the equipment waslast turned off.




CONTROL OR DISPLAY FUNCTION/DESCRIPTION

ACTIVE FREQUENCY DISPLAY The active frequency (frequency to which the VHF-22A is tuned) and dashes (----) on dIAG during self test.

PRESET FREQUENCY DISPLAY The preset (inactive) frequency and diagnostic messages are displayed in the lower window.

COMPARE ANNUNCIATOR ACT momentarily illuminates when frequencies are being changed. If ACT continues to flash, the actual radio frequency is not identical to the frequency shown in the active window.

ANNUNCIATORS The COMM control contains MEM (memory), and TX (transmit) annunciators. The MEM annunciator illuminates whenever a frequency is displayed in the lower window. The TX annunciator illuminates whenever the VHF-22A is transmitting.

POWER AND MODE SWITCH The OFF–ON positions switch system power to turn the system on or off. The SQ OFF position disables the receiver squelch circuits, so you should hear

noise. Use this position to set the volume control or, if necessary, to try to receive a very weak signal which cannot break squelch.

LIGHT SENSOR The built-in light sensor automatically controls the display brightness. The ANN PUSH BRT control knob/push button can be used to override the automatic dim controls and force the display to go to full bright.

XFR/MEM SWITCH This switch is a three-position, spring-loaded toggle switch. When moved to the XFR position, the preset frequency is transferred up to the active display and the VHF-22A retunes. The previously active frequency becomes the new preset frequency and is displayed in the lower window. When this switch is moved to the MEM position, one of the six stacked memory frequencies is loaded into the preset display. Successive pushes cycle the six memory frequencies through the display (...2, 3, 4, 5, 6, 1, 2, 3...).

FREQUENCY SELECT KNOBS Two concentric knobs control the preset or active frequency display. The larger knob changes the three digits to the left of the decimal point in 1-MHz steps. The smaller knob changes the two digits to the right of the decimal point in 50-kHz steps (or in 25-kHz steps for the first two steps after the direction of rotation has been reversed). Numbers roll over at the upper and lower frequency limits.

ACT BUTTON Push the ACT button for approximately two seconds to enable the frequency select knobs to directly retune the VHF-22A. The bottom window will display dashes and the upper window will continue to display the active frequency. Push the ACT button a second time to return the control to the normal two-display tune/preset mode of operation. The active tuning feature is not affected by power removal. If active tuning is selected (one push of the ACT button) and power is removed from the control, active tuning will still be enabled the next time power is reapplied to the control.

STO BUTTON The STO button allows up to six preset frequencies to be selected and entered into the control’s nonvolatile memory. To store a frequency, simply toggle the MEM switch until the upper window displays the desired channel number (CH 1 through CH 6), rotate the frequency select knobs until the lower window displays the frequency to be stored, and push the STO button twice within five seconds.

After approximately five seconds, the control will return to the normal two-display tune/preset mode of operation.

TEST Push the TEST button to initiate the radio self-test diagnostic routine. The transceiver performs a complete self-test routine requiring about five seconds.

Table 16-1. CTL-22 COMM CONTROL, CONTROLS AND INDICATIONS




Moving the mode switch (Figure 16-2) to SQOFF and then adjusting the background noiselevel can help set the radio volume. After acomfortable level has been established, re-turn the mode switch to ON to return thesquelch back to normal.

Whenever a new frequency is selected in theactive window, an ACT annunciator (com-pare annunciator) illuminates to indicate thetransceiver is changing frequencies and thenextinguishes after the process is complete. Ifit continues to flash, then the selected fre-quency on the CTL-22 is not the frequencybeing used by the transceiver. A recycling ofthe frequency should be accomplished (veri-fying the ACT annunciator extinguishes) or theuse of a different radio.

For frequency selection and storing refer toTable 16-1.

Stuck MIC Protection

Each time the Push-to-Talk (PTT) button ispushed the microprocessor in the transceiverstarts a two-minute timer. If the transmitter isstill on at the end of two minutes the micro-processor turns it off. This protects the ATCchanne l f r om long - t e rm in t e r f e r ence .Transmission is indicated by a TX annuncia-tor on the CTL-22 control and continuous il-lumination without pressing the PTT buttonwould indicate this malfunction (Figure 16-2and Table 16-1). This annunciator should ex-tinguish at the two-minute time limit.

MEMORYANNUNCIATOR

COMPAREANNUNCIATOR

PRESET COMMFREQUENCY

DISPLAY

CTL-22 COMM CONTROL

POWER ANDMODE SELECT

SWITCH

COMM VOLUMECONTROL

LIGHTSENSOR

TESTBUTTON

MEMORYSTORE

BUTTON

ACTIVE TUNEBUTTON(ACTIVE TUNING/PRESET TUNING)

FREQUENCYSELECTKNOBS (2)

TRANSMITANNUNCIATOR

ACTIVE COMMFREQUENCYDISPLAY

TRANSFER/MEMORYSWITCH

Figure 16-2. VHF-22 COMM Radio Controls/Displays

Overtemperature ProtectionThe temperature of the transmitter is contin-uously monitored by a microprocessor. If thetemperature exceeds +160°C (+320°F) thetransmitter is shut down. This immediatelyeliminates sidetone. A release of the PTT willsound two beeps, and as long as the tempera-ture remains above the limit, the transmitterwill not operate. If you must transmit, youcan override the protection by rapidly keyingthe PTT button twice and holding it on thesecond push.

Self TestAn extensive self-test diagnostic routine can beinitiated in the transceiver by pushing the TESTbutton on the radio. During the self test, theupper and lower displays modulate from mini-mum to maximum lighting intensity to indicatethe self test is in progress. Several audio toneswill be heard from the audio system during thistest. At the completion of the self test the radiousually displays four dashes (- - - -) in the upperdisplay and 00 in the lower display (Figure 16-3). If any out-of-limit conditions are found, theupper display will read dIAG and the lower dis-play will contain a two-digit diagnostic code(Figure 16-3).

Navigation/DME EquipmentTwo navigation receivers (VIR-32) are installedand are each controlled by a CTL-32 control.This control is shown in Figure 16-4 and detailedfunction/descriptions are listed in Table 16-2.

Two DME transceivers (DME-42) are also in-stalled and they show information on the re-spective indicator (IND-42). These are shownin Figure 16-5 with a detailed function/de-scription in Table 16-3.

OperationThe navigation radios provide VOR, LOC,GS and marker beacon capabilities. On-sideor cross-side course display information canthen be selected on the pilot and copilot dis-plays via push buttons or an EFIS system.

Selection of frequencies and storing are iden-tical to that described in the communicationssection above.

The DME equipment will indicate the slant-range distance when the on-side navigationradio contains a DME associated frequency.The DME identifier is sent once every 30seconds and when received, the indicator




TEST DISPLAY FAILURE PRESENTTEST DISPLAY NO FAULT PRESENT

DIAGNOSTICANNUNCIATOR

DIAGNOSTICCODE

Figure 16-3. VHF-22 COMM Radio Self Test Displays




Table 16-2. CTL-32 NAV/DME CONTROL, CONTROLS AND INDICATIONS


ACTIVE FREQUENCY DISPLAY Active frequency is the frequency to which the DME-42 channel 1 is tuned.In DME or NAV self test, diagnostic messages are displayed in the upperwindow.

PRESET FREQUENCY DISPLAY The preset (inactive) frequency is the frequency to which the DME-42channel 2 is tuned. In DME or NAV self test, diagnostic messages aredisplayed in the lower window.

COMPARE ANNUNCIATOR ACT momentarily illuminates when frequencies are being changed. If ACTcontinues to flash, the actual tuned frequency is not identical to thefrequency shown in the active display.

ANNUNCIATORS

MEM

HLD

The NAV control contains MEM (memory) and HLD (hold) annunciators.

The MEM annunciator illuminates when a frequency is displayed in thelower window.

The HLD annunciator indicates the DME is in DME hold. In this mode it isnormally tuned to the frequency displayed in the active window at the timeof selection. After selecting hold, the upper window displays the NAVfrequency and the lower window displays the DME hold frequency. Tuningof the active frequency can take place during this time. When completed,the unit will always revert back to display of the DME hold frequency in thelower window.

VOLUME CONTROL The volume control is concentric with the power and mode switch. Itcontrols only the NAV receiver volume.

POWER AND MODE SWITCH The NAV control power and mode switch contains three detentedpositions. The positions are: OFF–ON–HLD.

The OFF–ON positions switch system power.

The HLD position allows the NAV frequency to be changed but holds theDME to the current active NAV frequency.

LIGHT SENSOR The built-in light sensor automatically controls the display brightness. TheANN PUSH BRT control knob/push button can be used to override theautomatic dim controls and force the display to go to full bright.

XFR/MEM SWITCH This switch is a three-position, spring-loaded toggle switch. When moved tothe XFR position, the preset frequency is transferred up to the activedisplay and the NAV/DME retunes. The previously active frequencybecomes the new preset frequency and is displayed in the lower window.When this switch is moved to the MEM position, one of the four stackedmemory frequencies is loaded into the preset display. Successive pushescycle the four-memory frequencies through the display (...2, 3, 4, 1, 2, 3....).

FREQUENCY SELECT KNOBS Two concentric knobs select the preset or active frequency displays. Thelarger knob changes the two digits to the left of the decimal point in 1-MHzsteps. The smaller knob changes the two digits to the right of the decimalpoint in 0.05-MHz steps. Frequencies roll over at the upper and lowerlimits. The two frequency select knobs are independent of each other suchthat the upper and lower limit rollover of the 0.1-MHz digit will not causethe 1.0-MHz digit to change.




Table 16-2. CTL-32 NAV/DME CONTROL, CONTROLS AND INDICATIONS (Cont)

V

MARKER BEACONVOLUME CONTROL

MARKER BEACONSENSITIVITY SWITCH

MEMORYANNUNCIATOR DME HOLD

ANNUNCIATOR

COMPAREANNUNCIATOR

NAV VOLUMECONTROL


SWITCH

MEMORYSTORE

BUTTON

TESTBUTTON



ACTIVE VOR/LOCFREQUENCYDISPLAY

PRESET VOR/LOCFREQUENCY

DISPLAY

LIGHTSENSOR


CTL-32 NAV CONTROL

Figure 16-4. VIR-32 NAV Receiver Controls/Displays


ACT BUTTON Push the ACT button for approximately two seconds to enable the frequencyselect knobs to directly retune the VIR-32 and DME. The bottom window willdisplay dashes and the upper window will continue to display the activefrequency. Push the ACT button a second time to return the control to the normaltwo-display tune/preset mode of operation. The active tuning feature is notaffected by power removal. If active tuning is selected (one push of the ACTbutton) and power is removed from the control, active tuning will still be enabledthe next time power is reapplied to the control.

STO BUTTON The STO button allows up to four preset frequencies to be selected and enteredinto the control’s nonvolatile memory. To store a frequency, simply toggle theMEM switch until the upper window displays the desired channel number (CH 1through CH 4), rotate the frequency select knobs until the lower window displaysthe frequency to be stored, and push the STO button twice within five seconds.After approximately five seconds, the control will return to the normal two-displaytune/preset mode of operation.

TEST BUTTON Push the TEST button to initiate the radio self test diagnostic routine. (In the caseof the VIR-32 NAV receiver, self-test is active only while the TEST button ispushed or about 15 seconds maximum. In the case of the DME-442 transceiver,the self test routine requires about 10 seconds for completion.)




VOL VOL

DME1 2

ENCDALTM

1

COLLINS

SEL PWR

18.8.8.18.8.8.1 2 3 1 2 3 WPT WPT NMNM

8 8 8 8HLD HLD KT KT MIN MIN IDID

COLLINS

CH SEL PWR

18.8.8.18.8.8.1 2 3 1 2 3 WPT WPT NMNM

8 8 8 8HLD HLD KT KT MIN MIN IDID

DISPLAYGROUND SPEEDTIME TO STATIONSTATION IDENT

No. 2 DMEAUDIO VOLUMENo. 1 DME

AUDIO VOLUME

DISTANCE DISPLAY

CHANNELANNUNCIATORS

DISTANCE LABEL

IND-42A (PILOT'S DISPLAY)

DISPLAY ANNUNCIATORS

LIGHT SENSOR

POWERSWITCH

SELECTOR SWITCH(KT, MIN, OR ID)

CHANNELSWITCH(1, 2, 3)

IND-42C (COPILOT'S DISPLAY)

X X X X

X X X X

Figure 16-5. DME-42 Systems IND-42 Displays

will show the DME identifier (2-, 3-, or 4-let-ter identifier). This information can be alteredto show the ID, KT (velocity) and MIN (time-to/from station) through repeated pushes of theSEL button (Figure 16-5 and Table 16-3).

A HOLD function is provided on the NAV powerand mode switch (Figure 16-4) and allows se-lection of a different NAV frequency withoutchanging DME stations (e.g., flying with anILS frequency tuned in the active selection but

ANNUNCIATOR DESCRIPTION

CONTROL/INDICATOR FUNCTION/DESCRIPTION

NUMERIC DISPLAY The numeric display presents the NM (distance) and diagnostic code.

ALPHANUMERIC DISPLAY The alphanumeric display presents the KT (velocity), MIN (time-to-station), ID(2-, 3-, or 4-letter station identifier), and diagnostic identifier.

POWER (PWR) SWITCH The latching push-on/push-off PWR switch controls the power applied to theIND-42.

MODE SELECTOR (SEL) SWITCHALPHANUMERIC

The non latching pushbutton SEL switch selects the information to be displayedin the display. (When power is initially applied, NM (distance) is shown in thenumeric display and ID (DME station identifier) is shown in the alphanumericdisplay.) Pressing the SEL switch will sequentially select KT (velocity), MIN(time-to-station), and ID (2-, 3-, or 4-letter station identifier).KT, MIN, and ID are shown in the alphanumeric display and NM (distance) iscontinuously shown in the numeric display, provided the DME is locked on asignal.

CHANNEL (CH) SWITCH(IND-42A ONLY)

The momentary pushbutton CH switch sequentially selects the information fromthe next DME channel and lights the appropriate channel annunciator 1, 2, or 3.The copilot’s IND-42C will always power up on channel 2

ANNUNCIATORS The annunciators provide an indication of which DME channel is selected,system operational information, and units of measure. The following listdescribes the annunciators.

1 2 3Sequentially controlled by the channel (CH) button to indicate which DMEchannel is providing the information being displayed in the numeric andalphanumeric displays.

NM Automatically illuminates after power on when valid DME data is available.Indicates that the numbers displayed in the numeric display are slant rangeDME distance in nautical miles.

HLD Indicates that DME hold has been selected on the CTL-32 NAV Control. Whenin HLD, KTS and MIN will revert to ID after approximately 5 seconds.

KT Indicates that the value displayed in the alphanumeric display is the computedrate of change of DME distance.

MINIndicates that the value displayed in the alphanumeric display is the computedtime-to-station in minutes.

ID

Automatically illuminates after power on. The DME ident is transmitted onceevery 30 seconds and it is possible that 2 minutes could elapse before thestation ident is displayed in the alphanumeric display. The station identifier isusually 3 letters, but can be 2, 3, or 4 letters, depending on the type of facilitybeing received.




Table 16-3. IND-42A/C DME INDICATOR, CONTROLS AND INDICATIONS

using the DME from a VOR). To operate in theHOLD mode, first select the desired DME sta-tion and then move the power and mode switchto HOLD. The frequency being held will appearin the preset window. The CTL-32 and DME in-dicator will now show HLD in the annunciatorsection as a reminder of the HOLD selection.(KT and MIN can be selected on the DME in-dicator during HOLD operations; however, theindicator will default to ID after five seconds.)Although, the preset frequency is showing the“held” station, any previously selected fre-quency still remains in memory and a movementof the XFR/MEM switch to XFR will move thatpreset frequency to the active window. If fre-quency selection is done by a movement of thefrequency select knobs, only the active win-dow will change. To return to preset frequencyselection the power and mode switch must beturned to ON. Frequencies can still be retrievedfrom the memory during HOLD operations asdiscussed in Table 16-2.

