King Air C90 A-B

KING AIR C90A/B PILOT TRAINING MANUAL VOLUME 2

Record of Revision No. .02

This is a revision of the King Air C90A/B Pilot Training Manual.

A solid vertical line in the margin indicates the content of the adjacent text or figurehas been changed. A vertical line adjacent to a blank space indicates material hasbeen deleted.

Any page affected by the revision is marked “Revision .02” in the lower left or rightcorner. If a page has “Revision .02” in the lower left or right corner and no verticalline in the margin, it is a page in which format only has been changed.

The changes made in this revision will be further explained at the appropriate timein the training course.

FlightSafetyinternational

COURSEWARE SUPPORT—HURST 8900 Trinity Blvd. Hurst, Texas 76053 (817) 276-7500 FAX (817) 276-7501

the best safety device in any aircraft is a well-trained crew . . .

KING AIR C90A/BPILOT TRAINING MANUAL

VOLUME 2AIRCRAFT SYSTEMS

FlightSafety

international

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Flushing, New York 11371(718) 565-4100

www.flightsafety.com

Courses for the King Air C90A/B and other Beech aircraft are taught at thefollowing FlightSafety learning centers:

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iii

NOTICE

The material contained in this training manual is based on information obtainedfrom the aircraft manufacturer’s Pilot Manuals and Maintenance Manuals. It is tobe used for familiarization 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 manufactureror the FAA, 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 ourtraining program.

FOR TRAINING PURPOSES ONLYFOR TRAINING PURPOSES ONLY

FOR TRAINING PURPOSES ONLY

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

APPENDIX

ANNUNCIATOR PANEL

INSTRUMENT PANEL POSTER


1-i

CHAPTER 1

CONTENTS

Page

AIRCRAFT GENERAL

INTRODUCTION ..................................................................................................................

1-1

GENERAL ...............................................................................................................................

1-1

AIRPLANE SYSTEMS...........................................................................................................

1-2

General .............................................................................................................................

1-2

Chapters............................................................................................................................

1-2

CHANGES DISTINGUISHING MODEL C90B FROM MODEL C90A..............................

1-4

BEECHCRAFT KING AIR C90A AND C90B DESCRIPTION ...........................................

1-7

King Air C90A and C90B Configuration.......................................................................

1-12

Cabin Entry and Exits.....................................................................................................

1-17

Emergency Exit ..............................................................................................................

1-19

Nose Baggage Door (Optional) ......................................................................................

1-19

Cabin Compartments ......................................................................................................

1-20

Flight Deck .....................................................................................................................

1-21

C90A Instrument Panel/Avionics...................................................................................

1-30

Control Surfaces .............................................................................................................

1-30

Tiedown and Securing....................................................................................................

1-33

Taxiing............................................................................................................................

1-34

Servicing Data ................................................................................................................

1-34

Product Support ..............................................................................................................

1-36

Preflight Inspection ........................................................................................................

1-36


1-iii

Figure Title Page

ILLUSTRATIONS

1-1

Beechcraft King Air C90A ......................................................................................

1-7

1-2

General Arrangement—C90A ..................................................................................

1-8

1-3

Three-View Diagram—C90A ..................................................................................

1-9

1-4

General Arrangement—C90B ................................................................................

1-10

1-5

Three-View Diagram—C90B.................................................................................

1-11

1-6

Engine Air Inlet ......................................................................................................

1-12

1-7

Optional Cabin Seating...........................................................................................

1-12

1-8

King Air C90A Front Three-Quarter View (Engines Primary)..............................

1-15

1-9

King Air C90B in Flight .........................................................................................

1-16

1-10

Entrance and Exit Provisions..................................................................................

1-17

1-11

Dual Door Cables ...................................................................................................

1-18

1-12

Nose Compartment Door ........................................................................................

1-19

1-13

Cabin Areas ............................................................................................................

1-20

1-14

Cabin Seating Layout (Typical)..............................................................................

1-20

1-15

Flight Deck Layout .................................................................................................

1-21

1-16

Control Wheels and Fuel Control Panel .................................................................

1-22

1-17

Instrument Panels....................................................................................................

1-23

1-18

Instrumentation .......................................................................................................

1-24

1-19

Engine Instruments—Prior to LJ-1361...................................................................

1-25

1-20

Engine Instruments—LJ-1361, LJ-1363, and After ...............................................

1-26

1-21

Pedestal and Right Side Panel ................................................................................

1-27

1-22

Pilot’s and Copilot’s Subpanels..............................................................................

1-28

1-23

Annunciators...........................................................................................................

1-29

1-iv


1-24

Overhead Light Control Panel................................................................................

1-30

1-25

Avionics Installation (C90A) .................................................................................

1-31

1-26

Flight Control Surfaces ..........................................................................................

1-31

1-27

Flight Control Locks ..............................................................................................

1-32

1-28

Tiedowns ................................................................................................................

1-32

1-29

Propeller Boots.......................................................................................................

1-33

1-30

Turning Radius.......................................................................................................

1-34

1-31

Danger Areas..........................................................................................................

1-34

1-32

Servicing Data ........................................................................................................

1-35

1-33

Exterior Inspection .................................................................................................

1-36


1-v

Table Title Page

TABLES

1-1

Specifications—C90A ............................................................................................

1-13

1-2

Specifications—C90B.............................................................................................

1-14


1-1

CHAPTER 1AIRCRAFT GENERAL

INTRODUCTION

A good basic understanding of the airplane will help in studying the individual systems andtheir operation. This chapter provides basic and background information needed to learn thedetails of airplane operation and performance to be studied in other chapters.

GENERAL

This chapter of the training manual presents anoverall view of the airplane. This includes exter-nal familiarization, cabin arrangements, andcockpit layout.

In this chapter of the training manual you willfind diagrams and data describing the airplane ingeneral and its systems that are not included inthe

Pilot’s Operating Handbook (POH)

.

Reference material in this training manual isorganized into 15 chapters (with two unusedtabs) covering all airplane systems. Each chapteris complete and independent and can be referredto in any sequence.

Following are brief descriptions of the subjectmatter in each chapter. All material is discrete tothe Beechcraft King Air C90A and C90B.

1-2


AIRPLANE SYSTEMS

GENERAL

The “Systems Description” section of the

POH

gives a brief description of all the systems incor-porated in the King Air C90A and C90B.Additional description and details of these sys-tems are included in separate chapters of thistraining manual. The

POH

information isupdated as required and always supersedes anyinformation in this training manual.

CHAPTERS

Aircraft General

Chapter 1, “Aircraft General,” presents an overallview of the airplane. This includes externalfamiliarization, cabin arrangement, and cockpitlayout. In this chapter you will find diagrams anddata describing the airplane in general that arenot included in the

Pilot’s Operating Handbook

.

Electrical Power Systems

Chapter 2, “Electrical Power Systems,” describesthe airplane electrical system and its compo-nents. The electrical system is discussed to theextent necessary for pilot management of all nor-mal and emergency operations. The location andpurpose of switches, indicators, lights, and cir-cuit breakers are noted. DC and AC generationand distribution are described. This chapter alsoincludes electrical system limitations and a dis-cussion of potential electrical system faults.

Lighting

Chapter 3, “Lighting,” discusses cockpit lighting,cabin lighting, and exterior lighting. All lights areidentified and located. The location and use ofcontrols for the lighting system are also included.

Master Warning System

Chapter 4, “Master Warning System,” presents adescription and discussion of the warning, cau-tion, and advisory annunciator panels. Eachannunciator is described in detail, including itspurpose and associated cause for illumination.Emphasis is on corrective action required by thepilot if an annunciator is illuminated.

Fuel System

Chapter 5, “Fuel System,” presents a descriptionand discussion of the fuel system. The physicallayout of fuel cells are described. Correct use ofthe boost pumps, transfer pumps, crossfeed, andfirewall shutoff valves are discussed. Locationsand types of fuel drains and correct proceduresfor taking and inspecting fuel samples aredetailed. This chapter includes a list of approvedfuels and procedures for fuel servicing.