In a dual DME installation the copilot’s indi-cator will usually allow selection of differentDME channels. By repeatedly pushing the CHbutton, these channels can be cycled. The cur-rent selected option will be indicated on thedisplay as 1, 2, or 3 (Figure 16-5). A typicalinstallation of channel usage is shown in Figure16-6. If the pilot does not have channel se-lection then channel 1 is the default indication.

Self TestLike the communication radios, the NAV ra-dios provide the ability for an extensive self-

test sequence. During the self test, the upperand lower displays modulate from minimumto maximum lighting intensity to indicate theself test is in progress. The DME will be placedin self test at the same time.

VOR Self TestSelect a VOR frequency on the CTL-32 NAVcontrol. (108.20 MHz will do. A specific fre-quency is not required for test.) A signal on thefrequency will not interfere with the self test.

• Select VOR-1 or -2 (as required) as theactive course sensor on the EHSI.

• Rotate the Course Select knob to ap-proximately 0°.

• Push and hold the TEST button on theCTL-32.

• The active course sensor VOR1 or 2annunciator on the EHSI will turn red.

• After approximately two seconds, theVOR1 or 2 annunciator will turn green,the EHSI lateral deviation bar will ap-proximately center, and a TO indicationwill appear. The RMI pointers con-nected to the VIR-32 will indicate ap-proximately 0° magnetic bearing.

• Release the TEST button. (The VIR-32will return to normal operation after ap-proximately 15 seconds, even if theTEST button is held.)




CH1 - NAV 1 ACTIVE2 - NOT USED3 - NAV 1 PRESET

1 2 3 DME 1

NAV 1ACT ACTPRE PRE 1 2 3 DME 2

CH1 - NOT USED2 - NAV 2 ACTIVE3 - NAV 3 PRESET

DME 2

NAV 2

Figure 16-6. Typical Pro Line II Dual DME Installation

ILS (Localizer and Glideslope)Self-TestSelect a localizer frequency on the CTL-32NAV control. (108.10 MHz will do. A specificfrequency is not required for test.)

• Select LOC1 or 2 (as required) as the ac-tive course sensor on the EHSI.

• Push and hold the TEST button on theCTL-32.

• The active course sensor LOC1 or 2annunciator on the EHSI will turn redand the red GS flag will come into view.

• After approximately three seconds,the LOC1 or 2 annunciator will turngreen and the GS flag will go out ofview, the EHSI lateral deviation barwill deflect right approximately two-thirds of full scale, and the glides-l ope po in t e r w i l l de f l e c t downapproximately two-thirds of full scale.

• Release the TEST button. (The VIR-32will return to normal operation after ap-proximately 15 seconds, even if theTEST button is held.)

Marker Beacon Self-TestThe marker beacon assembly is tested auto-matically when the TEST button on the CTL-32 is pushed and either a VOR or localizerfrequency is selected. For No. 1 NAV receiverproper operation of the marker beacon as-sembly is indicated by all three-marker an-nunciators on the EADI cycling through inorder. For No. 2 NAV receiver the indicationwill be the three marker annunciators flick-ering at 30Hz. In addition, a tone will also bepresent in the marker beacon audio output.

DME Self Test• Turn power on to the DME, NAV, and

EFIS systems.

• Ensure that VOR or LOC is selected asthe NAV source on the HSI.

• On the CTL-32, select ON. Use the fre-quency select knobs and select the fre-quency for any DME or VORTAC stationthat is within range.

• Read the distance to the station on theIND-42A/C and the left side of the EHSIdisplay. Verify the station ID next to thedistance display.

NOTEThe DME can require at least 30 sec-onds, and as much as two minutes, toproperly decode the station ident.

• On the CTL-32, push TEST. On the IND-42 the following happens:

• Initially the IND-42A/C display mod-ulates in intensity between maximumand minimum brightness.

• LH display on IND-42A/C shows atest distance of 100.0(nm). After about10 seconds, the RH display shows anAOK (Figure 16-7).

• Listen to DME audio and note thataud io i s a Mor se code A O K(• – – – – – • –).

• Push SEL to annunciate KT and read100 (knots) in RH window.

NOTEIf the 10-second self test expires be-fore reaching this point, select TSTagain and continue with the test.

If there are any detected faults in the system onthe IND-42A/C a diagnostic code will appearin place of the AOK display (Figure 16-7). TheEFIS display will only show dashes for a fault.

The diagnostic routines are intended as an ex-tension of the self-test capability. The opera-tor should first observe the deviation indicatorsand associated flags for the proper self-test re-sponses. If an out-of-limit condition exists,then the problem may be verified in more de-tail by the diagnostics.

For the first two or three seconds immediatelyafter the TEST button on the CTL-32 is pushed,a two-digit diagnostic code may be displayedin the lower window based on the conditions ex-isting immediately before the TEST button waspushed. Four dashes will be displayed along




with the code 00, indicating normal operation,no trouble found (Figure 16-8). If an out-of-limitcondition is detected during self test, that two-digit code will also be displayed on the CTL-32 along with the word dIAG (diagnostic) orFLAG. FLAG will be displayed along with atwo-digit code when something is abnormal buta failure has not occurred (i.e., low signal level,etc.) (Figure 16-8). dIAG is displayed alongwith a two-digit code to indicate a failure hasbeen detected in the VIR-32 (Figure 16-8).

Completion of self test is indicated when theNAV control displays either the normal activeand preset frequencies in the upper and lowerwindows, respectively, or a two-digit code.

ADF EQUIPMENT

GeneralThe B200 aircraft can have either single ADFreceiver installed or dual ADFs. The ADF-60A

is controlled through the CTL-62 control. Thiscontrol is shown in Figure 16-9 and detailedfunction/descriptions are listed in Table 16-4.

The ADF receives transmissions from a se-lected ground station, indicates relative bear-ing to that station, and provides audio fordetermining station identification and listen-ing to voice announcements.

The ground station must be within the normaloperating range of 190 to 1749.5kHz. Thereare three functional modes of operation. In




IND-42A SELF-TEST DISPLAY NO FAULTS FOUND

IND-42C SELF-TEST DISPLAY 03 (TRANSMITTER) FAULT FOUND

2

Figure 16-7. IND-42 Self Test Displays

TEST DISPLAY NO FAULT PRESENT

TEST DISPLAY ABNORMAL OPERATION PRESENT

FLAGANNUNCIATOR

DIAGNOSTICCODE

TEST DISPLAY FAILURE PRESENT


DIAGNOSTICCODE

Figure 16-8. VIR-32 NAV ReceiverSelf-Test Displays

ANT mode, the ADF receiver functions as anaural receiver, providing only an aural outputof the received signal. In ADF mode, it func-tions as an automatic direction finder receiverin which bearing-to-the-station is presented onan associated bearing indicator and an auraloutput of the received signal is provided. ATONE mode provides a 1000-Hz aural outputtone when a signal is being received to allowidentification of keyed CW signals.

ADF Self Test• Apply power to the ADF and RMI, and

EFIS systems.

• Select appropriate bearing pointer on theRMI and EHSI for the ADF to be tested.

• Set the control mode to ADF. This ap-plies power to the system.

• Tune the control to any frequency,preferably one that is known to give agood signal.

• Note that the RMI pointer should indi-cate the correct relative bearing to thestation. Adjust the audio volume, as nec-essary, for a comfortable listening level.

NOTEIf the RMI pointer remains parked,the system may not be receiving areliable signal. In this case, try twoor three other stations if possible.

• Push and hold the self-test switch. Notethe RMI and EHSI bearing pointer ro-tates 90° counterclockwise. Releaseself-test switch.

NOTEIf the signal is weak or of poor qual-ity, the bearing pointer can rotaterather slowly. Degraded receiversensitivity might give the same re-sponse.

• Tune to any CW station if one is avail-able. Set the mode to TONE and listento the audio for a 1000-Hz tone iden-tifying the CW station.




MEMORYANNUNCIATOR

COMPAREANNUNCIATOR

NAV VOLUMECONTROL


SWITCH

MEMORYSTORE

BUTTON

TESTBUTTON



ACTIVE ADFFREQUENCYDISPLAY

PRESET ADFFREQUENCY

DISPLAY

LIGHTSENSOR


CTL-62 ADF CONTROL

Figure 16-9. ADF-60A ADF Receiver Controls/Displays




Table 16-4. CTL-62 ADF CONTROL, CONTROLS AND INDICATIONS


ACTIVE FREQUENCY DISPLAY The active frequency; the frequency to which the ADF-60A is tuned. In self-testmode and if an out-of-tolerance condition is detected, the word “dIAG” isdisplayed in the upper window while the diagnostic code is displayed in thelower window.

PRESET FREQUENCYDISPLAY

The preset frequency is displayed in the lower window. In self-test mode and ifan out-of-tolerance condition is detected, the diagnostic code is displayed inthe lower window.

COMPARE ANNUNCIATOR ACT momentarily illuminates when frequencies are being changed. If the ACTannunciator continues flashing, the receiver is not tuned to the displayedactive frequency.

ANNUNCIATORSThe ADF control contains a MEM (memory) annunciator. The MEMannunciator illuminates whenever a frequency is displayed in the lower window

VOLUME CONTROL The volume control, is concentric with the power and mode switch andcontrols ADF audio volume.

LIGHT SENSOR The built-in light sensor automatically controls the display brightness. The annpush brt control knob/push button can be used to override the automatic dimcontrols and force the display to go to full bright.

POWER AND MODE SWITCH

OFF

ANT

ADF

TONE

The power and mode switch contains four detented positions.

The OFF position interrupts system power ( Turns the ADF off). Selecting ANT,ADF, or TONE applies power to the ADF system and establishes the systemmode of operation.

In ANT mode, the ADF receiver functions as an aural receiver, providing onlyan aural output of the received signal

In ADF mode, it functions as an automatic direction finder receiver in whichbearing-to-the-station is presented on an associated bearing indicator and anaural output of the received signal is provided.

TONE mode provides a 1000-Hz aural output tone when a keyed CW signal isbeing received.

XFR/MEM SWITCH This switch is a 3-position, spring-loaded toggle switch. When moved to theXFR position, the preset frequency is transferred up to the active display andthe ADF-60 retunes. The previously active frequency becomes the new presetfrequency and is displayed in the lower window. When this switch is moved tothe MEM position, one of the four stacked memory frequencies is loaded intothe preset display. Successive pushes to the MEM position cycles the fourmemory frequencies through the display (...2, 3, 4, 1, 2, 3....). The frequencythat was in the preset window is:

1. Maintained in memory if it was originally assigned as a stored frequency, or2. Discarded if it was originally direct tuned either as a preset or active.

1/2 SWITCH (DUAL ADFINSTALLATION WITH ONECONTROL)

The 1/2 switch connects the tuning knobs to either the upper window or lowerwindow for tuning, 1 for upper and 2 for lower




Table 16-4. CTL-62 ADF CONTROL, CONTROLS AND INDICATIONS (Cont)


TUNING Normally, tuning is accomplished by entering a frequency into the presetwindow and then either storing that frequency in memory (STO) or entering itinto the active window (XFR) to tune the receiver. An alternate method is topress the ACT button for at least 2 seconds (this gives direct tuning access tothe upper window) and insert the desired frequency directly into the activewindow.

FREQUENCY SELECT KNOBS Two concentric knobs control the preset or active frequency displays. Thelarger knob changes the 1000’s and 100’s kHz digits. The smaller knobchanges the 10’s, units, and tenths kHz digits. Each detent of the larger knobchanges the frequency in 100-kHz steps. Each detent of the smaller knobchanges the frequency in 1-kHz steps with the exception that the first twodetent positions following a change in rotational direction will cause a 0.5-kHzchange. Rapid rotation of the smaller knob will cause frequency changesgreater than 1 kHz as a function of the rate of rotation. Frequencies roll overat the upper and lower limits. The two frequency select switches areindependent of each other such that the upper and lower limit rollover of the10-kHz digit will not cause the 100-kHz digit to change.

ACT BUTTONPush the ACT button for approximately 2 seconds to directly change theactive display window with the frequency select knob. The bottom windowwill display dashes. Push the ACT button a second time for about 2 secondsto return the control to the normal 2-display tune/preset mode of operation.The active tuning feature is not affected by power removal. If active tuning isselected (one push of the ACT button) and power is removed from thecontrol, active tuning will still be enabled the next time power is reapplied tothe control.

STO BUTTON The STO button allows up to four preset frequencies to be selected andentered into the control’s nonvolatile memory. To store a frequency, simplytoggle the MEM switch until the upper window displays the desired channelnumber (CH 1 through CH 4), rotate the frequency select knobs until thelower window displays the frequency to be stored, and press the STO buttontwice within 5 seconds. After approximately 5 seconds, the control will returnto the normal 2-display tune/preset mode of operation.

TEST BUTTON Push the TEST button to initiate the radio self-test routine. Self-test is activeonly while the TEST button is pushed. The display modulate in intensity whilethe TEST button is pushed.

When the CTL-62 is used with the CAD-62to control the ADF-60A system, certain di-agnostics codes can be displayed on theCTL-62 in the self-test mode. The diagnos-tic display appears when the self-test buttonis pressed as above for the ADF self test. Ifthe diagnostics detect no faults, the CTL-62displays four dashes (- - - -) and code 00(Figure 16-10). If the diagnostics detect afault dIAG will be displayed along the faultcode on the CTL-62 display (Figure 16-10).

TRANSPONDER EQUIPMENT

GeneralTwo TDR-94 transponders are installed in theB200 aircraft with only one operating at anyone time. The TDR-94 is a mode-A, mode-Cand mode-S transponder and is an integralpart of the Air Traffic Control Radar BeaconSystem. These transponders are controlled bya CTL-92 and is shown in Figure 16-11 withdetailed function/descriptions in Table 16-5.




TEST DISPLAY NO FAULT PRESENT TEST DISPLAY FAILURE PRESENT

DIAGNOSTICCODE


Figure 16-10. ADF-60A ADF Receiver Self-Test Displays

COMPAREANNUNCIATOR


SWITCHIDENT

BUTTON

TESTBUTTON

CODESELECTKNOBS (2)

ACTIVE CODEDISPLAY

LIGHTSENSOR

NO. 1 OR NO. 2SELECT SWITCH

IDENTDISPLAY

(DISPLAYED WHENIDENT BUTTON USED)

PRESETBUTTON

TRANSPONDER REPLYANNUNCIATOR

ENCODINGALTIMETER

SOURCE SELECTSWITCH

Figure 16-11. TDR-94 Transponder Controls/Displays




FUNCTION/DESCRIPTION

ANNUNCIATOR

CONTROL OR DISPLAY

UPPER DISPLAY WINDOW

LOWER DISPLAY WINDOW

The ATC code (code with which the active transponder replies) and diagnostic messages are displayed in the upper display window. During normal operation, the CTL-92 has only a single display (the transponder code) shown in the upper window.

The lower display window is normally blank. It is active only during self test. If a fail/warn condition is detected, dIAG will be displayed. Press the TEST button to view the diagnostic code.

COMPARE ANNUNCIATOR ACT momentarily illuminates when codes are being changed. If ACT flashes, the actual reply code is not identical to the code shown in the active code display.

The ATC control annunciator contains a TX (transmit) annunciator. The TX annunciator illuminates when the transponder replies to an interrogation.

POWER AND MODE SWITCH

The ATC control power and mode switch contains four detented positions. Available positions are: OFF–STBY–ON–ALT.

Power is removed in the OFF position and is applied when any of the other modes is selected.

In the STBY mode, power is applied to the transponder but it is prevented from transmitting replies. STBY should be used only during taxi or when requested by ATC.