Powerplant

Chapter 7, “Powerplant,” presents a discussion ofthe Pratt and Whitney PT6A turboprop engines.Engine theory and operating limitations aredescribed, and normal pilot procedures aredetailed. Crewmembers must have sufficientknowledge of the PT6A series engines to under-stand all normal and emergency procedures.

This chapter also describes the propeller system.Location and use of propeller controls, principleof operation, reversing, and feathering arediscussed.

Fire Protection

Chapter 8, “Fire Protection,” describes the firewarning and protection systems. Operation andtesting information for the fire detection and fire-extinguishing systems is included.


1-3

Pneumatics

Chapter 9, “Pneumatics,” presents a discussion ofpneumatic and vacuum systems. Sources andoperation of pneumatic and vacuum air aredescribed. Acceptable gage readings and normaland abnormal system indications are outlined.

Ice and Rain Protection

Chapter 10, “Ice and Rain Protection,” presents adescription and discussion of the anti-ice anddeice systems. All of the anti-ice, deice, and rainprotection systems in this airplane are described,showing location, controls, and how they areused. The purpose of this chapter is to acquaintthe pilot with all the systems available for flightin icing or heavy rain conditions and their con-trols. Procedures in case of malfunction in anysystem are included. This also includes informa-tion concerning preflight deicing and defrosting.

Air Conditioning

Chapter 11, “Air Conditioning,” presents adescription of the air-conditioning, heating, andfresh air systems. Each subsystem discussionincludes general description, principle of opera-tion, controls, and emergency procedures.

Pressurization System

Chapter 12, “Pressurization,” presents a descrip-tion of the pressurization system. The function ofvarious major components, their physical loca-tion, and operation of the pressurization systemcontrols are discussed. Where necessary, refer-ences are made to the environmental system as itaffects pressurization.

Landing Gear and Brakes

Chapter 14, “Landing Gear and Brakes,” presentsa description and discussion of the landing gearsystem, landing gear controls, and operating lim-itations. The indicator system and emergencylanding gear extension are also described.

This chapter also discusses the wheel brakesystem. Correct use of the brakes and parkingbrakes, along with brake system description,and what to look for when inspecting brakesare detailed.

Flight Controls

Chapter 15, “Flight Controls,” describes the four-segment Fowler-type flap system. System con-trols and limitations are considered, withreference to operation as outlined in the

Pilot’sOperating Handbook

.

This chapter also describes the rudder boost sys-tem. This system is designed to reduce piloteffort if single-engine flight is encountered.

Avionics

Chapter 16, “Avionics,” describes the standardavionics installation for the King Air C90A andC90B. The avionics controls, along with theweather radar, are mounted on an isolation panelin the center of the instrument panel so that it iseasily available to the pilot or copilot. Individualaudio switches, across the top of the panel, con-trol audio to the speakers or headphones for thepilot and copilot. There are separate sets of con-trols for pilot and copilot so that each can selectaudio from any nav or comm receiver.

A glossary of avionics terminology is included inan Appendix at the back of this training manual.

This chapter also presents a discussion of thedual pitot-static system, which is vital to airspeedindications in the airplane. The principle of oper-ation, sources of static and pitot pressure,instruments that depend on the system, and thepilot’s alternate static air source are covered.

1-4


Miscellaneous Systems

Chapter 17, “Miscellaneous Systems,” presents asummary of the oxygen system and its compo-nents . General descr ipt ion, pr inciple ofoperation, system controls, and emergency pro-cedures are included. Use of the oxygen durationchart involves working simulated problems undervarious flight conditions. FAR requirements forcrew and passenger oxygen needs are part of thediscussion, as well as the types and availability ofoxygen masks. Local servicing procedures refer-enced in the


are alsoincluded.

CHANGES DISTINGUISHING MODEL C90B FROM MODEL C90A

The following are significant changes that differ-entiate the Model C90B from the Model C90A.The C90B serial numbers are LJ-1288, LJ-1295,LJ-1302, LJ-1303, LJ-1305 through LJ-1308, LJ-1311, LJ-1312, LJ-1314 through LJ-1316, LJ-1318, LJ-1320 and subsequent.

●

Significantly reduced cabin sound andvibration levels.

●

Four-blade dynamically balanced 90-inch-diameter McCauley or Hartzell pro-pellers. Includes:

• Improved low-friction hub toimprove propeller synchrophasing.

• New streamlined, more aerodynamicpropeller spinners.

• A gated ground fine power lever posi-tion which provides improved groundhandling, as well as reduced acceler-ate-stop and landing distances. Theground fine position allows a flatterpropeller blade angle to be used fortaxi and for deceleration duringaccelerate-stop and landing.

• Dynamic propeller balancing toreduce propeller vibration and associ-ated airframe vibration to improvecomfort and reduce fatigue.

●

Hardwire installation on airplane to sim-p l i f y i n -fie ld dynamic p rope l l e rbalancing. Canon plug for Chadwick-Helmuth dynamic propeller balancinganalyzer is located just aft of copilot’sseat on cockpit sidewall.

●

Super King Air 350 follow-up type flapselector switch. Flaps follow position offlap selector with three positions: up,approach, and down.

●

Super King Air 300/350 type approachchart holder on pilot’s and copilot’s con-trol wheels (optional).

●

Avionics compartment moisture barrierto prevent infiltration of water into theavionics bay through the avionics bayaccess door. Improves avionics reliabilityand life. This same barrier is used on theB200 and 350.

●

Digital outside air temperature systemsimilar to Super King Air 300/350.

●

Cockpit flashlight and flashlight holdermounted on control pedestal similar toSuper King Air Model 350.

●

Changed chip detect warning annuncia-tion to caution annunciation—alsochanged chip detect procedures in

PilotOperating Handbook

and checklist. Newannunciator and checklist eliminate needfor emergency shutdown of affectedengine. Procedure now simply requiresthe pilot to monitor engine performance.


1-5

●

COMPLETELY REDESIGNED INTERIOR

• Incorporates 26 electronically tuneddynamic vibration absorbers mountedin strategic locations on specific fuse-lage frames. Provides 12 to 15 dBreduction in specific frequency rangesto significantly reduce propeller-induced sound and vibration.

• New lightweight composite soundabsorbing headliner, sidewall panels,and floorboards.

• New thermal and acoustic insulat-ing materials in cockpit and cabinsidewalls (bagged insulation andScotch damp).

• Window reveals covered withstretched wool headliner material.

• Indirect fluorescent lighting behindwindow reveals now has bright anddim lighting intensities.

• Sidewall panels incorporate integralsidewall armrest and Super King Air350 style tables. Integral sidewallarmrests allow for greater comfort byoffering greater seat width betweenarmrests when seat is lateral lytracked. New tables offer greaterstrength and improved operation.Adjustable set screws are provided toallow easy adjustment to ensure thetable is level in the extended position.

• Redesigned side facing seat with par-tition. Redesign of the side facingseat allowed the vertical partition tobe moved three inches further aft toprovide an additional three inches oflegroom for right side of the club.

• Redesigned and restyled cabin chairs.New chairs feature Super King Air350 styling. The cushion on the seatbacks were retailored and reshaped.The new shape allows the occupant tosit 1 1/2 inches further back in theseat. This offers a total of threeinches of additional legroom betweenthe seats in the club for greater pas-senger comfort.

• Combined seat redesign and reloca-tion of aft partition increases legroomin club by a total of six inches. Stan-dard C90B seat pitch is 59 inches.B200 seat pitch in club is 57 andC90A was 53 inches.

• New carpet installation covers all butseven inches of the seat tracks toallow for seat travel. Covered seatt racks provide greater comfortthrough reduced noise and cold thatradiates from the exposed track.Cabin image is also improved.

• Seat belt chime and no smokings ign re loca ted and upda ted toSuper King Air 350 type. Nowlocated on forward and aft parti-tions, for better visibility.

• Standard no smoking configurationremoves cigarette lighters and ash-trays. No smoking sign will remainilluminated during all operations.Ashtrays and cigarette lighters arepackaged in a smoking option.

• Forward right side cabinet nowincludes, in addition to the hot bever-age dispenser, a cold beveragedispenser. This unit has a servinglight similar to the Super King Airrefreshment centers.

1-6


• A new cabinet (optional) has beendesigned for the aft right side of thecabin. This cabinet features the fourliquor decanters and an insulatedice drawer.

• Forward sliding door partition is nowstandard.

• All cabinet drawer slides featureroller-bearing-type guides to improveoperation of drawers. C90A slideswere friction-type guides.

●

Optional quick-disconnect second doorcable allows for greater ease in loadingbaggage/cargo but retains second cablefor strength.

●

New stylized C90B logo under cockpitD-windows provides common image andstyling with Super King Air 350.

●

New updated and distinctive paintscheme to establish C90B as new and dif-ferent from all previous King Airs.

●

Significant performance improvementssubstantially improve safety.

• Reduced V

MCA

now 80 KIAS was9 0 K I A S . V

M C A

wa s r e d u c e dbecause the new four-blade propel-lers with reduced diameter producelower thrust at high power settingsand low speeds.

• Accelerate-stop distance reduced9.7%. Now 3,650 feet, was 4,042feet. (Sea level, standard day, nowind at maximum weight.)

• Accelerate-go distance reduced18.8%. Now 3,650 feet over 35-footobstacle, was 4,500 feet over 50-footobstacle (sea level, standard day, nowind at maximum weight).

• Landing distance reduced 6.3%.Now 2,290 feet, was 2,443 feet(sea level, standard day, no wind atmaximum weight).

●

New

POH

features Abnormal and Emer-gency Sections with bold-face actionitems. The

POH

has been revised toreflect all the changes to the airplane, aswell as to make the

POH

more consistentwith the other King Air models.


1-7

BEECHCRAFT KING AIR C90A AND C90B DESCRIPTION

The Beechcraft King Air C90A and C90B arehigh-performance, conventional tail, pressur-ized, twin-engine turboprop airplanes (Figures1-1 through 1-5). They are designed andequipped for flight in IFR conditions, day ornight, into high-density air traffic zones, and intoknown or forecast icing conditions. They are alsocapable of operating in and out of small unim-proved airports within the

POH

operating limits.

The King Air design is a blend of a highly effi-cient airframe with proven current technologycomponents, providing a reliable, economical,versatile, and cost-productive airplane.

The structure is all-metal, low-wing mono-plane. It has fully cantilevered wings and aconventional-tail empennage. The wings are anefficient, high-aspect ratio design. The airfoilsection provides an excellent combination oflow drag for cruise conditions, and easy han-dling for the low-speed terminal conditions orsmall airport operations.

Figure 1-1 Beechcraft King Air C90A

1-8


Figure 1-2 General Arrangement—C90A


1-9

Figure 1-3 Three-View Diagram—C90A

1-10


Figure 1-4 General Arrangement—C90B


1-11

Figure 1-5 Three-View Diagram—C90B

1-12


A faired, oval, minimum frontal area nacelle isinstalled on each side of the wing center sectionto house both the engine and landing gear. The“pitot” type intakes (Figure 1-6) boost perfor-mance by reducing drag, and the exhaust stacksare shaped for smaller frontal area to reducedrag. The nacelles are designed and located tomaximize propeller/ground clearance, minimizechain noise, and provide a low-drag installationof the powerplants on the wing.

The fuselage is conventional monocoque struc-ture using high-strength aluminum alloys. Thebasic cross-sectional shape of the cabin is afavorable compromise between passengercomfort and efficient cruise performance. Thecabin profile is squared-oval, not round. Pas-sengers can sit comfortably without leaningtheir heads to accommodate sloping walls. Thefloors are flat from side to side for passengerease in entering and leaving the cabin. TheBeechcraft King Air C90A and C90B are cer-tificated for up to 10 people (Figure 1-7). Themost popular configuration provides comfort-able seating for six passengers and a crew oftwo. Almost any arrangement is possible.

KING AIR C90A AND C90B CONFIGURATION

The King Air C90A and C90B are powered byPratt & Whitney, 550 shp (flat-rated) PT6A-21turboprop engines. In addition to the standardairplane configurations, Beechcraft offersmany optional items which are available atadditional cost and weight. The basic configu-r a t i o n s , d i m e n s i o n s , w e i g h t s , a n dspecifications are summarized in Tables 1-1and 1-2. Refer to the respective airplane

POH

for detailed, up-to-date information.

Figure 1-6 Engine Air Inlet

Figure 1-7 Optional Cabin Seating

Arrangement


1-13

Table 1-1 SPECIFICATIONS—C90A

Model Designation - Passenger........................................................... C90AMinimum Crew........................................................................................... 1Occupants - Max. FAA Cert. (incl. crew)................................................. 13Passengers - Normal Corp. Config. ............................................................ 6Engines - P&W Turboprop ......................................................... 2 PT6A-21Propellers - 3-Blade, Constant-speed,

Full-feathering, Counter-weighted,Hydraulically-actuated.......................................................... 2 McCauley

Landing Gear - Retractable, Tricycle............................................ HydraulicWing Area ................................................................................ 293.94 sq. ft.

Maximum Certificated Weights

LJ1138 andLJ1063-1137 Subsequent

Certificated Weights and 1146 Except 1146Maximum Ramp Weight ................ 9,710 pounds............... 10,160 poundsMaximum Take-off Weight ............ 9,650 pounds............... 10,100 poundsMaximum Landing Weight ..............9,168 pounds................. 9,600 poundsMaximum Zero Fuel Weight................................. No Structural LimitationMaximum Weight in Baggage Compartment:

Rear Baggage Compartment .................................................. 350 poundsNose Baggage Compartment

(Baggage and Avionics)..................................................... 350 pounds

Cabin and Entry Dimensions

Cabin Width (Maximum)............................................................... 54 inchesCabin Length (Partition to Partition)........................................... 155 inchesCabin Length (Maximum between pressure bulkheads)............. 214 inchesCabin Height (Maximum) ............................................................. 57 inchesAirstair Entrance Door Width (Minimum).................................... 27 inchesAirstair Entrance Door Height (Minimum)................................ 51.6 inchesSill Height (Maximum) ................................................................. 48 inchesPressurized Compartment Volume...................................... 313.6 cubic feet

Rear Baggage Compartment Volume................................ 53.5 cubic feet

Nose Avionics/Baggage Compartment Volume.................... 16 cubic feet

Specific LoadingsLJ1138 and

LJ1063-1137 Subsequentand 1146 Except 1146

Wing Loading............................. 32.8 pounds/Ft2........... 34.4 pounds/Ft2

Power Loading ........................... 8.8 pounds/H.P............. 9.2 pounds/H.P.

1-14


Table 1-2 SPECIFICATIONS—C90B

Model Designation - Passenger ........................................................... C90BCrew - FAA Certified.................................................................................. 1Occupants - Max. FAA Cert. (incl. crew)................................................. 13Passengers - Normal Corp. Config. ............................................................ 6Engines - P&W Turboprop ......................................................... 2 PT6A-21Propellers - 4-Blade, Constant-speed,

Full-feathering, Counter-weighted,Hydraulically-actuated................................. 2 McCauley (full reversing)

Hartzell after LJ 1542Landing Gear - Retractable, Tricycle............................................ HydraulicWing Area ................................................................................ 293.94 sq. ft.