The ALT position is the normal operating position and allows the transponder to reply to the interrogation pulses, as well as transmitting uncorrected barometric altitude when the transponder is interrogated in mode C.

The ON position deletes the altitude code and is normally used when requested by ATC.

1/2 SWITCH The 1/2 switch selects which of two transponders is active.

LIGHT SENSORThe built-in light sensor automatically controls the display brightness. The ANN PUSH BRT control knob/push button can be used to override the automatic dim controls and force the display to go to full bright.

CODE SELECT KNOBS

Two concentric knobs control the active code display. The larger knob changes the two more significant digits, and the smaller knob changes the two less significant digits. The less significant digit is incremented or decremented for each detent of the smaller knob if the knob is slowly turned. Rapid rotation of either knob will cause changes proportional to the rate of rotation. Rollover of the less significant digits will occur at 0 and 7, and will cause the more significant digits to be incremented or decremented. The left two digits and the right two digits are independent of each other. The various codes used for normal operation are listed in the Aeronautical Information Manual. Codes 7600 or 7700 are selected for in-flight emergency operation and will be annunciated by the codes flashing in the active code display for a couple of seconds before transmission begins.

PRE BUTTONPush and hold the PRE button while turning the code select knobs to select a preset code for storage. The preset code will be stored in nonvolatile memory and can be recalled by momentarily pressing the PRE button again.

IDENT BUTTONThe IDENT button causes the transponder to transmit a special identification pattern that is displayed on the ground controller’s radar scope. This button should be pushed only when you are requested to “squawk ident” by the ground controller.

Table 16-5. CLT-92 ATC CONTROL, CONTROLS AND INDICATIONS

OperationBoth transponders provide identification(mode-A) of the aircraft on the ATC groundcontroller ’s plan posit ion indicator. TheTDR-94 transponder also provides aircraftpressure altitude to the ground controller’sindicator (mode-C). The transponders aresent altitude data information from the pilot’saltimeter or the copilot’s altimeter (selectionis via an encoding altimeter switch on theaudio panel) (Figure 16-11). In normal mode-A or mode-C operation, the TDR94 is inter-rogated by radar pulses from a ground stationand replies automatically with a series ofpulses. The TDR-94/94D can operate inmode-S and provide a unique aircraft iden-tification code as well as air-to-air and air-to-ground interrogation replies. The unitalso has data link capability, which allows itto perform additional air traffic control andaircraft separation assurance functions. TheTDR-94 transponder can also transmit anident pattern when requested to squawk identby the ground controller. This can be donevia the IDENT button on the transponder orwith a push of the IDENT button on the backof each yoke.

For mode-S operation an air-ground signal froma strut switch is input to the transponder. This

operates the “on ground” or “in air” state of themode-S information.

Self TestTo carry out the transponder self-test posi-tion the mode switch to ON and select the de-sired transponder to be tested. Set the desiredcode using the code select knobs. Push theTEST button on the CTL-92. During self testthe CTL-92 display flashes from minimumto maximum brightness. If there are no di-agnostic conditions detected, uncorrectedbarometric altitude is displayed on the bot-tom display line in hundreds of feet and theannuncia tor AL on the top d isp lay l ine(Figure 16-12).

If no altitude data is present the display willshow dashes (- - - -) without a diagnosticcode (Figure 16-12).

If a diagnostic condition is detected in theTDR-94/94D during self test, the upper win-dow displays the word dIAG while the lowerwindow displays a two-digit diagnostic code(Figure 16-12). To carry out self test on theopposite transponder use the select switch onCLT-92 to select that transponder.




TEST BUTTON Push theTEST button to initiate the radio self-test routine. In dual version units, the 1/2 switch determines which transponder responds to the test command.

ENCODING ALIMETER SELECT SWITCH

This switch selects which altimeter, the pilot’s (ALTM1) or copilot’s (ALTM2), will provide encoding altimeter information to the transponders.

SELF-TEST DISPLAY

NO FAILURE

During self test, the active code display intensity will modulate from minimum to maximum. If the transponder is functioning properly and an altitude encoder is connected to the CTL-92 and operating, AL will be displayed in the upper window and the altitude in thousands of feet in100-foot increments will be displayed in the lower window.

FAILUREIf an out-of-tolerance condition is detected, the upper window shows the word DIAG while the lower window shows a two-character diagnostic code.

FUNCTION/DESCRIPTIONCONTROL OR

DISPLAY

Table 16-5. CLT-92 ATC CONTROL, CONTROLS AND INDICATIONS (Cont)

If during normal operation a “fail/warn” con-dition is detected the CTL-92 will display thedIAG message in the lower window withouthaving pressed the TEST button (Figure 16-12). When this occurs, press the TEST but-ton to view the associated diagnostic code.

AntennasA typical installation of antennas is shownin Figure 16-13.

EFIS and other AvionicsFor specific EFIS and avionics equipment notdiscussed here refer to the appropriate pilotguides, AFM supplements and other appro-priate information.

FLIGHT INSTRUMENTS

PITOT AND STATIC SYSTEMThe pitot and static system (Figure 16-14)provides a source of impact air and static airfor operation of the flight instruments. Twoheated pitot masts (Figure 16-15) are locatedon each side of the lower portion of the nose.Tubing from the left pitot mast is connectedto the pilot’s airspeed indicator, and tubingfrom the right pitot mast is connected to thecopilot’s airspeed indicator.





DIAGNOSTICCODE

TEST DISPLAY FAULT PRESENT

FOUR DASHES

TRANSPONDER IN ALT MODE OF OPERATIONWITH NO ALTITUDE DATA PRESENT

ALTITUDEANNUNCIATOR

CURRENT AIRCRAFTALTITUDE IN 100FT

INCREMENTS

TEST DISPLAY NO FAULT PRESENT

ACTIVE CODEDISPLAY


NORMAL TRANSPONDER OPERATIONWITH A FAULT DETECTED

Figure 16-12. TDR-94 Transponder Self-Test Displays




WE

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PILOT'S STATICAIR SOURCE

NORMAL ALTERNATE

+ + +

SEE FLIGHT MANUAL PERFORM-ANCE SECTION FORINSTR CAL ERROR

(POSITIONED ABOVE S1)

S1 S2

DRAIN VALVE

PILOT'S ALTERNATESTATIC AIR

COPILOT'S STATIC AIR

PILOT'S STATIC AIR

PILOT'S ALTERNATESTATIC AIR

TO COPILOT'SINSTRUMENTS

PILOT'S PITOT

COPILOT'S PITOT

DIFFERENTIAL PRESSURE SWITCH(FOR LANDING GEAR WARNING

ON BB-324 THRU BB-452THAT ARE NOT IN COMPLIANCE

WITH SI 1047)

A

TO PILOT'SINSTRUMENTS

PILOT'S STATIC AIR

COPILOT'S STATIC AIR

PRESSURE BULKHEAD

NOTE: ALTIMETERS ANDVERTICAL INDICATORS OMITTEDFROM THIS VIEW FOR CLARITY

PILOT'S STATIC AIRSOURCE CONTROL VALVE(VALVE IN "NORMAL" POSITION)

TO COPILOT'S INSTRUMENTS

TO PILOT'S INSTRUMENTS

COPILOT'S AIRSPEEDINDICATOR

PILOT'S AIRSPEEDINDICATOR

ADS-65 AUTOPILOT AIR DATA SENSOR

ADC-85 AIR DATA COMPUTER

MANIFOLD MANIFOLD

ASI

IVSI

ALT

ALT

IVSI

ASI

PPI

CDPI

REAR PRESSURE BULKHEAD

R/H PITOT MAST

DRAIN

DRAIN

DRAIN DRAIN

ALTERNATE STATIC SELECTOR VALVE

ALTERNATE STATIC PORT

R/H STATIC PORTS

L/H STATIC PORT

ALT = ALTIMETERIVSI = INSTANTANEOUS VERTICAL SPEED INDICATORASI = AIRSPEED INDICATORPPI = PNEUMATIC PRESSURE INDICATORCDPI = CABIN DIFFERENTIAL PRESSURE INDICATOR

S2 S1

S1 S2P2

P1

PILOT'S PITOT

COPILOTS PITOT

PILOTS STATIC

COPILOTS STATIC

ALTERNATE STATIC

LEGEND

PNEUMATIC PRESSURE

CABIN PRESSURE

ELECTRIC SIGNAL

FORWARD PRESSURE BULKHEAD

APC-65 AUTOPILOT COMPUTER

L/H PITOT MAST

BOTTOM TOP

BOTTOM TOP

Figure 16-14. Pitot and Static System Diagram

The normal static system provides two sepa-rate sources of static air: one for the pilot’sflight instruments and one for the copilot’sflight instruments. Each of the two static airlines open to the atmosphere through two staticports (Figure 16-16) on each side of the aftfuselage.

An alternate static air line is also provided forthe pilot’s flight instruments. In the event ofa failure of the pilot’s normal static air source,which could be caused by ice accumulationsobstructing the static ports (the static portsare not heated), the alternate static source maybe selected by lifting the spring-clip retaineroff the PILOT’S STATIC AIR SOURCE valveswitch (Figure 16-17), located under the copi-lot’s right side circuit-breaker panel, and plac-ing the switch in the ALTERNATE position.This connects the alternate line to the pilot’sflight instruments only. It obtains static air aft

of the rear pressure bulkhead from inside theunpressurized area of the fuselage.

The pilot’s airspeed and altimeter nor-mal indications are changed when thealternate static air source is in use.Refer to the Airspeed Calibration-Alternate System, and the AltimeterCorrection Alternate System graphsin the Flight Manual (PerformanceSection) for operation when the alter-nate static air source is in use. (The ver-tical speed indicator is also affected,but no correction table is available.)

When the alternate static air source is not re-quired, the pilot should ensure the PILOT’SSTATIC AIR SOURCE valve switch is held inthe NORMAL (forward) posi t ion by thespring-clip retainer.

OUTSIDE AIR TEMPERATUREGAGEThe outside air temperature (OAT) gage onvery early models is located on the overheadpanel adjacent to the oxygen controls. Theprobe and sunshield protrude through the skinat the top of the fuselage. A button located onthe overhead panel must be depressed to illu-minate a post light next to the gage for nightflight.

WARNING




Figure 16-16. Static Ports Location

Figure 16-17. Pilot’s Static Air SourceValve Switch

Figure 16-15. Pitot Mast Location

On later models, the OAT gage (Figure 16-18)is located on the left sidewall panel below thepilot’s left arm. The probe is mounted belowthe pilot’s side window and directly oppositeof the gage. The ON–OFF button for the postlights is located next to the gage on the sidepanel.

On BB-1439, 1444 and subsequent, a digitaldisplay is located on the sidewall, and it indi-cates the free air temperature in Celsius. Whenthe adjacent button is depressed, Fahrenheit isdisplayed. The probe is located on the lowerfuselage under the pilot position (Figure 16-18).

AUTOFLIGHT SYSTEMYAW DAMPERA yaw damper function aids the pilot in main-taining directional control of the airplane. Thefunction may be used at any altitude; how-ever, it is required for flight above 17,000feet. Yaw damping should be deactivated fortakeoff and landing.

If the airplane has an autopilot system, theoperation of the yaw damper is covered inthe applicable Flight Manual supplement






Figure 16-18. Typical OAT Gage and Probe

(vendor’s manual). If an autopilot system isnot installed, yaw damping functions as anindependent system. The components of thissystem consist of a yaw sensor, amplifier,and a control valve (Figure 16-19).

STALL WARNINGSYSTEM

The formation of ice at the trans-ducer vane results in erroneous in-dications during flight.

The stall warning system consists of a trans-ducer, a lift computer, a warning horn, and a testswitch. Angle of attack is sensed by air pressureon the transducer vane (Figure 16-20) locatedon the left wing leading edge. When a stall isimminent, the transducer output is sent to a liftcomputer which activates a stall warning hornat approximately 5 to 13 knots above stall withflaps retracted, and at 5 to 12 knots above stallwith flaps in the 40% position, and at 8 to 14knots above stall with flaps fully extended.

The left main-gear squat switch disconnectsthe stall warning system when the aircraft ison the ground.

The system has preflight test capability throughthe use of the STALL WARN TEST switch(Figure 16-21) on the copilot’s left subpanel.This switch, held in the TEST position, raisesthe transducer vane, which actuates the warn-ing horn for preflight test purposes.

In the ICE group located on the pilot’s rightsubpanel, a STALL WARN switch (Figure16-22) controls electrical heating of the trans-ducer vane and mounting plate.

WARNING




YAWENG

APENG

SR

I/20

DN

L

UP

R

YAW ALT

YAW

HDG

COLLINS



Figure 16-19. YAW Damp Switch

Figure 16-21. STALL WARN TEST Switch(Copilot’s Left Subpanel)

Figure 16-20. Stall WarningTransducer Vane

COMMUNICATIONSYSTEM

STATIC DISCHARGINGDESCRIPTIONA static electrical charge, commonly referredto as “P” (precipitation static), builds up onthe surface of an airplane while in flight andcauses interference in radio and avionicsequipment operation. The charge is also dan-gerous to persons disembarking after land-ing , a s we l l a s t o pe r sons pe r fo rmingmaintenance on the airplane. Fifteen staticwicks (Figure 16-23) are installed on the trail-ing edges of the flight surfaces and the wingtips. The wicks aid in the dissipation of theelectrical charge. Nineteen are installed andonly three may be broken or missing.




Figure 16-22. STALL WARN Heat Switch(Pilot’s Right Subpanel)

Figure 16-23. Static Wicks

LIMITATIONS

AIRSPEED INDICATORRefer to Table 16-6 for airspeed indicatorlimitations.

OUTSIDE AIR TEMPERATUREGAGEDo not operate the airplane when the outsideair temperature is beyond the following limits:

• Minimum limit at all altitudes is –53.9°C(–65.02°F) for Super King Air 200 and–60°C (–76°F) for Super King Air B200.

• Maximum limit as follows:

1. Sea level to 25,000 feet—ISA +37°C

2. Above 25,000 feet—ISA + 31°C

AUTOPILOTRefer to the applicable FAA-approved FlightManua l supp l emen t f o r FAR Pa r t 91Operational Limitations for the autopilot.Except for minimum altitude, refer to the samesupplement for limitations imposed by FARPart 135, Operations, which establishes thesetwo limitations as well:

1. Enroute—500 feet above terrain is mini-mum altitude.

2. Coupled Approach—Observe decisionheight (DH) or minimum descent altitude(MDA).




MARKING KCAS VALUE KIAS VALUE SIGNIFICANCE OR RANGE OR RANGERED LINE 91 86 Air Minimum Control Speed (VMCA)

WHITE ARC 80 to 144 75 to 146 Full-flap Operating Range †80 to 155 †75 to 157

WIDE WHITE ARC 80 to 102 75 to 99 Lower limit is the stalling speed at (VSO) maximum weight with full flaps (ALL AIRPLANES) (100%) and idle power.

NARROW WHITE ARC 102 to 144 99 to 146 Lower limit is the stalling speed (VS) at †102 to 155 †99 to 157 maximum weight with Flaps Up (0%) and idle power. Upper limit is the maximum speed permissible with flaps extended beyond approach (more than 40%).

WHITE TRIANGLE 200 200 Maximum flaps to/at approach (40%) (ALL AIRPLANES) speed.

BLUE LINE (ALL AIRPLANES) 122 121 One engine-inoperative best rate of climb speed.

RED & WHITE HASH- ‡270 KCAS (269 KIAS) or MARKED POINTER value equal to .48 Mach, whichever is lower. Maximum speed for any operation ¥260 KCAS (259 KIAS) or value equal to .48 Mach,

* The airspeed indicator is marked in CAS values for 200 aircraft. † Applicable to Super King Air B200. ‡ SN BB-2 through BB-198, except airplanes modified by Beechcraft Kit Number 101-5033-1 S in compliance with Beechcraft Service Instructions Number 0894. ¥ SN BB-199 and subsequent, BL-1 and subsequent, and any earlier airplanes modified by Beechcraft Kit Number 101-5033-1 S in compliance with Beechcraft Service Instructions Number 0894.

Table 16-6. AIRSPEED INDICATOR MARKINGS*

17-i

CHAPTER 17MISCELLANEOUS SYSTEMS

CONTENTS

Page

INTRODUCTION................................................................................................................. 17-1

OXYGEN SYSTEMS ........................................................................................................... 17-1

Manual Plug-In System ................................................................................................. 17-4

Autodeployment System................................................................................................ 17-4

TOILET ................................................................................................................................. 17-8

RELIEF TUBES.................................................................................................................... 17-8

QUESTIONS....................................................................................................................... 17-10




17-iii

ILLUSTRATIONS

Figure Title Page

17-1 Oxygen System Diagram (BB-1439, 1444 and Subsequent)................................. 17-2

17-2 Oxygen System Diagram (Prior to BB-1444, Except 1439).................................. 17-3

17-3 Oxygen Mask Stowed ............................................................................................ 17-4

17-4 O2 Mask Selector (BB-1439, 1444 and After—Puritan Bennett) ......................... 17-5

17-5 O2 Mask Selector (Prior to BB-1444, Except 1439) ............................................. 17-5

17-6 First Aid Mask Access Panel ................................................................................. 17-6

17-7 Oxygen System Push-Pull Handles (BB-1439, 1444 and Subsequent) ................ 17-6

17-8 Oxygen System Push-Pull Handles (Prior to BB-1439) ....................................... 17-6

17-9 Oxygen Bottle and Shutoff Valve ......................................................................... 17-6

17-10 Oxygen System Annunciators ............................................................................... 17-7

17-11 Passenger Oxygen Mask Deployed ....................................................................... 17-7

17-12 Oxygen Available with Partially Full Bottle .......................................................... 17-8

17-13 Toilet ...................................................................................................................... 17-8

17-14 Relief Tube............................................................................................................. 17-9

TABLES

Table Title Page

17-1 Average Time of Useful Consciousness................................................................. 17-5

17-2 Oxygen Duration—200 and B200 ......................................................................... 17-9




INTRODUCTIONThe miscellaneous systems include the oxygen system, toilet, and the relief tubes.

OXYGEN SYSTEMSThe Super King Air has two oxygen systemsavailable:

1) A plug-in system for SNs BB-2 throughBB-54, and

2) An automatic deployment system for SNsBB-55 and subsequent, including the B200(Figures 17-1 and 17-2).

On the Super King Air 200, these systems arebased on an adequate flow for an altitude of31,000 feet. The Oxygen Duration Chart in theFlight Manual and the masks are based on 3.7

SLPM (Standard Liters Per Minute). The di-luter-demand crew mask is the only excep-t ion when used in the 100% mode . Forcomputation purposes, each diluter-demandcrew mask being used in the 100% mode countsas two masks at 3.7 SLPM.

On the Super King Air B200, the oxygen sys-tems are based on an adequate flow for an al-titude of 35,000 feet. The duration chart andmasks are based on a flow rate of 3.9 LPM-NTPD (Liters Per Minute-Normal TemperaturePressure Differential). The diluter demandcrew masks are an exception also, and com-

CHAPTER 17MISCELLANEOUS SYSTEMS







FORWARD PRESSURE BULKHEADCOCKPIT OXYGENPRESSURE GAGE

DILUTER DEMANDCREW MASK

PASSENGER MANUALOVERRIDE HANDLE

CABLE

PASSENGER MANUALOVERRIDE SHUTOFFVALVE

SOLENOID

OFF

ON

BAROMETRICPRESSURESWITCH

CONTROLCABLE

OXYGEN PRESSURESENSE SWITCH

PASSENGER SINGLE MASK OUTLET

FIRST AID OXYGEN MASK STOWEDIN MANUALLY OPERATED BOX

CONTROL CABLE

PRESSURE REGULATORAND SHUTOFF VALVE

COMPOSITE OXYGEN CYLINDER

HIGH PRESSURE OVERBOARD RELIEF

AFT PRESSURE BULKHEAD

OPTIONAL OXYGEN MASKCONTAINER, LINES AND

OUTLET FOR FOLD-UP SEATS

PASSENGER 2 MASK OUTLET(TYPICAL 5 PLACES)

OXYGEN PRESSURE GAGE FILL VALVE

ANNUNCIATOR PASS OXYGEN ON

DILUTER DEMAND CREW MASK

CONSOLEPULL ON SYSTEMREADY CONTROL

DETAIL A

DETAIL B

DETAIL C

DETAIL D

B

AD

CHIGH PRESSURE LINE

LOW PRESSURE LINE

CONTROL CABLE

FLEXIBLE HOSE

LEGEND

Figure 17-1. Oxygen System Diagram (BB-1439, 1444 and Subsequent)




B

A

PRESSURE REGULATORAND SHUTOFF VALVE

CONTROL CABLE

FIRST AID OXYGEN MASK STOWEDIN MANUALLY OPERATED BOX

PASSENGER SINGLE OXYGEN MASK

OXYGEN PRESSURESENSE SWITCH

OXYGEN PRESSUREGAGE

FILL VALVE

DETAIL B

DETAIL ABAROMETRICPRESSURESWITCH

OFF

ON

PASSENGER MANUALOVERRIDE SHUTOFFVALVE

SOLENOID

OXYGEN OUTLET

PASSENGER MANUALOVERRIDE CONTROL

COCKPIT OXYGENPRESSURE GAGE

FORWARD PRESSURE BULKHEADANNUNCIATOR PASS OXYGEN ON

DILUTER DEMAND CREW MASK

PULL ON SYSTEMREADY CONTROL

OXYGEN OUTLET

PASSENGER 2 MASK OUTLET(TYPICAL 5 PL ACES)

OPTIONAL OXYGEN MASKCONTAINER, LINES AND

OUTLET FOR FOLD-UP SEATS

AFT PRESSURE BULKHEAD

HIGH PRESSURE OVERBOARD RELIEF

STEEL OXYGEN CYLINDER

IN

OUT

NOTICE: AVIATORS BREATHING OXYGENKEEP FILL AREA CLEAN, DRY & FREE FROM OILPRESSURE TO 1850 PSI @ 14.7 PSI & 70°F

OXYGEN

PRECHARGE

NOTE:BB-55-309, 311–342, 344–382, 384–414, 417 AND 419 HAVE ONLY THREE PASSENGER MASK OUTLETS WITH THREE LATERALLY-PLACED MASKS APIECE. (PRIOR TO BB-55, IT IS THE SAME WITH THE OPTIONAL AUTODEPLOYMENT SYSTEM.)

HIGH PRESSURE LINE

LOW PRESSURE LINE

CONTROL CABLE

FLEXIBLE HOSE

LEGEND

Figure 17-2. Oxygen System Diagram (Prior to BB-1444, Except 1439)

putation is identical with that used on theSuper King Air 200. At cabin altitudes above20,000 feet, the 100% mode is required.

MANUAL PLUG-IN SYSTEM

(Super King Air 200)The manual plug-in system (SNs BB-2 throughBB-54) is the constant-flow type with eachmask plug having its own regulating orifice.The crew oxygen masks are stowed under thepilot’s and copilot’s seats. Oxygen outlets arelocated on the forward cockpit sidewalls. Thepassenger masks are stowed in pockets be-hind the seat backs. However, with respect tothe couch, the masks are stowed underneath.The cabin outlets, located on the cabin head-liner at the top center at the forward and aftends of the cabin, are protected by accessdoors when not in use. Pushing the plug in

firmly and then turning clockwise one-quar-ter turn easily connects the masks. Reversingthis procedure unplugs the mask.

AUTODEPLOYMENT SYSTEMThe autodeployment system (Figures 17-1 and17-2) is available for all Super King Air air-planes after SN BB-54 and is factory installedon all Super King Air B200 airplanes.

The crew utilizes diluter-demand, quick-don-ning oxygen masks (Figure 17-3) which areheld in the overhead panel. (Prior to BB-1444,except 1439, they hang on the aft cockpit par-tition behind and outboard of the crew seats.)Since these masks deliver oxygen only upon in-halation, there is no oxygen loss when the masksare plugged in and the PULL ON–SYS READYhandle is pulled out. For BB-1439, 1444 and sub-sequent this is located to the left of the power





BB-1439, 1444 AND AFTER BB-1439, 1444 AND AFTER—PURITAN BENNETT

Figure 17-3. Oxygen Mask Stowed




quadrant. For prior aircraft it is located aft ofthe overhead lighting control panel.

The PULL ON–SYS READY han-dle shall be pulled out to arm theoxygen system prior to flight. Thisis mandatory since the oxygen bot-t le cable or l inkage may freeze.Should this cable or linkage freezewhen the handle is in the OFF po-sition (pushed in), the handle can-not be pulled out, and oxygen wouldnot be available.

Table 17-1 sets forth the average time of use-ful consciousness (time from onset on hy-poxia until loss of effective performance) atvarious altitudes.

On BB-1439, 1444 and subsequent, the crewmask has three modes of operation—normal,100%, and emergency (Figure 17-4). The nor-mal position mixes cockpit air with oxygensupplied through the mask. This mode reducesthe rate of oxygen depletion. When 100% isselected, only oxygen directly from the oxy-gen bottle is breathed by the crewmember.The emergency position supplies a positivepressure to the face piece and should be usedif smoke and/or fumes are present in the cabin.

Prior to BB-1444 except BB-1439, a smalllever (Figure 17-5) on each crew mask permitsthe selection of two operational modes—NORMAL and 100%. In the NORMAL posi-tion, cockpit air is mixed with the oxygensupplied through the mask. This mode reducesthe oxygen depletion, plus it is more com-fortable to use than 100% oxygen. However,when smoke or contaminated air is in the cock-pit, 100% oxygen must be used. The selectorlevers must be kept in the 100% position whenthe masks are stowed so that no adjustment isnecessary when the masks are donned.

When the primary oxygen supply l ine ischarged, oxygen can be obtained from the first

WARNING

Table 17-1. AVERAGE TIME OF USEFULCONSCIOUSNESS

35,000 feet ............................... 1/2 to 1 minute

30,000 feet ................................ 1 to 2 minutes

28,000 feet ........................... 2 1/2 to 3 minutes

25,000 feet ................................ 3 to 5 minutes

22,000 feet ............................... 5 to 10 minutes

12,000-18,000 feet ............. 30 minutes or more

Figure 17-4. O2 Mask Selector (BB-1439, 1444and After—Puritan Bennett)

Figure 17-5. O2 Mask Selector (Prior toBB-1444, Except 1439)

aid oxygen mask located in the toilet area.The first aid mask is actuated by manuallyopening the overhead access panel (Figure17-6) marked FIRST AID OXYGEN–PULLand opening the on-off valve inside the box.There is a placard which reads: NOTE: CREWSYS MUST BE ON to remind the user that thePULL ON–SYS READY handle in the cock-pit must be armed before oxygen flows throughthe first aid mask.

The PULL ON–SYS READY push-pull han-dle (Figure 17-7) is located to the left of thepower quadrant (Prior to BB-1444, except 1439it is aft of the overhead light control panel;Figure 17-8). The PASSENGER MANUALO’RIDE (override) push-pull handle (Figure17-7) is located on the right side of the powerquadrant (prior to BB-1444, except 1439 it is

next to the PULL ON–SYS READY handle inthe overhead panel; Figure 17-8). Both are op-erated the same way. Pushing in the handledeactivates the selected function, while pullingout the handle actuates the desired function.

The system ready handle operates a cable thatopens and closes the shutoff valve on the oxy-gen bottle (Figure 17-9) in the aft fuselage, be-hind the aft pressure bulkhead. When thehandle is pushed in, no oxygen supply is avail-able anywhere in the airplane. It must be pulledout before engine starting to ensure oxygen isavailable any time it is needed. If the oxygenbottle is not empty when the handle is pulledout, the primary oxygen supply line chargeswith oxygen. This supply line, when charged,delivers oxygen to the two-crew oxygen out-lets, to the first aid oxygen mask, and to themanual override shutoff valve. The crew canmonitor the pressure in the oxygen bottle byreading the oxygen gage on the copilot’s rightsubpanel. It should be noted the filler gage alsoreads bottle and system pressure.

The passenger oxygen system is the constantflow type. Any time the cabin-pressure altitudeexceeds approximately 12,500 feet, a baro-




Figure 17-6. First Aid Mask Access Panel

Figure 17-8. Oxygen System Push-PullHandles (Prior to BB-1439)

Figure 17-7. Oxygen System Push-PullHandles (BB-1439, 1444 and Subsequent)

Figure 17-9. Oxygen Bottle and ShutoffValve

metric pressure switch automatically energizesa solenoid causing passenger manual overrideshutoff valve to open. Also, the crew can man-ually open this valve any time by pulling out thePASSENGER MANUAL O’RIDE handle. Oncethe shutoff valve is opened, either automati-cally or manually, oxygen flows into the pas-senger oxygen supply line. When this happens,a pressure-sensitive switch in the supply linecauses the PASS OXY ON advisory annuncia-tor to illuminate (Figure 17-10). On SNs BB-310, -343, -383, -415, -416, -418, -448, -450 andsubsequent (including the B200 airplanes), andwith all serial numbers in the 1979 model year,this switch also causes the cabin lights (whichincludes all fluorescent lights, the vestibulelight, and the center baggage compartment light)to illuminate in the full bright mode.

This occurs regardless of the position of theCABIN LIGHTS switch on the copilot’s leftsubpanel.

Automatic deployment of the passenger con-stant-flow oxygen masks is accomplishedwhen the pressure of the oxygen in the supplyline causes a plunger to extend against eachof the mask dispenser doors, which forces thedoor open (Figure 17-11). When the doorsopen, the masks drop down approximatelynine inches below the doors.

NOTEThe lanyard valve pin at the top of theoxygen mask hose must be pulledout in order for oxygen to f lowthrough the mask.

A lanyard valve pin is connected to the maskwith a flexible cord. When the mask is pulleddown for use, the cord pulls the pin out of thelanyard valve. When this occurs, oxygen willflow continuously from the mask until thepassenger shutoff valve is closed. If the PAS-SENGER MANUAL O’RIDE handle is pushedin (and cabin altitude is below 12,500 feet),or the oxygen control circuit breaker in the en-vironmental group is pulled (regardless ofcabin altitude), this will isolate the remainingoxygen for the crew and first aid outlets. Referto Table 17-2 for the oxygen duration and seeFigure 17-12 for oxygen bottle capacity.

NOTE (200 AND B200)For duration time with crew usingdiluter-demand, quick-donning oxy-gen masks with selector on 100%,increase computation of NUMBEROF PEOPLE USING by a factor oftwo (e.g., with four passengers, enterthis table at eight).




AIR COND N 1 LOWLEC TRIM O FF

Figure 17-10. Oxygen SystemAnnunciators

Figure 17-11. Passenger Oxygen MaskDeployed

NOTE (200)Oxygen duration is computed for aPuri tan-Zep oxygen system thatuses either the red, color-coded,plug-in-type or the autodeployed-t y p e m a s k , b o t h r a t e d a t 3 . 7Standard Liters Per Minute (SLPM)flow. Both are approved for alti-tudes up to 31,000 feet.

NOTE (B200)Oxygen duration is computed for anautodeployed-type mask, 3.9 LitersPer Minute (LPM-NTPD), color-coded orange and white, and ap-proved for altitudes up to 35,000feet.