Maximum Certificated Weights

\Maximum Ramp Weight ..................................................... 10,160 poundsMaximum Take-off Weight ................................................... 10,100 poundsMaximum Landing Weight....................................................... 9,600 poundsMaximum Zero Fuel Weight................................. No Structural LimitationMaximum Weight in Baggage Compartment:

Rear Baggage Compartment .................................................. 350 poundsNose Baggage Compartment

(Baggage and Avionics)..................................................... 350 pounds

Cabin and Entry Dimensions

Cabin Width (Maximum)............................................................... 54 inchesCabin Length (Partition to Partition)........................................... 155 inchesCabin Length (Maximum between pressure bulkheads)............. 214 inchesCabin Height (Maximum) ............................................................. 57 inchesAirstair Entrance Door Width (Minimum).................................... 27 inchesAirstair Entrance Door Height (Minimum)................................ 51.6 inchesSill Height (Maximum) ................................................................. 48 inchesPressurized Compartment Volume...................................... 313.6 cubic feet

Rear Baggage Compartment Volume................................ 53.5 cubic feet

Nose Avionics/Baggage Compartment Volume.................... 16 cubic feet

Specific Loadings

Wing Loading................................................... 32.8 pounds per square footPower Loading.......................................... 8.8 pounds per shaft horsepower


1-15

C90A Operating Speeds

The Beechcraft King Air C90A (Figure 1-8)qualifies as one of the most maneuverable corpo-rate airplanes in the world. Insistence onhandling ease in all flight regimes and tough con-struction techniques contribute to the followingKIAS data (calculated at maximum takeoffweight of 9,650 pounds):

Maximum operating speed (V

MO

) ....... 226 knots

Maneuvering speed (V

A

) .............. 153/169 knots

Maximum landing gear operating speed (V

LO):

Extensions/extended ...................... 182 knots

Retraction....................................... 163 knots

Maximum flap extension/extended (VFE):

Approach ....................................... 184 knots

Down.............................................. 148 knots

Stall (100% flaps, power off) ................. 76 knots

Air minimum control (VMCA) ............... 90 knots

Figure 1-8 King Air C90A Front Three-Quarter View (Engines Primary)

1-16 FOR TRAINING PURPOSES ONLY

C90B Operating SpeedsThe Beechcraft King Air C90B (Figure 1-9)qualifies as one of the most maneuverable corpo-rate airplanes in the world. Insistence onhandling ease in all flight regimes and tough con-struction techniques contribute to the followingKIAS data (calculated at maximum takeoffweight of 10,100 pounds):

Maximum operating speed (VMO) ...... 226 KIAS

Maneuvering speed (VA) .................... 169 KIAS

Maximum landing gear operating speed (VLO):

Extension/extended ....................... 182 KIAS

Retraction ...................................... 163 KIAS

Maximum flap extension/extended (VFE):

Approach....................................... 184 KIAS

Down............................................. 148 KIAS

Stall (100% flaps, power off)................. 78 KIAS(with four-blade propeller installed)

Air minimum control (VMCA)............... 80 KIAS(with four-blade propeller installed)

Figure 1-9 King Air C90B in Flight

FOR TRAINING PURPOSES ONLY 1-17

CABIN ENTRY AND EXITSThe cabin entry airstair door is on the left side ofthe fuselage, just aft of the wing (Figure 1-10). Aswing-down door, hinged at the bottom, providesa convenient stairway for entry and exit.

Two of the four steps are movable and automati-cally fold flat against the door in the closedposition. A self-storing platform automaticallyfolds down over the door sill when the dooropens to provide a stepping platform for doorseal protection.

Figure 1-10 Entrance and Exit Provisions


A plastic-encased cable provides support for thedoor in the open position, a handhold for passen-gers, and a means of closing the door from insidethe airplane. A hydraulic dampener permits thedoor to lower gradually during opening. It isimportant that not more than one person be onthe airstair door at a time as excessive weightscould cause structural damage to the door.

Dual Door Cables with One Detachable (Optional)Dual stair assist cables are available as anoption on the C90B (Figure 1-11). Door assistcables provide passengers a way to stabilizethemselves when going up or down the stairs.The forward assist cable is easily detachable toprovide more room for loading large baggageor cargo into the airplane.

Airstair Locking MechanismThe door locking mechanism is operated byeither of the two vertically staggered handles,one inside and the other outside the door. Theinside and outside handles are mechanicallyinterconnected.

When either handle is rotated per placard instruc-tions, two latch bolts at each side of the door, andtwo latch hooks at the top of the door, lock intothe doorframe to secure the airstair door. A but-ton adjacent to the door handle must bedepressed before the handle can be rotated toopen the door. For security of the airplane on theground, the door can be locked with a key.

To secure the airstair door inside, rotate the han-dle clockwise as far as it will go. The releasebutton should pop out, and the handle should bepointing down. Check the security of the airstairdoor by attempting to rotate the handle counter-clockwise without depressing the release button;the handle should not move.

Next lift the folded stairstep that is just below thedoor handle. Ensure the safety lock is in positionaround the diaphragm shaft when the handle is inthe locked position.

To observe this area, depress a red switch nearthe window that illuminates a lamp inside thedoor. If the arm is properly positioned around theshaft, proceed to check the indication in each ofthe visual inspection ports located near each cor-ner of the door (see Figure 1-10). Ensure thegreen stripe on the latch bolt is aligned with theblack pointer in the visual inspection port.

Figure 1-11 Dual Door Cables


WARNING

Never attempt to unlock or check thesecurity of the door in flight. If theCABIN DOOR annunciator illumi-nates in flight, or if the pilot has anyreason to suspect that the door may notbe securely locked, the cabin pressureshould be reduced to zero differential,and all occupants instructed to remainseated with their seat belts fastened.After the airplane has made a full-stoplanding, only a crewmember shouldcheck the security of the airstair door.

EMERGENCY EXITThe emergency exit door is located at the thirdcabin window on the right side of the fuselage(see Figure 1-10). A placard at the window givesinstructions for access to the release mechanism.

The door is released from the inside with twohooks, a trigger button, and a latch-release pull-up handle. A placard on the emergency exit hatchrelease cover lists proper opening procedures.

A pressure lock prevents the door from beingopened when the cabin is pressurized. If pressur-ized, pulling the hooks overrides the pressurelock and al lows the t r igger but ton to bedepressed. This releases the latch-release handle.When the handle is pulled up and the securinglatches are released, a hinge at the bottom allowsthe hatch to swing outward and downward foremergency exit.

NOSE BAGGAGE DOOR (OPTIONAL)Prior to LJ-1531 the King Air C90A and C90Bhave an optional 16 cubic-foot nose baggagecompartment which is accessible through a doorlocated on the left side of the nose (Figure 1-12).This compartment is limited to 350 pounds,which includes the weight of the avionics equip-ment within the compartment.

The baggage door is hinged at the top to allowthe door to swing upward. A flush-mounted doorhandle with a push-to-release button activatesthree bayonet-type latching bolts that, whenengaged, will hold the door securely closed.

When not engaged, a switch at the forward latch-ing bolt will close, and the BAG DOOR OPENannunciator will illuminate. In addition, the dooris equipped with a secondary safety latch to holdthe door in a partially closed position in the eventthe primary latching bolt is not engaged.

The push-to-release button, adjacent to the doorhandle, will prevent the door inadvertentlyopening. For security of the unattended air-plane, the nose baggage compartment doorfeatures a key-lock latch.

Figure 1-12 Nose Compartment Door

(Baggage Compartment Not Shown)


CABIN COMPARTMENTSThe pressurized cabin interior consists of theflight deck, passenger seating area, and anaft baggage area (Figure 1-13). The flightdeck provides side-by-side seating for thepilot and copilot.

Typically for corporate use, the cabin is arrangedin a five-passenger club seating and aisle-facingcabinet seat layout (Figure 1-14).

A lavatory area is located in the aft compartment,with a padded seat which can be used as the sixthpassenger seat.

Aft of the cabin area is the baggage area. Thispressurized area is capable of holding 53.5cubic feet of luggage, cargo, or clothing (allaccessible in flight). The location of the bag-gage area next to the airstair door makesloading and unloading easy.

If an operation requires, some or all of theseats, wall partitions, and lavatory can bequickly removed to configure the airplane forcargo transport.

Figure 1-13 Cabin Areas

Figure 1-14 Cabin Seating Layout (Typical)


FLIGHT DECKThe flight deck layout is a time-proven designthat has optimized crew efficiency and comfort(Figure 1-15). The pilot and copilot sit side-by-side in individual chairs, separated by the controlpedestal. The seats are adjustable fore and aft aswell as vertically. Seat belts and inertia-typeshoulder harnesses are provided for each seat.