TOILETThe side-facing toilet (Figure 17-13) is in-stalled in the foyer and faces the airstair door.The foyer can be closed off from the cabin bysliding the two partition-type door panels to thecenter of the fuselage, where they are heldclosed by magnetic strips. The forward-facingtoilet, when installed, is located in the aft cargo

area and is enclosed by the cargo partition. Thetoilet may be either the chemical type or theelectrically-flushing type. In either case, thetwo-hinged lid half-sections must be raised togain access to the toilet. A toilet tissue dis-penser is contained in a slide out compartmenton the forward side of the toilet cabinet.

If a Monogram electrically-flushingtoilet is installed, the sliding knifevalve should be open at all times,except when actually servicing theunit. The cabinet below the toiletmust be opened in order to gain ac-cess to the knife valve actuator han-dle.

RELIEF TUBESA relief tube (Figure 17-14) is contained in a spe-cial tilt-out compartment at the aft side of thetoilet cabinet. A relief tube may also be in-stalled in the cockpit and stowed under the pilotor copilot seat. The hose on the cockpit relieftube is of sufficient length to permit use by ei-ther pilot or copilot.

CAUTION




Figure 17-12. Oxygen Available withPartially Full Bottle Figure 17-13. Toilet




A valve lever is located on the side of the relieftube horn. This valve lever must be depressedat all times while the relief tube is in use. Eachtube drains into the atmosphere through its ownspecial drain port, which protrudes from thebottom of the fuselage. Each drain port atom-izes the discharge to keep it away from the skinof the airplane.

NOTE The relief tubes are for use duringflight only.

OXYGEN DURATION—200CYLINDER NUMBER OF PEOPLE USINGVOLUME 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15CU FT DURATION IN MINUTES

22 150 72 48 36 30 24 21 18 16 15 13 12 11 10 *

49 336 168 108 84 66 54 48 42 37 33 30 27 25 24 22

64 438 216 144 108 84 72 60 54 48 43 39 36 33 31 28

76 552 261 173 130 104 87 74 66 57 52 47 43 40 37 34

115 792 396 264 198 158 132 113 99 88 79 72 66 60 56 52

OXYGEN DURATION—B200CYLINDER NUMBER OF PEOPLE USINGVOLUME 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17CU FT DURATION IN MINUTES22 143 71 47 35 28 23 20 17 15 14 13 11 11 10 * * *49 320 160 106 80 64 53 45 40 35 32 29 26 24 22 21 20 1866 431 215 143 107 86 71 61 53 47 43 39 35 33 30 28 26 2576 496 248 165 124 99 82 70 62 55 49 45 41 38 35 33 31 29115 751 375 250 187 150 125 107 93 83 75 68 62 57 53 50 46 44

OXYGEN DURATION—BB-1439, 1444 and SubsequentCYLINDER NUMBER OF PEOPLE USINGVOLUME 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 **16 **17CU FT DURATION IN MINUTES22 144 72 48 36 28 24 20 18 16 14 13 12 11 10 * * *50 317 158 105 79 63 52 45 39 35 31 28 26 24 22 21 19 1877 488 244 162 122 97 81 69 61 54 48 44 40 37 34 32 30 28115 732 366 244 183 146 122 104 91 81 73 66 61 56 52 48 45 43

Table 17-2. OXYGEN DURATION—200 AND B200

* Will not meet oxygen requirements.

** For oxygen duration computations, count each diluter-demand crew mask in use as 2(e.g. with 4 passengers and a crew of 2, enter the table at 8 people using).

Figure 17-14. Relief Tube

GEN-i


CONTENTS

Page

FLIGHT MANEUVERS AND PROFILES...................................................................... GEN-1

Takeoff ....................................................................................................................... GEN-1

FLIGHT PROFILES ......................................................................................................... GEN-1

LANDING ...................................................................................................................... GEN-18

Flaps-Up Approach and Landing ............................................................................ GEN-18

Single-Engine Approach and Landing .................................................................... GEN-18

Crosswind Approach and Landing .......................................................................... GEN-18

WINDSHEAR................................................................................................................. GEN-18

General .................................................................................................................... GEN-18

Microbursts.............................................................................................................. GEN-19

Acceptable Performance Guides ............................................................................. GEN-19

COCKPIT RESOURCE MANAGEMENT.................................................................... GEN-20




GEN-iii

ILLUSTRATIONS

Figure Title Page

GEN-1 Normal Takeoff and Departure................................................................... GEN-2

GEN-2 Engine Loss at or Above V1....................................................................... GEN-3

GEN-3 Rejected Takeoff......................................................................................... GEN-4

GEN-4 Steep Turns ................................................................................................. GEN-5

GEN-5 Approach to Stall—Clean........................................................................... GEN-6

GEN-6 Approach to Stall—Takeoff Configuration ................................................ GEN-7

GEN-7 Approach to Stall—Landing Configuration ............................................... GEN-8

GEN-8 Emergency Descent .................................................................................... GEN-9

GEN-9 Standard Holding Pattern—Direct Entry ................................................. GEN-10

GEN-10 Standard Holding Pattern—Teardrop Entry ............................................. GEN-11

GEN-11 Standard Holding Pattern—Parallel Entry ............................................... GEN-12

GEN-12 Visual Approach and Landing.................................................................. GEN-13

GEN-13 One Engine Inoperative—Visual Approach and Landing........................ GEN-14

GEN-14 ILS Approach—Landing in Sequence from an ILS................................. GEN-15

GEN-15 Non-Precision Approach—Procedure Turn ............................................. GEN-16

GEN-16 Circling Approach and Landing ............................................................... GEN-17

GEN-17 Situational Awareness in the Cockpit....................................................... GEN-20

GEN-18 Command and Leadership........................................................................ GEN-20

GEN-19 Communication Process ........................................................................... GEN-21

GEN-20 Decision-Making Process......................................................................... GEN-21




FLIGHT MANEUVERSAND PROFILES

TAKEOFF

Crosswind TakeoffFollow procedures for normal takeoff except:

• Hold aileron into wind.

• Maintain runway heading with rudderuntil rotation then crab to hold center line.

Instrument TakeoffFollow procedures for normal takeoff except:

• Transition to flight instruments at or be-fore 100 feet AGL.

Obstacle Clearance TakeoffFollow procedures for normal takeoff except:

• Maintain V2 until clear of obstacle.

FLIGHT PROFILESSpecific flight profiles are graphically de-picted on the following pages.


GEN-1FOR TRAINING PURPOSES ONLY



GEN-2 FOR TRAINING PURPOSES ONLY



TAKEOFF

IN POSITION

BEFORE TAKEOFF

VYSE OR ABOVE

CLIMB-OUT

1. CHECKLIST — COMPLETE2. RECHECK V1 AND V2

1. HOLD BRAKES2. PROPS — 2,000 RPM (ON GOVERNORS)3. RELEASE BRAKES4. SET TORQUE

TAKEOFF ROLL1. RECHECK TORQUE/ITT2. ANNUNCIATORS — CHECK

1. ROTATE AT V1 TO APPROX 7˚ NOSE UP2. ESTABLISH POSITIVE RATE OF CLIMB3. LANDING GEAR — UP

1. FLAPS — UP2. YAW DAMPER — ON3. CLIMB POWER — SET

1. ACCELERATE TO 160 KIAS2. LANDING/TAXI LIGHTS — OUT3. COMPLETE CLIMB CHECKLIST

AREA DEPARTURE/CLIMBPROFILE

1. 160 KIAS TO 10,000 FT2. 140 KIAS 10,000 - 20,000 FT3. 130 KIAS 20,000 - 25,000 FT4. 120 KIAS 25,000 - 35,000 FT

CRUISE1. ACCELERATE TO CRUISE SPEED2. SET CRUISE POWER3. COMPLETE CRUISE CHECKLIST

Figure GEN-1. Normal Takeoff and Departure




V2

TAKEOFF

BEFORE TAKEOFF

1,000 FT AGL

ENGINE LOSS

1. FOLLOW NORMAL TAKEOFF PROCEDURES UNTIL AT OR ABOVE V1

1. ROTATE AT V1 TO APPROX 7˚ NOSE UP2. ESTABLISH POSITIVE RATE OF CLIMB3. LANDING GEAR — UP

1. MAINTAIN RUNWAY HEADING

1. CHECK MAX POWER2. AIRSPEED AT V2

3. VERIFY PROP FEATHERED

1. VYSE (BLUE LINE)2. FLAPS — UP

CLIMB

1. COMPLETE ENGINE FAILURE CHECKLIST CLEAN-UP ITEMS2. LAND AS SOON AS PRACTICAL

NOTE: IT MAY BE NECESSARY TO BANK AS MUCH AS 5˚INTO THE GOOD ENGINE TO MAINTAIN RUNWAYHEADING. IT WILL TAKE ALMOST FULL RUDDERON THE SIDE OF THE GOOD ENGINE TO KEEPTHE BALL SLIGHTLY OFF CENTER.

NOTE:

DO NOT RETARD FAILED ENGINE POWER LEVERUNTIL THE AUTOFEATHER SYSTEM HAS COMPLETELYSTOPPED PROPELLER ROTATION.

Figure GEN-2. Engine Loss at or Above V1




BEFORE TAKEOFF

EMERGENCY OR MALFUNCTIONAT OR BELOW V1

CLEAR OF RUNWAY

1. FOLLOW NORMAL TAKEOFF PROCEDURES UNTIL INITIATING ABORT AT OR BELOW V1

1. RECOGNIZE REASON FOR REJECTING TAKEOFF2. POWER LEVERS — IDLE3. BRAKING — AS NECESSARY4. REVERSE — AS NECESSARY5. MAINTAIN RUNWAY HEADING

1. COMPLETE AFTER LANDING CHECKLIST

NOTE:

IF REJECTED TAKEOFF IS DUE TO REASONS OTHER THAN ONE ENGINE POWER LOSS, REVERSE IS MOST EFFECTIVE AT HIGH SPEEDS; BRAKING IS MOST EFFECTIVE AT LOW SPEEDS.

Figure GEN-3. Rejected Takeoff




ROLLOUT

1. RETURN TO AND HOLD ENTRY PARAMETERS

THROUGH 30° BANK

1. ADD APPROX 100 LBS TORQUE

2. ONE UNIT NOSEUP TRIM

3. SMALL PITCH INCREASE

THROUGH 30° BANK

1. REDUCE TORQUE 100 LBS

2. REDUCE PITCH

3. TAKE OUT TRIMROLL INTO TURN

1. MAINTAIN INITIAL ALTITUDE

INITIAL ENTRY

1. AIRSPEED — 180 KNOTS

2. TORQUE — APPROX 1,000-1,200 LBS

3. HEADING BUG — SET

4. FD — OFF

5. CHECK ADI PITCH REFERENCE

HOLD 45° BANK

1. SMALL PITCH CORRECTIONS

2. MAINTAIN AIRSPEED

ROLL OUT OF TURN

1. START ROLLOUT 25° PRIOR TO ROLLOUT HEADING

Figure GEN-4. Steep Turns




BEGINNING OF MANEUVER STALL AND RECOVERY COMPLETION OF MANEUVER

HORNOR BUFFET

INITIAL CONDITION:

1. TORQUE — 200 LBS2. PROPELLERS — 1,700 RPM3. MAINTAIN INITIAL HEADING4. MAINTAIN INITIAL ALTITUDE5. PITCH ATTITUDE PRIOR TO HORN OR BUFFET MAY REACH 10˚-15˚, DEPENDING ON TECHNIQUE6. HORN WILL SOUND APPROX 10 KTS ABOVE BUFFET

AT HORN OR BUFFET — RECOVER:

1. SIMULTANEOUSLY ADVANCE THE POWER LEVERS TOWARD MAX TORQUE, REDUCE THE PITCH ATTITUDE AS NECESSARY TO STOP THE STALL WARNING, AND ROLL THE WINGS LEVEL2. ESTABLISH POSITIVE RATE OF CLIMB

COMPLETION:

1. LEVEL OFF AT NEW ALTITUDE AND INITIAL HEADING2. RESET POWER AS REQUIRED

V2

Figure GEN-5. Approach to Stall—Clean





HORNOR BUFFET


1. REDUCE THE PITCH ATTITUDE AS NECESSARY TO STOP THE STALL WARNING, AND ROLL THE WINGS LEVEL2. ESTABLISH POSITIVE RATE OF CLIMB3. FLAPS — UP, AT OR ABOVE VYSE (BLUE LINE)

COMPLETION:


V2

INITIAL CONDITION:

1. TORQUE — 200 LBS 2. PROPELLERS — 2,000 RPM 3. MAINTAIN INTITIAL HEADING 4. MAINTAIN INTITIAL ALTITUDE 5. FLAPS — APPROACH (BELOW TRIANGLE) 6. AT 110 KIAS OR LESS, SIMULTANEOUSLY SET THE TORQUE TO 1,100 LBS (SIMULATED 100% TORQUE), ESTABLISH A BANK ANGLE OF 20˚ (NO MORE THAN 30˚), AND RAISE THE NOSE AND CLIMB 7. STUDENT MAY BE REQUIRED TO PERFORM THIS MANEUVER WHILE MAINTAINING 15˚ - 30˚ ANGLE OF BANK OR WHILE MAINTAINING A HEADING 8. CLEAR AREA IN DIRECTION OF TURN 9. DECREASE SPEED APPROX 1 KT PER SECOND10. PITCH ATTITUDE PRIOR TO HORN OR BUFFET MAY REACH 15˚ - 25˚, DEPENDING ON TECHNIQUE

Figure GEN-6. Approach to Stall—Takeoff Configuration





HORNOR BUFFET


1. SIMULTANEOUSLY ADVANCE THE POWER LEVERS TOWARD MAX TORQUE, PROPELLER LEVERS FULL FORWARD, REDUCE THE PITCH ATTITUDE AS NECESSARY TO STOP THE STALL WARNING, AND ROLL THE WINGS LEVEL2. ESTABLISH POSITIVE RATE OF CLIMB3. FLAPS — UP, AT OR ABOVE 100 KIAS4. GEAR — UP

COMPLETION:


V2

INITIAL CONDITION:

1. TORQUE — 200 LBS2. PROPELLERS — 1,700 RPM3. MAINTAIN INTITIAL HEADING4. MAINTAIN INTITIAL ALTITUDE5. FLAPS — APPROACH (BELOW TRIANGLE)6. GEAR — DOWN (BELOW VLE)7. FLAPS — DOWN 100% (BELOW TOP OF WHITE ARC)8. PITCH ATTITUDE PRIOR TO HORN OR BUFFET MAY REACH 10˚ - 15˚, DEPENDING ON TECHNIQUE9. HORN WILL SOUND APPROX 10 KTS ABOVE BUFFET

Figure GEN-7. Approach to Stall—Landing Configuration




INITIAL

1. OXYGEN SYSTEM — VERIFY ARMED

2. CREW MASK — ON

3. MIC SWITCH — OXYGEN MASK POSITION

4. SPEAKER (AS REQUIRED)

5. PASSENGER OXYGEN (AS REQUIRED)

6. POWER LEVERS — IDLE

7. PROP LEVERS — SMOOTHLY FULL FORWARD

8. FLAPS — APPROACH (BELOW TRIANGLE)

9. GEAR — DOWN (BELOW VLE)

NOTE: IF INITIAL INDICATED AIRSPEED IS

ABOVE VLE, MAINTAIN THE INITIAL ALTITUDE UNTIL THE IAS IS AT OR BELOW VLE.

DESCENT

1. INITIAL PITCH ATTITUDE — 20° NOSEDOWN

2. PRIOR TO VLE, REDUCE PITCH ATTITUDE TO APPROXIMATELY 14° NOSEDOWN

3. MAXIMUM IAS SHOULD BE VLE

4. ADVISE ATC

5. RESET ALTIMETER AND ALTITUDE ALERTER TO LEVEL-OFF ALTITUDE

NOTE: DESCENT FROM 35,000 TO 12,500

FEET REQUIRES APPROXIMATELY SIX MINUTES

LEVEL-OFF

1. APPROXIMATELY 500 FEET BEFORE LEVEL-OFF ALTITUDE, SMOOTHLY REDUCE RATE OF DESCENT

2. FLAPS — UP

3. GEAR — UP (BELOW VLO RETRACTION)

4. ADD POWER AS REQUIRED

5. MIC SWITCH — NORMAL POSITION

6. REMOVE MASK

7. SET PROP RPM

8. COMPLETE DESCENT CHECKLIST

REDUCE RATE OF DESCENT APPROXIMATELY 500 FEET ABOVE LEVEL-OFF ALTITUDE

20° NOSEDOWN

VLE — APPROXIMATELY14° NOSEDOWN

LEVEL OFF

Figure GEN-8. Emergency Descent




TURN INBOUND

1. ADJUST LEG LENGTH TO PROVIDE 1 MINUTE AT 14,000 FEET AND BELOW OR 1.5 MINUTES ABOVE 14,000 FEET

70°

110°

ENTERING HOLDING PATTERN

1. REPORT ENTERING HOLD

2. TURN TO PARALLEL OUTBOUND COURSE

3. START TIMING OVER OR ABEAM FIX, WHICHEVER OCCURS LATER

*MAX HOLDING SPEEDS

• 6,000 FEET & BELOW — 200 KIAS

• 6,001-14,000 FEET — 230 KIAS

• 14,001 & ABOVE — 265 KIAS

INITIAL

1. SLOW TO HOLDING AIRSPEED* — 160 KIAS WITHIN 3 MINUTES OF FIX

2. TORQUE — APPROX 800-1,000 LBS

Figure GEN-9. Standard Holding Pattern—Direct Entry




TURN INBOUND


70°

*MAX HOLDING SPEEDS


• 6,001-14,000 FEET — 230 KIAS

• 14,001 & ABOVE — 265 KIAS



2. TURN 30° FROM OUTBOUND COURSE

3. START TIMING OVER FIX

INITIAL

1. SLOW TO HOLDING AIRSPEED* — 160 KIAS WITHIN 3 MINUTES OF FIX


Figure GEN-10. Standard Holding Pattern—Teardrop Entry




TURN INBOUND


110°

*MAX HOLDING SPEEDS


• 6,001-14,000 FEET — 230 KIAS

• 14,001 & ABOVE — 265 KIAS



2. TURN TO PARALLEL OUTBOUND COURSE

3. START TIMING OVER OR ABEAM FIX, WHICHEVER OCCURS LATER

INITIAL

1. SLOW TO HOLDING AIRSPEED — 160 KIAS* WITHIN 3 MINUTES OF FIX


Figure GEN-11. Standard Holding Pattern—Parallel Entry




INITIAL

1. OBTAIN ATIS2. DESCENT CHECKLIST — COMPLETE

ARRIVAL

1. TORQUE — APPROX 800 LBS2. 150 - 175 KIAS (TYPICAL)3. START BEFORE LANDING CHECKLIST

DOWNWIND

1. FLAPS — APPROACH2. 130 - 140 KIAS

ABEAM TOUCHDOWN POINT

1. GEAR — DOWN2. BEFORE LANDING CHECKLIST — COMPLETE

BASE

1. 130 KIAS (MIN REC)

FINAL

1. 130 - 140 KIAS (VYSE MIN)WHEN LANDING ASSURED: 2. FLAPS — DOWN 3. TRANSITION TO VREF 4. YAW DAMPER — OFF

LANDING

1. PROPS — FULL FORWARD 2. BETA OR REVERSE3. BRAKES — AS NECESSARY

THRESHOLD

1. GEAR — RECHECK DOWN2. AIRSPEED — VREF3. POWER — IDLE

REJECTED LANDING

1. POWER — MAX 2. PITCH — 10˚ NOSE UP3. AIRSPEED — 100 KIAS4. ESTABLISH NORMAL CLIMB WHEN CLEAR OF OBSTACLES5. FLAPS — UP6. GEAR — UP

CAUTION

TO ENSURE CONSTANT REVERSINGCHARACTERISTICS, THE PROPELLERCONTROL MUST BE IN FULL INCREASERPM POSITION.

NOTE: REVERSE IS MOST EFFECTIVE AT HIGHER SPEEDS; BRAKING IS MOST EFFECTIVE AT LOWER SPEEDS

CAUTION

IF POSSIBLE, PROPELLERS SHOULD BE MOVED OUT OFREVERSE AT APPROXIMATELY 40 KNOTS TO MINIMIZEBLADE EROSION. CARE MUST BE EXERCISED WHENREVERSING ON RUNWAYS WITH LOOSE SAND, DUST,OR SNOW ON THE SURFACE. FLYING GRAVEL WILLDAMAGE PROPELLER BLADES, AND DUST OR SNOWMAY IMPAIR THE PILOT'S VISIBILITY.

Figure GEN-12. Visual Approach and Landing




INITIAL

1. OBTAIN ATIS2. DESCENT CHECKLIST — COMPLETE

ARRIVAL

1. TORQUE — APPROX 1,600 LBS2. 150 - 175 KIAS (TYPICAL)3. START ONE-ENGINE-INOPERATIVE APPROACH AND LANDING CHECKLIST

DOWNWIND

1. FLAPS — APPROACH2. 130 - 140 KIAS

ABEAM TOUCHDOWN POINT

1. GEAR — DOWN2. PROP — FULL FORWARD

BASE

1. 130 KIAS (MIN REC)

FINAL

1. 130 - 140 KIAS (VYSE MIN)WHEN LANDING ASSURED: 2. FLAPS — DOWN 3. TRANSITION TO VREF 4. YAW DAMPER — OFF 5. ONE-ENGINE-INOPERATIVE APPROACH AND LANDING CHECKLIST — COMPLETE

LANDING

1. BETA OR REVERSE — AS NECESSARY2. BRAKES — AS NECESSARY

THRESHOLD


GO-AROUND

1. POWER — MAX 2. GEAR — UP3. FLAPS — UP4. AIRSPEED — INCREASE TO VYSE (BLUE LINE)

Figure GEN-13. One Engine Inoperative—Visual Approach and Landing




INITIAL

ARRIVAL

1. OBTAIN ATIS2. REVIEW APPROACH AND MISSED APPROACH3. NAVAIDS — TUNE/IDENT4. DESCENT CHECKLIST — COMPLETE

1. TORQUE — APPROX 800 LBS2. 150 - 175 KIAS (TYPICAL)3. FD — AS DESIRED4. START BEFORE LANDING CHECKLIST

APPROACH INBOUND1. FLAPS — APPROACH2. 130 - 140 KIAS

APPROACHING GLIDE SLOPE1. GEAR — DOWN2. COMPLETE BEFORE LANDING CHECKLIST

DH-VISUAL AND LANDING ASSURED

1. FLAPS — DOWN2. TRANSITION TO VREF3. YAW DAMPER — OFF

DH

MM

OM

THRESHOLD


LANDING1. PROPS — FULL FORWARD2. BETA OR REVERSE3. BRAKES — AS NECESSARY

DH-MISSED APPROACH

1. POWER — MAX 2. PITCH — 7˚ - 8˚ NOSE UP (FD-GA)3. FLAPS — UP4. GEAR — UP5. COMPLETE MISSED APPROACH PROCEDURE

GLIDE SLOPE INTERCEPT

1. TORQUE — APPROX 600 - 800 LBS2. 130 - 140 KIAS (VYSE MIN)

CAUTION

TO ENSURE CONSTANT REVERSING CHARACTERISTICS,THE PROPELLER CONTROL MUST BE IN FULL INCREASERPM POSITION.


CAUTION

IF POSSIBLE, PROPELLERS SHOULD BE MOVED OUT OFREVERSE AT APPROXIMATELY 40 KNOTS TO MINIMIZEBLADE EROSION. CARE MUST BE EXERCISED WHENREVERSING ON RUNWAYS WITH LOOSE SAND, DUST,OR SNOW ON THE SURFACE. FLYING GRAVEL WILLDAMAGE PROPELLER BLADES, AND DUST OR SNOWMAY IMPAIR THE PILOT'S VISIBILITY.

Figure GEN-14. ILS Approach—Landing in Sequence from an ILS




FAF

FAF

INITIAL

1. OBTAIN ATIS2. REVIEW APPROACH AND MISSED APPROACH3. NAVAIDS — TUNE/IDENT4. DESCENT CHECKLIST — COMPLETE

ARRIVAL

1. TORQUE — APPROX 800 LBS2. 150 - 175 KIAS (TYPICAL)3. FD — AS DESIRED4. START BEFORE LANDING CHECKLIST

STATION PASSAGE

1. START TIMING2. SET ALTITUDE ALERTER

1. START TIMING2. FLAPS — APPROACH3. 130 - 140 KIAS

PROCEDURE TURN OUTBOUND

1. FD — AS DESIRED2. RESET ALTITUDE ALERTER

PROCEDURE TURN INBOUND

INTERCEPT FINAL APPROACH

1. COURSE INBOUND

APPROACH INBOUND

1. RESET ALTITUDE ALERTER

FINAL APPROACH FIX

1. START TIMING2. GEAR — DOWN3. TORQUE — APPROX 200 LBS4. COMPLETE BEFORE LANDING CHECKLIST5. 130 - 140 KIAS

MINIMUM DESCENT ALTITUDE (MDA)

1. LEVEL OFF AT MDA AT LEAST 1 MILE PRIOR TO MAP, IF POSSIBLE2. TORQUE — 1,100 - 1,300 LBS3. 130 - 140 KIAS (VYSE MIN)

MDA

MAP

MAP-MISSED APPROACH

1. POWER — MAX 2. PITCH — 7˚ - 8˚ NOSE UP (FD-GA)3. FLAPS — UP4. GEAR — UP 5. COMPLETE MISSED APPROACH PROCEDURE

MAP-LANDING ASSURED

1. FLAPS — DOWN2. TRANSITION TO VREF3. YAW DAMPER — OFF

THRESHOLD


LANDING

1. PROPS — FULL FORWARD2. BETA OR REVERSE3. BRAKES — AS NECESSARY

CAUTION

TO ENSURE CONSTANT REVERSING CHARACTERISTICSCHARACTERISTICS, THE PROPELLER CONTROL MUSTBE IN FULL INCREASE RPM POSITION.


CAUTION

IF POSSIBLE, PROPELLERS SHOULD BE MOVED OUT OFREVERSE AT APPROXIMATELY 40 KNOTS TO MINIMIZEBLADE EROSION. CARE MUST BE EXERCISED WHENREVERSING ON RUNWAYS WITH LOOSE SAND, DUST, ORSNOW ON THE SURFACE. FLYING GRAVEL WILL DAMAGEPROPELLER BLADES, AND DUST OR SNOW MAY IMPAIRTHE PILOT'S VISIBILITY.

Figure GEN-15. Non-Precision Approach—Procedure Turn




CAUTION

TO ENSURE CONSTANT REVERSING CHARACTERISTICS,THE PROPELLER CONTROL MUST BE IN FULL INCREASERPM POSITION.


CAUTION

IF POSSIBLE, PROPELLERS SHOULD BE MOVED OUT OFREVERSE AT APPROXIMATELY 40 KNOTS TO MINIMIZEBLADE EROSION. CARE MUST BE EXERCISED WHENREVERSING ON RUNWAYS WITH LOOSE SAND, DUST, ORSNOW ON THE SURFACE. FLYING GRAVEL WILL DAMAGEPROPELLER BLADES, AND DUST OR SNOW MAY IMPAIRTHE PILOT'S VISIBILITY.

1 NM

MDAMAP

ARRIVAL

1. PLAN CIRCLING MANEUVER2. FOLLOW NORMAL APPROACH PROCEDURES TO MDA

MINIMUM DESCENT ALTITUDE (MDA)

1. LEVEL OFF AT MDA AT LEAST 1 MILE PRIOR TO MAP, IF POSSIBLE2. TORQUE — 1,100 - 1,300 LBS3. 130 - 140 KIAS (VYSE MIN)4. MANEUVER WITHIN VISIBILITY CRITERIA5. MAINTAIN MDA

MAP AND DURING CIRCLING MANEUVER

1. DETERMINE THAT VISUAL CONTACT WITH THE RUNWAY ENVIRONMENT CAN BE MAINTAINED AND A NORMAL LANDING CAN BE MADE FROM A CIRCLING APPROACH, OR INITIATE A MISSED APPROACH2. MAINTAIN MDA DURING CIRCLING MANEUVER

BASE

1. COMMENCE DESCENT FROM A POINT WHERE A NORMAL LANDING CAN BE MADE

1. 130 - 140 KIAS (VYSE MIN)WHEN LANDING ASSURED: 2. FLAPS — DOWN 3. TRANSITION TO VREF 4. YAW DAMPER — OFF

FINAL

THRESHOLD


NOTE: THIS IS A CATEGORY B AIRCRAFT, BUT AIRSPEEDS OF 121 THROUGH 140 KIAS REQUIRE USING CATEGORY C MINIMUMS.

Figure GEN-16. Circling Approach and Landing




LANDING

FLAPS-UP APPROACH ANDLANDING Follow normal approach and landing proce-dures except:

• Complete the flaps up landing checklist.

• Refer to the flaps up VREF.

• Airspeed 140 knots until established onfinal.

• When landing assured—reduce the air-speed to the flaps up VREF.

SINGLE-ENGINE APPROACHAND LANDINGFollow normal approach and landing proce-dures except:

• Complete the one-engine-inoperativeapproach and landing checklist.

• The target torque settings are approxi-mately doubled.

• Smoothly push the propeller lever fullforward (2,000 rpm) prior to the IAF ordownwind.

• Maintain the airspeed at least 10 knotsabove VREF until landing assured.

• Cautiously use reverse, if necessary.

• If performance is limited when accom-plishing a circling approach, circle withthe flaps positioned for approach andthe gear up until it is certain the field canbe reached with the gear down.

CROSSWIND APPROACH ANDLANDINGFollow normal approach and landing proce-dures except:

• Crab into the wind to maintain the de-sired track across the ground.

• Immediately prior to touchdown, lowerthe upwind wing by use of the aileronand align the fuselage with the runwayby use of the rudder. During the rollout,hold the aileron control into the wind andmaintain directional control with therudder and brakes.

WINDSHEAR

GENERALThe best windshear procedure is avoidance.Recognize the indications of potential wind-shear and then:

AVOID AVOID AVOID

The key to recovery from windshear is to flythe aircraft so it is capable of a climb gradientgreater than the windshear-induced loss of per-formance. Normally, the standard wind/gustcorrection factor 1/2 gust will provide a suf-ficient margin of climb performance. If a shearis encountered that jeopardizes safety, initiatea rejected landing procedure. If the sink rateis arrested, continue with the procedure formicrobursts.

MICROBURSTSIf a microburst is encountered, the first indi-cation will be a rapid increase in the rate ofdescent accompanied by a rapid drop belowglide path (visual or electronic).

1. Initiate normal rejected landing pro-cedures (10° pitch).

2. Do not change the aircraft configurationuntil a climb is established.

3. If the aircraft is not climbing, smoothlyincrease pitch until a climb is estab-lished or stall warning is encountered.If stall warning is encountered, de-crease pitch sufficiently to depart thestall warning regime.

4. When positively climbing at a safe al-titude, complete the rejected landingmaneuver.

NOTEThe positive rate of climb should beverified on at least two (2) instru-ments. Leave the gear down untilyou have this climb indication, as itwill absorb some energy on impactshould the microburst exceed yourcapability to climb.

If a decision is made to rotate to thestall warning, extreme care should beexercised so as not to over rotate be-yond that point as the aircraft is onlya small percentage above the stallwhen the aural warning activates.

ACCEPTABLE PERFORMANCEGUIDELINES

• Understand that avoidance is primary.