The general layout of the flight deck shows thelocation of the instruments and controls. Conven-tional dual controls are installed so that theairplane can be flown by either pilot (Figure1-16). The controls and instruments are arrangedfor convenient single-pilot operation or for a pilotand copilot crew.

Figure 1-15 Flight Deck Layout


Figure 1-16 Control Wheels and Fuel Control Panel


The fuel control panel (Figure 1-16) is located onthe left sidewall, next to the pilot. Fuel quantitygages and switches, firewall valve switches, andcircuit breakers are located on this panel.

The instrument panels (Figure 1-17) contain theflight instruments, engine instruments, and avi-onics panel. The airspeed indicator and othermiscellaneous system gages which provide limi-tations markings are shown in Figure 1-18.

The engine instruments (Figures 1-19 and1-20) are mounted in a vertical double rownext to the avionics panel. The avionics panelin the center contains the nav/comm controlsand weather radar unit.

Extending aft from the center subpanel is theengine control quadrant and pedestal (Figure1-21). Engine controls, flap control handle, rudderand aileron trim knobs, and pressurization controlsare mounted on this pedestal. The flight directorand autopilot systems are also installed here.

On the right side panel next to the copilot is themain circuit-breaker panel (Figure 1-21), wherethe majority of the system circuit breakers arelocated. The static air selector handle is mountedjust below the circuit-breaker panel.

Just below the instrument panel are the pilot’s(left) and copilot’s (right) subpanels (Figure1-22). Aircraft system controls, engine switches,master switches, and landing gear controls arelocated on these subpanels.

OFF

FLIGHTHOURS 1/10

GYROSUCTION

INCHES OF MERCURY

3 64 5

COPILOTAIR

PULLON

PNEUMATICPRESSURE

0 20

10

PSI

01500

2000

1000

OFF

CLOSED

OPEN

INCR

ENVIRONMENTALCABIN

HIGH

AUTO

LO

DECR

INCR

CABIN TEMP

CABIN TEMP MODE

OFF

LEFT RIGHT

BLEED AIR VALVES

MANUALTEMP

VENTBLOWER

NO SMOKE& FSB

MANHEAT

AUTO

MANCOOL

ELECHEAT

GND MAX

NORM

CABINAIR

PULLDECREASE

LANDINGGEAR

STALLWARN

ICE PROTECTIONWSHLD ANTI-ICE

NORMAL

SURFACEDEICE

SINGLE

DOWNLOCKREL

PITOT

LANDING TAXI ICE NAV RECOG BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

OFF

OFF

LIGHTS LIGHTS

LDG GEAR CONTROL

MANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

LEFT RIGHT

RELAY TEST

OFF

OFF

PROP FUEL VENT

UP

NOSE

L R

DN

HI

PARKING BRAKE

ENGINE ANTI-ICE

AUTOFEATHER

ON

ON

GENRESET

OFF

ON

OFF

STARTER ONLY

ON

MAIN

OFF

OFF

AVIONICSMASTER PWR

NORM

ACTUATORSSTANDBY

OFF

OFF

PROP GOVTESTARM

TEST OPEN

BUS SENSERESET

GEN TIESMAIN CLOSE

LEFT RIGHT

LEFT RIGHT

TEST

OFF

ARM

OFF

IGNITION ANDENGINE START

ENG AUTOIGNITION

OFF - RESET OXYGENMASK

EXT PWR MICNORMAL

PILOTAIR

PULLON

DEFROSTAIR

PULLON

NO 2

INVERTERNO 1

- +GYROSLAYING

SLEW MODE

COLLINS

INSTANTANEOUSX1000

1

1

3

3

2

2

0

.5

.5ADF

ADF

NAV NAV

COLLINS

- +GYROSLAYING

SLEW MODE

COLLINS

COLLINS

CH SEL PWR

L RINSTANTANEOUS

1

1

3

3

2

2

0

.5

.5

D

HDG

63

E

NW

S

12

3330

2421

15

COURSE

OFF

FSB

BRIGHT

DIM

OFF

BAR

OFF

OFF

BAT GEN 1 GEN 2

2

TEST ERASE

HEADSET600 OHMS

COCKPIT VOICE RECORDER

R ENG ICE FAIL

R ENG ANTI-ICE

PRESS

TO TEST

PRESS TO TEST

MASTERWARNING

PRESS TO RESET

MASTERCAUTION

PRESS TO RESET

EFIS

HORNON

TEST

OFF

SILENCE

AUX POWER

EADI/EHSIDIM

ELAPSEDTIME

DG

DG

DC

FAST

ERECT

220

260

200

180160 140

120

100

4060

AIRSPEEDKNOTS

AIR

10

10

20 20

10

10

OXYGENMASK

MICNORMAL

220

260

200

180160 140

120

100

4060

AIRSPEEDKNOTS

L R

DC

VOL VOL

DIM

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 BCNDME1 12 2 2 ADF1

COMM

COPILOT AUDIO OFF

NAV MKR BCNDME1 12 2 2 ADF1

BOTH RANGE

VOICE BOTH

COMM 1

NAV 1

TRANSPONDER

COMM 2

NAV 2

MKR BCN1 & 2

V OL

COMM 1

COMM 2

CABIN

V OL

COMM 1

COMM 2

CABIN

PUSHON/OFF

GNDCOMMPWR

ANNPUSH BRT25,0 00

SET ALTITUDE

SIDE-TONE

INTPHSENS

ENCDALTM 1

ENCDALTM 2

COMOFF

ONSTO

TESTACT

SQOFF

Collins

MEM MEM

XFR

NAVOFF

ONSTO

TESTACT

HLD

Collins

MEM MEM

XFR

ATCSTBYOFF

ONIDENT

TESTPRE

ALT

Collins

2

1

ADFANTOFF

ADFSTO

TEST ACT

TONE

Collins

MEM MEM

XFR

NAVOFF

ONSTO

TESTACT

HLD

Collins

MEM MEM

XFR

COMMOFF

ONSTO

TESTACT

SQOFF

Collins

MEM MEM

XFR

AVIONICS BY

BENDIX/KINGKLN 90B TS0

RADIO CALLN90KA

GPS APR

ARM ACTV

GPS CRS

OBS LEG

09

8 2

7

5

329.92

1

46

1 0 0

1013

100 FEET

MILLBARS

IN MG

CRSR CRSR

MSG ALT CLR ENTD

NAVFPL

MODETRIP

CALCSTAT

SETUPOTHER

NAVD/T

ACTVREFCTR

APTVORNDBINT

SUPL

PULLSCAN

GPS

PUSHONBRT

PRESENT POS KBEC CLR 125.00 - - - - - - +FR REF: KICT - - - - ,- NM GRND 121.70 - - * - - : - - ' TWR 126.80 - - - * - - : - - ' CTAF 126.80NAV 2 ENR-LEG APT+4