• Ability to recognize potential windshearsituations.

• Ability to fly the aircraft to obtain op-timum performance.

WARNING




COCKPIT RESOURCE MANAGEMENT




REMEMBER

2 + 2 = 2

— OR —

2 + 2 = 5(SYNERGY)

IT'S UP TO YOU!CLUES TO IDENTIFYING:• Loss of Situational Awareness• Links in the Error Chain

HU

MA

NO

PE

RA

TIO

NA

L

1. FAILURE TO MEET TARGETS2. UNDOCUMENTED PROCEDURE3. DEPARTURE FROM SOP4. VIOLATING MINIMUMS OR LIMITATIONS5. NO ONE FLYING AIRPLANE6. NO ONE LOOKING OUT WINDOW7. COMMUNICATIONS8. AMBIGUITY9. UNRESOLVED DISCREPANCIES10. PREOCCUPATION OR DISTRACTION11. CONFUSION OR EMPTY FEELING12.

GROUPS/A

CAPTAININDIVIDUAL

S/A

COPILOTINDIVIDUAL

S/A

SITUATIONAL AWARENESS IN THE COCKPIT

LEADERSHIP STYLES

AUTOCRACTICSTYLE

(EXTREME)

AUTHORITARIANLEADERSHIP

STYLE

DEMOCRATICLEADERSHIP

STYLE

LAISSEZ-FAIRESTYLE

(EXTREME)

PARTICIPATION

LOW HIGHCOMMAND — Designated by Organization — Cannot be SharedLEADERSHIP — Shared Among Crewmembers — Focuses on "What's Right," not "Who's Right"

COMMAND AND LEADERSHIP

Figure GEN-18. Command and Leadership

Figure GEN-17. Situational Awareness in the Cockpit




EVALUATERESULT

IMPLEMENTRESPONSE

RECOGNIZENEED

IDENTIFYAND

DEFINEPROBLEM

COLLECTFACTS

IDENTIFYALTERNATIVES

WEIGH IMPACTOF ALTERNATIVES

SELECT ARESPONSE

HINTS:• Identify the problem: • Communicate it • Achieve agreement • Obtain commitment

• Consider appropriate SOPs

• Think beyond the obvious alternatives

• Make decisions as a result of the process

• Resist the temptation to make an immediate decision and then support it with facts.

DECISION MAKING PROCESS

Figure GEN-20. Decision-Making Process

FEEDBACK

NEED SEND RECEIVE

INTERNALBARRIERS

EXTERNALBARRIERS

INTERNALBARRIERS

OPERATIONALGOAL

THINK:• Solicit and give feedback• Listen carefully• Focus on behavior, not people• Maintain focus on the goal• Verify operational outcome is achieved

ADVOCACY: To increase other's S/A

• State Position

• Suggest Solutions

• Be Persistent and Focused

• Listen Carefully

INQUIRY: To increase your own S/A

• Decide What, Whom, How to ask

• Ask Clear, Concise Questions

• Relate Concerns Accurately

• Draw Conclusions from Valid Information

• Keep an Open Mind

— REMEMBER —Questions enhance communication flow. Do not give in to the temptation to ask questions when advocacy is required. Use of advocacy or inquiry should raise a "red flag."

COMMUNICATION PROCESS

Figure GEN-19. Communication Process

WA-1FOR TRAINING PURPOSES ONLY



WALKAROUNDThe following section is a pictorial walkaround. It shows each itemcalled out in the exterior power-off preflight inspection.

The general location photographs do not specify every checklistitem. However, each item is portrayed on the large-scale photographsthat follow.

WA-2 FOR TRAINING PURPOSES ONLY



11

9 1087 4

5

6

3

21 113 114

115

116112

106

109

110

111

21 2019

1815

16 14

13

12

17

30

44

45

54

50

47 484649 43

29

24 25

31 27

28 26 23 22

35

32

38

36

68

56

5151 52 33 3462 63 64

37 404139425770695958

8. INVERTER COOLING LOUVERS—CHECK

1. CABIN DOOR SEAL—CHECK

3. OIL BREATHER VENT—CHECK

2. FLAPS—CHECK

4. LEFT MAIN GEAR, STRUT, TIRES, BRAKES—CHECK5. CHOCK—REMOVE6. BRAKE DEICE LINE (IF INSTALLED)—CHECK




7. FIRE EXTINGUISHER PRESSURE (IF INSTALLED)—CHECK

WALKAROUND INSPECTION




9. AILERON, AILERON TAB, STATIC WICKS (4)—CHECKED

10. FLUSH OUTBOARD DRAIN—DRAIN

11. NAVIGATION, RECOGNITION, STROBE LIGHT—CHECKED

14. TIEDOWN—REMOVED

13. STALL WARNING VANE—CHECK

12. MAIN FUEL TANK CAP—SECURE

16. STALL STRIP—CHECK 19. WING LEADING EDGE TANK SUMP—DRAIN

17. ICE LIGHT—CHECK

15. OUTBOARD DEICE BOOTS—CHECKED 18. RAM SCOOP FUEL VENT AND HEATED FUEL VENT—CLEAR




20. GRAVITY LINE DRAIN—DRAIN




21. FUEL FILTER STRAINER DRAIN AND STANDBY PUMPDRAIN—DRAIN

22. LANDING GEAR DOORS—CHECK

23. WHEEL WELL—CHECK

25. ENGINE OIL CAP—SECURE

24. ENGINE OIL—CHECK

26. ENGINE COMPARTMENT DOOR (OUTBOARD)—SECURE, BLEED VALVE EXHAUST CLEAR

29. NACELLE COOLING RAM AIR INLETS—CLEAR

27. EXHAUST STACK (OUTBOARD)—CHECK FOR CRACKS28. TOP COWLING LOCKS (OUTBOARD)—SECURE

30. PROPELLER—CHECK FOR NICKS, DEICE BOOTSECURE

31. ENGINE AIR INTAKE—CLEAR




32. ENGINE COMPARTMENT DOOR (INBOARD)—SECURE,BLEED VALVE EXHAUST CLEAR

33. TOP COWLING LOCKS (INBOARD)—SECURE

34. EXHAUST STACK (INBOARD)—CHECK FOR CRACKS

35. GENERATOR COOLING INLET—CLEAR




36. AUXILIARY FUEL TANK CAP—SECURE

37. HYDRAULIC GEAR SERVICE DOOR—SECURE 40. HYDRAULIC LANDING GEAR VENT LINES—CLEAR

38. HEAT EXCHANGER (INLET & OUTLET)—CLEAR 41. AUXILIARY FUEL TANK SUMP—DRAIN

39. INBOARD DEICE BOOTS—CHECKED

42. LOWER ANTENNAS AND BEACON—CHECKED




43. OAT PROBE—CHECK44. AVIONICS PANEL—SECURE45. CONDENSER BLOWER OUTLET—CLEAR

46. NOSE GEAR, DOORS, STRUT, TIRE—CHECKED47. CHOCK—REMOVED

48. NOSE GEAR STEERING STOP BLOCK—CHECK

49. NOSE GEAR WHEEL WELL—CHECK

50. LANDING AND TAXI LIGHTS—CHECK




51. PITOT MASTS—COVER REMOVED, CLEAR

52. WINDSHIELD, WINDSHIELD WIPERS—CHECK53. RAM AIR INLET—CLEAR54. RADOME—CHECK55. AVIONICS PANEL—SECURE56. AUXILIARY FUEL TANK CAP—SECURE57. INBOARD DEICE BOOTS—CHECKED58. HEAT EXCHANGER (INLET & OUTLET)—CLEAR59. AUXILIARY FUEL TANK SUMP—DRAIN60. ENGINE OIL—CHECK61. ENGINE OIL CAP—SECURE62. ENGINE COMPARTMENT DOOR (INBOARD)—SECURE,

BLEED VALVE EXHAUST CLEAR

63. TOP COWLING LOCKS (INBOARD)—SECURE64. EXHAUST STACK (INBOARD)—CHECK FOR CRACKS65. NACELLE COOLING RAM AIR INLETS—CLEAR66. PROPELLER—CHECK FOR NICKS, DEICE BOOT

SECURE67. ENGINE AIR INTAKE—CLEAR

For Right Wing, A Close Up Is Given For Those Components Different Than Left Wing (See Foldout Page For Specific Locations)




69. BATTERY BOX DRAIN—CLEAR

70. BATTERY AIR INLET (NICKEL-CADMIUM)—CLEAR,VALVE FREE

71. ENGINE COMPARTMENT DOOR (OUTBOARD)—SECURE

72. EXHAUST STACK (OUTBOARD)—CHECK FOR CRACKS

73. TOP COWLING LOCKS (OUTBOARD)—SECURE

74. GENERATOR COOLING INLET—CLEAR

75. FUEL FILTER STRAINER DRAIN AND STANDBY PUMPDRAIN—DRAIN

76. LANDING GEAR, DOORS, STRUT, TIRES, BRAKES—CHECKED

77. CHOCK—REMOVE

78. FIRE EXTINGUISHER PRESSURE (IF INSTALLED)—CHECK

79. EXTERNAL POWER DOOR—SECURE

68. BATTERY AIR EXHAUST (NICKEL-CADMIUM)—CLEAR

See Foldout Page For Specific Locations




80. RAM SCOOP FUEL VENT AND HEATED FUEL VENT—CLEAR

81. GRAVITY LINE DRAIN—DRAIN

82. INVERTER COOLING LOUVERS—CLEAR

83. WING LEADING EDGE TANK SUMP—DRAIN

84. ICE LIGHT—CHECK

85. OUTBOARD DEICE BOOTS—CHECKED

86. TIEDOWN—REMOVE

87. FLUSH OUTBOARD DRAIN—DRAIN

88. MAIN FUEL TANK CAP—SECURE

89. NAVIGATION, RECOGNITION, STROBE LIGHT—CHECKED

90. STALL STRIP—CHECK

91. AILERON, FLAPS—CHECKED

92. BRAKES—CHECK

93. BRAKE DEICE (IF INSTALLED)—CHECK

94. OIL BREATHER VENT—CLEAR

95. BENDABLE TAB (RIGHT AILERON ONLY)

See Foldout Page For Specific Locations




96. LOWER ANTENNAS—CHECKED 100. CABIN AIR EXHAUST—CLEAR

97. VENTRAL FIN DRAIN HOLES—CLEAR98. TIEDOWN—REMOVE

99. LOWER AFT CABIN ACCESS DOOR—SECURE

101. OXYGEN SERVICE ACCESS DOOR—SECURE102. RIGHT STATIC PORTS—CLEAR

103. ELT AFT ARMING SWITCH (PRIOR TO BB-1510)—ARMED




105. VENTRAL FIN, STATIC WICK (1)—CHECKED

104. ACCESS PANEL—SECURE

106. RUDDER, STATIC WICKS (4)—CHECKED

108. ELEVATOR, ELEVATOR TAB, STATIC WICKS (3 EACHSIDE)—CHECKED

109. POSITION LIGHT—CHECK

110. TAIL FLOODLIGHTS (LEFT AND RIGHT IFINSTALLED)—CHECKED

107. RUDDER TAB—CHECK

111. HORIZONTAL STABILIZER, DEICE BOOTS (TAIL)—CHECKED




112. ACCESS PANEL—SECURE

113. RELIEF TUBE DRAIN—CLEAR

116. LEFT STATIC PORTS—CLEAR

114. AFT COMPARTMENT DRAIN TUBE—CLEAR115. OXYGEN OVERPRESSURE DRAIN TUBE—CLEAR




91959192939496999897100105

107

106

108 111

110104 103 101 102

747578817776838279808687

89

88 85 90 7184 72 73

ANN-1FOR TRAINING PURPOSES ONLY



ANNUNCIATOR PANELSThe Annunciator section presents a color representation of all theannunciator lights in the plane.