ALT,765

PROP SYN

OFF

ON

COLLINS

10

HDG

APYD

GPS ALTARM

10

DH 2000

34.5NMGSPLIN

CRS

315

E

NW

S

12

63

3330

2421

15

MSG

GPS

COLLINS

WXR-270

MIN

TGTMAX

GAIN

MAP

WX

TST

OFF

SBY

HOLD

0

TILT

+10

-10

10

200

100

50

25

INTCollins

25

10

15

5

STBY

PULLSTABOFF

20

7

6

9 10

12

245

8 ITTSTART

˚C X 1006

9 10

12

245

8 ITTSTART

˚C X 100

16

14

6

4

2

12 FTLB X 100

0

810

TORQUE 0 16

14

6

4

2

12 FTLB X 100

0

810

TORQUE 0

PROP2423

2221

201918 1716

1514

13

10

50

25PROP24

2322

2120

1918 171615

1413

10

50

25

TURBINE

%RPM

110

100

9080 50

40

30

20

0

70 60

.0TURBINE

%RPM

110

100

9080 50

40

30

20

0

70 60

.0

6

0

12

3

4

5FUEL FLOW

PPH X 100

6

0

12

3

4

5FUEL FLOW

PPH X 100

120

80

40

0 0

50

100150

200

˚C

OIL

PSI

120

80

40

0 0

50

100150

200

˚C

OIL

PSI

AUXON

AUXARM

AUX TEST

STALL

WARNING

000

ALT

VERTICALSPEED

CHAN 1

CHAN 2

ADF

1 NM 10

NAV

VERTICALSPEED

X 1000

500

OXYGENSUPPLY PRESSURE

MADE IN USA

PSIUSE NOOIL

LEFT RIGHT

PUSHIN

V

V

V

V

V

35K

15K

COLLINS

V

4

21

56

7

89 0

3

ALT

ENCODING

1015

1010

1005

298

299IN Hg

mb

100 FEET

4

6

2 41

1

0.5

.5

2CABIN CLIMBTHDS FT PER MINFLAPS

DOWN

APPROACH

UP20

60

80

O MI DH

MASTER SWITCH

ALTALERT

63

E

NW

S

12

3330

2421

15

0

5

10

1520

25

30

35

1

2

34

5

6

1,000 FT

80

80

Figure 1-17 Instrument Panels


220

260

200

180160 140

120

100

4060

AIRSPEEDKNOTS

80

PNEUMATICPRESSURE

0 20

10

PSI

GYROSUCTION

INCHES OF MERCURY

3 64 535K

15K

10 20

0 30PROP AMPS

64

122

0 14

108

QTY

FUEL

LBS X 100

MAIN TANKONLY

0

5

10

1520

25

30

35

1

2

34

5

6

1,000 FT

FLAPS

DOWN

APPROACH

UP20

60

80

Figure 1-18 Instrumentation


Figure 1-19 Engine Instruments—Prior to LJ-1361


Figure 1-20 Engine Instruments—LJ-1361, LJ-1363, and After


TEST ERASE

HEADSET600 OHMS

COCKPIT VOICE RECORDER

YAWENG

APENG

SR

I/2Ø

DN

L

UP

R

HDG NAV APPR B/C CLIMB

ALT ALT SEL VS IAS DSC

CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

TEST OFF

PRESS

OFF

NAV DATA

GSPTTG

ET

COURSETIMER

EFIS

POWER TEST

PREACT

XFR

S/S

DH

MAP MAP

ARC ARCHSI

WX

TST

CRSSEL

WARNING DE-PRESSURIZE CABINBEFORE LANDING

CABINALT

CABINALT

RATE

1000FT

MIN

MAX

UP

DN

0

5

CAUTIONREVERSE

ONLY WITHENGINESRUNNING

LIFT

LIFTIDLE

GDFINE

ELEVATOR

TRIM

REVERSE

FRICTIONLOCK

AILERON TABLEFT RIGHT

FRICTIONLOCK

PROP

CONDITION

UP

FEATHER

RUDDER TABLEFT RIGHT

F L A P

CMPST

NORMAL

DN

10

2

34

5

6

78

910

-1

01

10

2022

18

26

16

12 14

24

ACFT ALT1000 FT

FUELCUTOFF

TEST

+

5

UP

10

APPROACH

DOWN

PUSH PUSHHDG CRS

SY N CDIRE CT

4

6

2 41

1

0.5

.5

2CABIN CLIMBTHDS FT PER MINFLAPS

DOWN

APPROACH

UP20

60

80

5

31

5

31

5

3

1

5

310

GO AROUND

GEAR HORNSILENCE

0

5

10

1520

25

30

35

1

2

34

5

6

1,000 FT

TRIM HDG NAVARM DR APPRARM B/C VNAV 1/2 0

YAW DIS AP

YAWDIS ALT ALTARM VS GSARM IAS GA DSC CLM APDIS

PILOT STATICAIR SOURCENORMAL ALTERNATE

SEE FLIGHT MANUAL PERFORM-ANCE SECTION FORINSTR CAL ERROR

5

20

5

5

5

1

1

55 5

5

1

5

5

5

5

5

5

10

5

15

15

15

5

10

5

25

5

3

1

1

5

225

2

71/2

2

32

2

3

5

2

WARN

WARNLANDING

GEAR

IND &CONTROL

TURN &SLIP

TURN &SLIP

ENCDALT

LDGFUELVENT

BLEEDAIR

CONTROL

BUSTIE

POWER

BUSTIE

OUTSIDEAIR

ENCDALTM

& ENGINSTR &SIDE PNL

NO SMKFSB &

CHIPDETR

IGNITORPOWER

FEATHER

FUELCONTROL

HEAT

ANN

POWER STALL LEFTLEFT LEFT AUTOLEFT AVIONICS

AVIONICS COMM NAV COMPASS

IND

BAT

BAT

GEN 1

GEN 2

BAT

BAT

GEN 1

GEN 2

FLAP

PILOT PILOT TEMP INDIRECT FLOOD ANN

RADAR XPNDR DME

DME

RM 1RADIO

ALTM PHONERESET NO 2NO 2

LEFT TEMP PRESS

CONTROL DEICE WIPER CONTROLPILOT

CIGAR LEFT LEFT NO 1 AURAL PILOT ADFCOPILOT

RMI GPS

FURNISHING

WEATHER

LIGHTS

ELECTRICAL

ENVIRONMENTAL

FLIGHT

WARNINGS ENGINES AVIONICS

RIGHT WARN AUDIOLIGHTER

RIGHT RIGHT NO 2

TOILET

AUDIO NO 2

ICE TAXI SURF WSHLD DSPDSPL

PRCSR EADI EHSI

MOTOR ALERT SYNCANTI-ICE

FLAP COPILOT

ELEVRUDDER COPLTSUB PNL

READING

AVIONICS XPNDR

TRIMCOPILOT BOOST OVHD &CONSOLE

INSTR CABIN AVIONICS GEN IND

FLTINSTR

RIGHTRIGHT RIGHT STBY NO 3 NO 1 NO 1 NO 1

ALT PLT FLT AVIONICS PROP LEFTLEFT LEFT AVIONICSNORMAL COMM NAV COMPASS

NO 2 NO 2 NO 2 NO 2

IND RIGHTRIGHT RIGHT TESTRIGHT NO 1 NO 1 NO 1NO 1

MASTERENGINSTR

POWER

OILPRESSWARN

EFISFANS

STARTCONTROL

BUSTIE

POWER

5

5

571/25 71/2 55

5 71/2 5 5

71/2

71/2

71/2

71/2

55

5

71/2

71/2

INV PWRSELECT

5

71/2 71/2

71/27 1/2

5

5

71/2

71/2

71/2

5 5

5

25

5

5

2

71/2

71/2 71/2

PROPGOV

52

2 15

VOICE CABIN

RCDR AUDIO

DISPL

AFCS EFIS

2

ADF

10

AP

SERVO NO 1AUX BAT

NO 2

RADIO

3

5 5

Figure 1-21 Pedestal and Right Side Panel


OFF

LANDINGGEAR

STALLWARN

ICE PROTECTIONWSHLD ANTI-ICE

NORMAL

SURFACEDEICE

SINGLE

DOWNLOCKREL

PITOT

LANDING TAXI ICE NAV RECOG BEACON STROBE

TAILFLOOD

HYD FLUIDSENSOR

HD LTTEST

GEARDOWN

OFF

OFF

LIGHTS LIGHTS

LDG GEAR CONTROL

MANUAL

PILOT COPILOT

LEFT RIGHT

LEFT RIGHT

LEFT RIGHT

RELAY TEST

OFF

OFF

PROP FUEL VENT

UP

NOSE

L R

DN

HI

PARKING BRAKE

ENGINE ANTI-ICE

AUTOFEATHER

ON

ON

GENRESET

OFF

ON

OFF

STARTER ONLY

ON

MAIN

OFF

OFF

AVIONICSMASTER PWR

NORM

ACTUATORSSTANDBY

OFF

OFF

PROP GOVTESTARM

TEST OPEN

BUS SENSERESET

GEN TIESMAIN CLOSE

LEFT RIGHT

LEFT RIGHT

TEST

OFF

ARM

OFF

IGNITION ANDENGINE START

ENG AUTOIGNITION

OFF - RESET OXYGENMASK

EXT PWR MICNORMAL

PILOTAIR

PULLON

DEFROSTAIR

PULLON

NO 2

INVERTERNO 1

OFF

BAT GEN 1 GEN 2

2

PROP SYN

OFF

ON

LEFT RIGHT

MASTER SWITCH

FLIGHTHOURS 1/10

GYROSUCTION

INCHES OF MERCURY

3 64 5

COPILOTAIR

PULLON

PNEUMATICPRESSURE

0 20

10

PSI

01500

2000

1000

OFF

CLOSED

OPEN

INCR

ENVIRONMENTALCABIN

HIGH

AUTO

LO

DECR

INCR

CABIN TEMP

CABIN TEMP MODE

OFF

LEFT RIGHT

BLEED AIR VALVES

MANUALTEMP

VENTBLOWER

NO SMOKE& FSB

MANHEAT

AUTO

MANCOOL

ELECHEAT

GND MAX

NORM

CABINAIR

PULLDECREASE

OFF

FSB

BRIGHT

DIM

OFF

BAR

OFF

OXYGENMASK

MICNORMAL

500

OXYGENSUPPLY PRESSURE

MADE IN USA

PSIUSE NOOIL

35K

15K

Figure 1-22 Pilot’s and Copilot’s Subpanels


The annunciator system (Figure 1-23) consists ofan annunciator panel centrally located in theglareshield, an annunciator panel dimming con-trol, a press-to-test switch, and a fault warninglight. The annunciators are word-readout type.