ANN-3FOR TRAINING PURPOSES ONLY



NOSE

L R

GNDCOMMPWR

CPS APR

ARM ACTV

GPSINTEG

GPS APR

ARM ACTV

GPS CRS

OBS LEG

GPWS

P/TEST

BELOWG/S

P/CANCEL

0I

MDH

ALTALERT

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

G/SCANCLD

GPWSINOP

GPWSFLAPOVRD


OKD

PARKING BRAKEOFF

COPILOTAIR

PULLON

INCR

ENVIRONMENTALCABIN

HIGH

AUTO

LO

DECR

INCR

CABIN TEMP

CABIN TEMP MODE

OFF

MANUALTEMP

VENTBLOWER

NO SMOKE& FSB

MAINHEAT

AUTO

MAINCOOL CABIN

AIR

PULLDECR

DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

PARKING BRAKE

ENGINE ANTI-ICE

AUTOFEATHER

ON

ON

GENRESET

OFF

ON

OFF

STARTER ONLY

ON

MAIN

OFF

OFF

AVIONICSMASTER PWR

ACTUATORSTANDBY

ICEEMERGENCY

EXTENSIONLEFT ENG

OFF

OFF

PROP GOVTEST

ARM

LEFT RIGHT

LEFT RIGHT

TEST

OFF

ARM

OFF

IGNITION ANDENGINE START

ENG AUTOIGNITION

OXYGENMASK

MICNORMAL

PILOTAIR

PULLON

COLLINS

CH SEL PWR

COLLINS

OFF

FSB

START/BRIGHT

DIM

OFF

FURNON

OFF

NO 2

INVERTERNO 1

OFF

BATT GEN 1 GEN 2

PROP SYN

OFF

ON

- +GYROSLAYING

SLEW MODE

COLLINS

DG

MASTER SWITCHGYRO

SUCTIONINCHES OF MERCURY

Beechcraft

OXYGENMASK

MICNORMAL

VOL VOL VOL

DIM

VOL VOL

AUTOCOMM

COMM

OFF

AUTOCOMM

OFFPILOT AUDIO OFF

NAV

AUDIOSPKR

SIDE-TONE

INTPHSENS

OFF

AUDIOSPKR

OFFNORM

HI

LO

DME

RANGE


OFF

HOTINTPH

MKR BCN DME1 12 2

1 2

2 2 ADF1 1COMM

COPILOT AUDIO OFF

NAV DME1 12 2 2 2 ADF1 1

BOTH

RANGE

VOICE BOTH

COMM 1

NAV 1

TRANSPONDER

COMM 2

NAV 2

MKR BCN1 & 2

COMM 1

COMM 2

CABIN COMM 1

COMM 2

CABIN

PUSHON/OFF

GNDCOMMPWR

ANNPUSH BRT3,0 00

ALTITUDE SET

ALTALERT

PUSH TO

CANCEL

COLLINS

SIDE-TONE

INTPHSENS

ENCDALTM

1

2

COM

OFFON

STO

TESTACT

SQOFF

Collins

MEM MEM

XFR

NAV

OFFON

STO

TESTACT

HLD

Collins

MEM MEM

XFR

ATC

STBYOFF

ONIDENT

TESTPRE

ALT

Collins

2

1

ADFANTOFF

ADFSTO

TESTACT

TONE

Collins

MEM MEM

XFR

NAV

OFFON

STO

TESTACT

HLD

Collins

MEM MEM

XFR

COM

OFFON

STO

TESTACT

SQOFF

Collins

MEM MEM

XFR


ACT

AUXDRIVE

EFIS

ALTALERT

CPS APR

ARM ACTV

GPSINTEG

GPS APR

ARM ACTV

GPS CRS

OBS LEG

PNEUMATICPRESSURE

0 20

10

Beechcraft

PSI

- +GYROSLAVING

SLEW MODE

COLLINS

L R

DC

DG

GPWS

P/TEST

BELOWG/S

P/CANCEL

ADF

ADF

NAV NAV

COLLINS

3033

3027

24

21 1815

129

6

EFIS

HORNON

TEST

OFF

SILENCE

AUX POWER

COLLINS

SEL PWR

FAST

ERECT

E

N

W

S

12

6

3

33

30

24

21

15

COLLINS

DIST ET00:00

CRSBRGFMS1 013

LOC1

10

20

10

DH 500

ALT

YD

HDG ALT

0

COLLINS

09

8 2

7

5

3

2 9 9 2

1 300200

BAROTEST

1 TCAS OFF

VERT SPEED

X1000 FPM

1 2

2

4

4

6

.5

.5

0

RA/B

262422

54

20

2018

FTLB X 1009

10121416

TORQUE 262422

54

20

2018

FTLB X 1009

10121416

TORQUE

START

912

8

76 5

4

2

ITT

PROP23

22

21

20

1918

17 16 1514

13

10

50

RPM X 100

PROP23

22

21

20

1918

17 16 1514

13

10

50

RPM X 100

TURBINE110

100

8040

20

0

60

%RPM

TURBINE110

100

8040

20

0

60

%RPM

FUEL FLOW

PPH X 100

65

4

3

2 1

0

FUEL FLOW

PPH X 100

65

4

3

2 1

0

ADF

ADF

NAV NAV

COLLINS

3033

3027

24

21 1815

129

6

ADF

ADF

NAV NAV

COLLINS

3033

3027

24

21 1815

129

6

GPWS

P/TEST

BELOWG/S

P/CANCEL

1 TCAS OFF

VERT SPEED

X1000 FPM

1 2

2

4

4

6

.5

.5

0

RA/B

E

N

W

S

12

6

3

33

30

24

21

15

DME2

9.3

GSPñ ñ ñ

CRS 359

VOR2

9.3 ICT

20

DH

TESTCollins

1510 5

RAD ALT

4

3

2

1

COLLINS

OFF

NAV

MAG1

VOR VRS 199FMS 1

134 00115

150 300

NRDR

PWR INT

NAV

RMT

PGE

EMG

RCL SKP CLR

CABIN AIR

50∞F0 100

80 FLIGHTHOURS 1/10

500

0

1500

2000

1000

SUPPLY PRESSURE

MADE IN USA

OXYGENSTALLWARN TEST

ELECHEAT

OFFOFF

OFF

AFTBLOWER

ON

INSTR & ENVIR OFF

ENVROFF

OPENLEFT RIGHT

BLEED AIR VALVES

COFFEE

OFF


LEXTR

LDET

R

0

10 20

30PROP AMPS

L IGNITION ON

L ENG ICE FAIL

L CHIP DETECT

L DC GEN

L BL AIR OFF

BRAKE DEICE ONLDG/TAXI LIGHT

ELEC TRIM OFF

BATTERY CHARGE

PROP SYNC ON

HYD FLUID LOW

FUEL CROSSFEED

PASS OXY ON

AIR COND N1 LOW

EXT PWR

DUCT OVERTEMP

RVS NOT READY

R BL AIR OFF

ELEC HEAT ON

R IGNITION ON

R ENG ANTI-ICE

L AUTO FEATHER R AUTO FEATHER

R ENG ICE FAIL

R CHIP DETECT

R DC GEN

L ENG ANTI-ICE

LANDINGGEAR

STALLWARN

BRAKEDEICE


NORMAL

SURFACEDEICE

SINGLE PITOT


OFF

OFF

LIGHTS

MANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

LEFT RIGHT

RELAY

OFF

OFF


HI

DEFROSTAIR

PULLON

2

FLAPS

DOWN

TAKEOFF

AND

APPROACH

UP20

60

80

4

6

2 41

1

0.5

.5


71

2

345

6

0PSI

CABINALT

00 FT

40 5

10

152025

30

35

0

PSI

3 64 5

L R

DC

31 002992

2

3

4

ALT

56

7

89 0 1

,

IN

PULL

. AIRBEECHCRAFT

10

10

20 20

10

10

0I

MDH

ALTALERT

x 00

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

4060

80

100

140120

180 160

220

200

240

260

300280

KNOTS

4060

80

100

140120

180 160

220200

240

260

300280

KNOTS

AUXON

AUX TEST

RIGHTLEFT

OFF

G/SCANCLD

GPWSINOP

GPWSFLAPOVRD

PRESS

TO TEST

L FUEL PRESS

INVERTER ALT WARNDOOR UNLOCKED

L BL AIR FAIL R BL AIR FAILA/P FAIL

L OIL PRESS

R FUEL PRESS

R OIL PRESSR GEN OVHTA/P TRIM FAIL

R ENG FIRE

L GEN OVHT

L ENG FIRE


OKD


OKD

PUSH

100

V

V V

V

ADF

DATA

CALCSTAT

SETUPOTHER

NAVFPL

MODETRIP

APTVORNDBINT

SUPL

NAVD/T

ACTVREFCTR

MSG ALT B CLR ENT

CRSR CRSR

BENDIX/KINGKLN 90B TSO

BR

GPSPUSH

ON

PULLSCAN

DIS 34.5NM

OBS IN - - -∞315∞130∞

ONANNUNRMI

OUT

ï ï ï ï ï ï ï ï ï ï

29

CRSR

BARO:APPROVE?

.92"

DATE/TIME23 FEB 88

21 :29 :38SDTALT 01300FT

ENRñLEG


START

912

8

76 5

4

2

ITT

C X 100

START

912

8

76 5

4

2

ITT

C X 100

MKR BCN

C PSIo

OIL140

100

100

150

200

60

200

0-20

50

C PSIo

OIL140

100

100

150

200

60

200

0-20

50

TCAS

A

OFF

35K15K

2 IDNM

VOL VOL

PRIOR TO BB-453BB-453 AND AFTER* OPTIONAL EQUIPMENT

WARNING ANNUNCIATOR PANEL—200 AIRCRAFT

WARNING ANNUNCIATOR PANEL—B200 AIRCRAFT(PRIOR TO BB-1444, EXCEPT BB-1439)

* OPTIONAL EQUIPMENT

WARNING ANNUNCIATOR PANEL—B200 AIRCRAFT(BB-1439, 1444 AND SUBSEQUENT)

*OPTIONAL EQUIPMENT

AUXDRIVE ALT

ALERT

Figure ANN-1. Annunciator Panels




ANN-4 FOR TRAINING PURPOSES ONLY



PARKING BRAKEOFF

COPILOTAIR

PULLON

INCR

ENVIRONMENTALCABIN

HIGH

AUTO

LO

DECR

INCR

CABIN TEMP

CABIN TEMP MODE

OFF

MANUALTEMP

VENTBLOWER

NO SMOKE& FSB

MAINHEAT

AUTO

MAINCOOL CABIN

AIR

PULLDECR

DOWNLOCK REL

BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

LIGHTS

LDG GEAR CONTROL

TEST

UP

NOSE

L R

DN

PARKING BRAKE

ENGINE ANTI-ICE

AUTOFEATHER

ON

ON

GENRESET

OFF

ON

OFF

STARTER ONLY

ON

MAIN

OFF

OFF

AVIONICSMASTER PWR

ACTUATORSTANDBY

ICEEMERGENCY

EXTENSIONLEFT ENG

OFF

OFF

PROP GOVTEST

ARM

LEFT RIGHT

LEFT RIGHT

TEST

OFF

ARM

OFF

IGNITION ANDENGINE START

ENG AUTOIGNITION

OXYGENMASK

MICNORMAL

PILOTAIR

PULLON

COLLINS

CH SEL PWR

COLLINS

OFF

FSB

START/BRIGHT

DIM

OFF

FURNON

OFF

NO 2

INVERTERNO 1

OFF

BATT GEN 1 GEN 2

PROP SYN

OFF

ON

- +GYROSLAYING

SLEW MODE

COLLINS

DG

MASTER SWITCHGYRO

SUCTIONINCHES OF MERCURY

Beechcraft

OXYGENMASK

MICNORMAL

VOL VOL VOL

DIM

VOL VOL

AUTOCOMM

COMM

OFF

AUTOCOMM

OFFPILOT AUDIO OFF

NAV

AUDIOSPKR

SIDE-TONE

INTPHSENS

OFF

AUDIOSPKR

OFFNORM

HI

LO

DME

RANGE


OFF

HOTINTPH

MKR BCN DME1 12 2

1 2

2 2 ADF1 1COMM

COPILOT AUDIO OFF

NAV DME1 12 2 2 2 ADF1 1

BOTH

RANGE

VOICE BOTH

COMM 1

NAV 1

TRANSPONDER

COMM 2

NAV 2

MKR BCN1 & 2

COMM 1

COMM 2

CABIN COMM 1

COMM 2

CABIN

PUSHON/OFF

GNDCOMMPWR

ANNPUSH BRT3,0 00

ALTITUDE SET

ALTALERT

PUSH TO

CANCEL

COLLINS

SIDE-TONE

INTPHSENS

ENCDALTM

1

2

COM

OFFON

STO

TESTACT

SQOFF

Collins

MEM MEM

XFR

NAV

OFFON

STO

TESTACT

HLD

Collins

MEM MEM

XFR

ATC

STBYOFF

ONIDENT

TESTPRE

ALT

Collins

2

1

ADFANTOFF

ADFSTO

TESTACT

TONE

Collins

MEM MEM

XFR

NAV

OFFON

STO

TESTACT

HLD

Collins

MEM MEM

XFR

COM

OFFON

STO

TESTACT

SQOFF

Collins

MEM MEM

XFR


ACT

AUXDRIVE

EFIS

ALTALERT

CPS APR

ARM ACTV

GPSINTEG

GPS APR

ARM ACTV

GPS CRS

OBS LEG

PNEUMATICPRESSURE

0 20

10

Beechcraft

PSI

- +GYROSLAVING

SLEW MODE

COLLINS

L R

DC

DG

GPWS

P/TEST

BELOWG/S

P/CANCEL

ADF

ADF

NAV NAV

COLLINS

3033

3027

24

21 1815

129

6

EFIS

HORNON

TEST

OFF

SILENCE

AUX POWER

COLLINS

SEL PWR

FAST

ERECT

E

N

W

S

12

6

3

33

30

24

21

15

COLLINS

DIST ET00:00

CRSBRGFMS1 013

LOC1

10

20

10

DH 500

ALT

YD

HDG ALT

0

COLLINS

09

8 2

7

5

3

2 9 9 2

1 300200

BAROTEST

1 TCAS OFF

VERT SPEED

X1000 FPM

1 2

2

4

4

6

.5

.5

0

RA/B

262422

54

20

2018

FTLB X 1009

10121416

TORQUE 262422

54

20

2018

FTLB X 1009

10121416

TORQUE

START

912

8

76 5

4

2

ITT

PROP23

22

21

20

1918

17 16 1514

13

10

50

RPM X 100

PROP23

22

21

20

1918

17 16 1514

13

10

50

RPM X 100

TURBINE110

100

8040

20

0

60

%RPM

TURBINE110

100

8040

20

0

60

%RPM

FUEL FLOW

PPH X 100

65

4

3

2 1

0

FUEL FLOW

PPH X 100

65

4

3

2 1

0

ADF

ADF

NAV NAV

COLLINS

3033

3027

24

21 1815

129

6

ADF

ADF

NAV NAV

COLLINS

3033

3027

24

21 1815

129

6

GPWS

P/TEST

BELOWG/S

P/CANCEL

1 TCAS OFF

VERT SPEED

X1000 FPM

1 2

2

4

4

6

.5

.5

0

RA/B

E

N

W

S

12

6

3

33

30

24

21

15

DME2

9.3

GSPñ ñ ñ

CRS 359

VOR2

9.3 ICT

20

DH

TESTCollins

1510 5

RAD ALT

4

3

2

1

COLLINS

OFF

NAV

MAG1

VOR VRS 199FMS 1

134 00115

150 300

NRDR

PWR INT

NAV

RMT

PGE

EMG

RCL SKP CLR

CABIN AIR

50∞F0 100

80 FLIGHTHOURS 1/10

500

0

1500

2000

1000

SUPPLY PRESSURE

MADE IN USA

OXYGENSTALLWARN TEST

ELECHEAT

OFFOFF

OFF

AFTBLOWER

ON

INSTR & ENVIR OFF

ENVROFF

OPENLEFT RIGHT

BLEED AIR VALVES

COFFEE

OFF


LEXTR

LDET

R

0

10 20

30PROP AMPS

L IGNITION ON

L ENG ICE FAIL

L CHIP DETECT

L DC GEN

L BL AIR OFF

BRAKE DEICE ONLDG/TAXI LIGHT

ELEC TRIM OFF

BATTERY CHARGE

PROP SYNC ON

HYD FLUID LOW

FUEL CROSSFEED

PASS OXY ON

AIR COND N1 LOW

EXT PWR

DUCT OVERTEMP

RVS NOT READY

R BL AIR OFF

ELEC HEAT ON

R IGNITION ON

R ENG ANTI-ICE

L AUTO FEATHER R AUTO FEATHER

R ENG ICE FAIL

R CHIP DETECT

R DC GEN

L ENG ANTI-ICE

LANDINGGEAR

STALLWARN

BRAKEDEICE


NORMAL

SURFACEDEICE

SINGLE PITOT


OFF

OFF

LIGHTS

MANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

LEFT RIGHT

RELAY

OFF

OFF


HI

DEFROSTAIR

PULLON

2

FLAPS

DOWN

TAKEOFF

AND

APPROACH

UP20

60

80

4

6

2 41

1

0.5

.5


71

2

345

6

0PSI

CABINALT

00 FT

40 5

10

152025

30

35

0

PSI

3 64 5

L R

DC

31 002992

2

3

4

ALT

56

7

89 0 1

,

IN

PULL

. AIRBEECHCRAFT

10

10

20 20

10

10

0I

MDH

ALTALERT

x 00

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

PRESS TO RESET

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

4060

80

100

140120

180 160

220

200

240

260

300280

KNOTS

4060

80

100

140120

180 160

220200

240

260

300280

KNOTS

AUXON

AUX TEST

RIGHTLEFT

OFF

G/SCANCLD

GPWSINOP

GPWSFLAPOVRD

PRESS

TO TEST

L FUEL PRESS

INVERTER ALT WARNDOOR UNLOCKED

L BL AIR FAIL R BL AIR FAILA/P FAIL

L OIL PRESS

R FUEL PRESS

R OIL PRESSR GEN OVHTA/P TRIM FAIL

R ENG FIRE

L GEN OVHT

L ENG FIRE


OKD


OKD

PUSH

100

V

V V

V

ADF

DATA

CALCSTAT

SETUPOTHER

NAVFPL

MODETRIP

APTVORNDBINT

SUPL

NAVD/T

ACTVREFCTR

MSG ALT B CLR ENT

CRSR CRSR

BENDIX/KINGKLN 90B TSO

BR

GPSPUSH

ON

PULLSCAN

DIS 34.5NM

OBS IN - - -∞315∞130∞

ONANNUNRMI

OUT

ï ï ï ï ï ï ï ï ï ï

29

CRSR

BARO:APPROVE?

.92"

DATE/TIME23 FEB 88

21 :29 :38SDTALT 01300FT

ENRñLEG


START

912

8

76 5

4

2

ITT

C X 100

START

912

8

76 5

4

2

ITT

C X 100

MKR BCN

C PSIo

OIL140

100

100

150

200

60

200

0-20

50

C PSIo

OIL140

100

100

150

200

60

200

0-20

50

TCAS

A

OFF

35K15K

2 IDNM

VO L VOL

*Optional Equipment

(PRIOR TO BB-453) (BB-453 AND AFTER)

(BB-1439, 1444 AND SUBSEQUENT) (PRIOR TO BB-144, EXCEPT 1439)

Figure ANN-1. Caution-Advisory Annunciator Panels

Documents

Super King Air 200/B200 Pilot Training Manualaviaco-va.es/WP/BE20_Technical_Manual.pdf · INTRODUCTION This pilot training manual covers all systems on the Super King Air 200 and