Whenever a condition covered by the annuncia-tor system occurs, a signal is generated, and theappropriate annunciator is illuminated.

The illumination of a green or yellow annuncia-tor light will not trigger the fault warning system,but a red annunciator will actuate the fault warn-ing flasher. After LJ-1353, a yellow light willtrigger a MASTER CAUTION flasher.

NOTE: CHIP DETECT - Lights red on the C90A DC GEN - Lights Red (Prior to LJ-1353 and after) NO FUEL XFR - Lights Red (Prior to LJ-1353) OIL PRESS - Optional prior to LJ-1353

Figure 1-23 Annunciators


In the overhead area, between the pilot and copi-lot, is the lighting control panel (Figure 1-24).The various rheostat controls for the flight deckand instrument lighting are mounted on thispanel, convenient to both pilot and copilot.

Also mounted on this panel are the windshieldwiper control, the generator load and voltagegages, the deice amps gage, and the invertermonitoring gage. Certain operation limitationsare also placarded on this panel.

C90A INSTRUMENT PANEL/AVIONICSThe C90A panel features as standard equipmenta Collins Pro Line II avionics package, includingan EFIS HSI (EHSI-74).

The airplane is approved for single-pilot opera-tion; however, a full set of copilot’s flightinstruments are installed as standard equipment.

Optionally available are Bendix/King GoldCrown and Silver Crown avionics packages. TwoRNAV/LORAN systems are available as options:the Bendix/King KLN 88 (Figure 1-25) and theFoster LNS616B.

CONTROL SURFACESThe King Air C90A and C90B are equipped withconventional ailerons, elevators, and rudder (Fig-ure 1-26). The control surfaces are pushrod- andcable-operated by conventional dual controls inthe flight deck.

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

THE OPERATING LIMITATIONS STATED IN THE FORM OF PLACARDS, MARKINGS AND MANUALS.

NO ACROBATIC MANEUVERS INCLUDING SPINS ARE APPROVED

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

CAUTIONSTALL WARNING IS INOPERATIVE WHEN MASTER SWITCH IS OFF

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

AND/OR ELECTRIC HEAT IS ON

PILOTFLIGHTINSTROFF

BRT

ENGINEINSTROFF

OFF

BRT

AVIONICSPANEL

OFF

BRT

OVHD PED& SUBPANEL

OFF

BRT

SIDEPANEL

OFF

BRT

COPILOTGYROINSTROFF

MASTERPANELLIGHTS

ON

BRT

COPILOTFLIGHTINSTROFF

BRT

DO NOT OPERATEON DRY GLASS

WINDSHIELDWIPER

OFFPARK SLOW

FAST

OVERHEADFLOOD

OFF

INSTRUMENTINDIRECT

OFF

BRT BRT

MAXAIRSPEED KNOTS

GEAR EXTENSION 182GEAR RETRACT 163GEAR EXTENDED 182

APPROACH FLAP 184FULL DOWN FLAP 148 MANEUVERING 169

40 60 8020

0 100DC % LOAD

40 60 8020

0 100DC % LOAD

10 20

0 30DC VOLTS

INSTRUMENTEMERG LIGHTS

ON

OFF

VOLTMETERBUS SELECT

GENERATOR

CTR TPLFED

BATEXTPWR

10 20

0 30PROP AMPS

LEFT RIGHT

FREQ400

390

110 120

130100

410

380 420

AC VOLTS

23033

FO

R0

3060

90120 150 180 210

24027

030

033

0

COMPASS CORRECTIONCALIBRATE WITH

RADIO ON

ST

EE

Figure 1-24 Overhead Light Control Panel


Any time the airplane is parked overnight or inwindy conditions, the rudder gust pin and controllocks should be installed to prevent damage tothe control surfaces and hinges or to the controls(Figure 1-27). Two items require particular atten-tion: the parking brake handle mounted justunder the left corner of the subpanel, and the rud-der gust lock bar mounted between the pilot’srudder pedals.

Before towing the airplane, the parking brakemust be released (brake handle pushed in), andthe rudder gust lock bar must be removed frombetween the rudder pedals. Serious damage to thetires, brakes, and steering linkage can result ifthese items are not released.

Figure 1-25 Avionics Installation (C90A)

Figure 1-26 Flight Control Surfaces


Figure 1-27 Flight Control Locks

Figure 1-28 Tiedowns


TIEDOWN AND SECURINGWhen the airplane is parked overnight or duringhigh winds, it should be securely moored withprotective covers in place (Figure 1-28). Placewheel chocks fore and aft of the main gearwheels and nosewheel. In severe conditions theparking brake should be set.

Using the airplane mooring points, tie the air-plane down with suitable chain or rope. Installthe control surface lock, and be sure the flaps areup. Secure the propellers with appropriatetiedown boots (one blade down) to prevent wind-milling (Figure 1-29).

This airplane has free spinning propellers thatcould be hazardous if not restrained. Windmillinggears and bearings without lubrication is notgood practice. When there is blowing dust orrain, install the pitot mast cover, as well as theengine inlet and exhaust covers.

Two items require particular attention: the park-ing brake handle mounted just under the leftcorner of the pilot’s subpanel and the rudderpedal gust lock. Before towing the airplane, theparking brake must be released (brake knobpushed in) and the rudder gust lock removed.Serious damage to tires, brakes, and steeringlinkage can result if these items are not released.

Figure 1-29 Propeller Boots


TAXIINGThe ground turning radii are predicated on theuse of partial braking action, differential power,and the nosewheel fully castored in the directionof the turn (Figure 1-30). Locking the insidebrake can cause tire or strut damage. When turn-ing the airplane, if the wingtip clears obstaclesthe tail will also. The turning radius for thewingtip is 35 feet 6 inches. While turning, thepilot should be aware of vertical stabilizer clear-ance, which is 14 feet 3 inches.

When taxiing, turning, and starting the engines,there is an area directly to the rear of the engineswhere the propeller windstream can be hazardousto persons or parked airplanes (Figure 1-31).While the velocities and temperatures cannot beaccurately measured, reasonable care should betaken to prevent inc idents wi th in thesedanger areas.

SERVICING DATAThe “Handling, Servicing, and Maintenance”section of the POH outlines to the Owner andOperator the requirements for maintaining theKing Air C90A and C90B in a condition equalto that of its original manufacture. This infor-mation sets time intervals at which the airplaneshould be taken to a Beechcraft Aviation Centerfor periodic servicing or preventive mainte-nance. All limits, procedures, safety practices,time limits, servicing and maintenance require-ments contained in the POH are mandatory.This section of the POH includes a ConsumableMaterials chart which lists approved and recom-mended materials for servicing the airplane(Figure 1-32). The “Servicing Schedule andLubrication Schedule” lists and illustrates ser-vicing points and materials required.

Figure 1-30 Turning Radius

Figure 1-31 Danger Areas


Figure 1-32 Servicing Data


PRODUCT SUPPORTBeech Aircraft has established service facilitiesthroughout the world, which are fully equippedand professionally staffed to provide total sup-port for the Super King Airs.

These facilities are listed in the Beechcraft Qual-ity Service Center Directory (USA) and theInternational Service Facility Directory, copies ofwhich are provided to each new Beechcraftowner. To support this worldwide service organi-zation, Beech Aircraft, through its Parts andEquipment Marketing Wholesalers and Interna-tional Distributors, provides a computer-controlled parts service that assures rapid ship-ment of equipment on a 24-hour basis.

PREFLIGHT INSPECTIONThe preflight inspection procedure in the POHhas been divided into five areas, as shown inFigure 1-33. The inspection begins in the flightcompartment, proceeds aft, then moves clock-wise around the aircraft, discussing the leftwing, landing gear, left engine and propeller,nose section, etc.

Exterior Inspection1. Cockpit check

2. Left wing, landing gear, engine, nacelle andpropeller

3. Nose section

4. Right wing, landing gear, engine, nacelle andpropeller

5. Empennage and tail

Figure 1-33 Exterior Inspection

Revision .02


2-i

CHAPTER 2

CONTENTS

Page

ELECTRICAL POWER SYSTEMS

INTRODUCTION ...................................................................................................................

2-1

GENERAL ..............................................................................................................................

2-1

Battery and Generator .....................................................................................................

2-1

DC Power Distribution....................................................................................................

2-8

Bus Tie System.................................................................................................................

2-8

Bus Isolation..................................................................................................................

2-11

Load Shedding...............................................................................................................

2-12

Battery ...........................................................................................................................

2-12

Starter/Generators..........................................................................................................

2-13

AC Power Distribution..................................................................................................

2-16

External Power ..............................................................................................................

2-18

Avionics Master Power .................................................................................................

2-19

Circuit Breakers.............................................................................................................

2-19


2-iii

Figure Title Page

ILLUSTRATIONS

2-1

Electrical System Component Locations.................................................................

2-2

2-2

Basic Electrical Symbols .........................................................................................

2-3

2-3

Pilot and Copilot Subpanels.....................................................................................

2-4

2-4

Overhead Meter Panel ..............................................................................................

2-5

2-5

Electrical System Buses and Feeders (1 of 2) .........................................................

2-6

2-6

Right Side and Fuel Management Circuit Breaker Panels ......................................

2-9

2-6A

Right Side and Fuel Management Circuit Breaker Panels............................................

2-10

(LJ-1361, LJ-1363, and After)

2-7

Battery Installation.................................................................................................

2-12

2-8

Starter/Generator Installation.................................................................................

2-13

2-9

Simplified Inverter Schematic ...............................................................................

2-16

2-10

Inverter Schematic—Condition 1 ..........................................................................

2-17

2-11


2-17

2-12


2-18

2-13

Avionics Master Power Schematic ........................................................................

2-20

2-14

Power Distribution Schematic ...............................................................................

2-22

2-15

Power Distribution—Battery OFF.........................................................................

2-23

2-16

Power Distribution—Battery ON ..........................................................................

2-24

2-17

Power Distribution—Battery ON (Generator Ties Manually Closed) ..................

2-25

2-18

Power Distribution—Right Engine Start (Generator Ties Normal) ......................

2-26

2-19

Power Distribution—Right Generator ON ............................................................

2-27

2-20

Power Distribution—Left Engine Cross-Start (Right Engine Running) ................

2-28

2-21

Power Distribution—Both Generators ON............................................................

2-29

2-22

Power Distribution—Both Generators ON (Generator Ties Open) ......................

2-30

2-iv


2-23

Bus Sense Test—Both Generators ON .................................................................

2-31

2-24

Both Generators Failed—Load Shedding .............................................................

2-32

2-25

Right Generator Bus Short—Bus Isolation...........................................................

2-33

2-26

Center Bus Short—Bus Isolation ..........................................................................

2-34

2-27

Triple-Fed Bus Short—Bus Isolation....................................................................

2-35

2-28

Power Distribution—External Power

(External Power and Battery Switches ON)...........................................................

2-36


2-1

CHAPTER 2ELECTRICAL POWER SYSTEMS

INTRODUCTION

Familiarity with, and an understanding of, the airplane electrical system will ease pilot work-load in normal operations in case of an electrical system or component failure. The pilot shouldbe able to locate and identify switches and circuit breakers quickly, and should also be familiarwith appropriate corrective actions in emergency situations.

GENERAL

The Electrical System section of the trainingmanual presents a description and discussion ofthe airplane electrical system and components.The electrical system is discussed to the extentnecessary for the pilot to cope with normal andemergency operations. The location and purposeof switches, indicators, and circuit breakers,along with DC and AC generation and distribu-tion are described. This section also includessome of the limits of, and possible faults with,systems or components.

BATTERY AND GENERATOR

The airplane electrical system is a 28-VDC(nominal) system with the negative lead of eachpower source grounded to the main airplanestructure. DC electrical power is provided by one34-ampere-hour, air-cooled, 20-cell, nickel-cad-mium battery (airplanes prior to LJ-1534) or one42-ampere-hour, sealed, lead-acid battery (air-planes LJ-1534 and after), and two 250-amperestarter/generators connected in parallel. Basicelectrical symbols are shown in Figure 2-2.

#1 S

ERVO

SYSTEM

BATT HOT

BAT OFF

AC

GEN

#1 D

C

GEN

#1 E

NG

OIL PL

2-2


Figure 2-1 Electrical System Component Locations


2-3

This system is capable of supplying power to allsubsystems necessary for normal operation of theairplane. The battery and generator switches onthe pilot’s left subpanel are used to controlpower from the battery and generators into theairplane electrical system.

The battery is always connected to the hot batterybus (Figure 2-21). Both are located in the rightwing center section. Operation of equipment onthe hot battery bus does not depend on the posi-tion of the battery switch. The battery switch, onthe pilot’s left subpanel, closes a battery bus tieand a battery relay which connect the battery tothe rest of the electrical system.

The generators are controlled by individual gen-erator control panels which allow constantvoltage to be presented to the buses during varia-tions in engine speed and electrical loadrequirements. The load on each generator is indi-cated by left and right loadmeters located on theoverhead meter panel (Figure 2-4). A normal sys-tem potential of 28.25 ±0.25 volts maintains thebattery at full charge.

This airplane utilizes a multi-bus system. Themain buses are the left and right generator buses,center bus, triple-fed bus, and battery emergencybus. Switches in the cockpit which receive powerfrom the center and triple-fed buses are identifiedby a white ring on the panel around the switch.

Electrical loads are divided among the buses asnoted on the Electrical System Buses and Feed-ers chart (Figure 2-5). Equipment on the buses isarranged so that all items with duplicate func-tions (such as right and left landing lights) areconnected to different buses. Circuit breakers onthe same feeder or sub-bus are connected bywhite lines on the right circuit breaker panel faceboard (Figure 2-6).

In normal operation, all buses are automaticallytied into a single-loop system where all sourcessupply power through individual protectivedevices. The triple-fed bus is powered from thebattery and both generator buses. The left andright generators supply power to their respectiveleft and right generator buses.

The center bus is fed by two generator busesand the battery, which automatically connectsthose components whenever the bus ties areclosed. The power distribution schematics(Figures 2-14 through 2-28) show how busesare interconnected.

BATTERY

FUSE

CURRENT LIMITER

(OR ISOLATION LIMITER) THIS ACTSAS A LARGE, SLOW TO OPEN FUSE

DIODE

THE DIODE ACTS AS A ONE-WAY"CHECK VALVE" FOR ELECTRICITY.(Triangle points in direction of power flow. Power cannot flow in opposite direction.)

CIRCUIT BREAKER

SWITCH - TYPECIRCUIT BREAKER

BUS TIE &SENSOR

RELAY OPEN

RELAY CLOSED

NO

RM

ALL

Y

OP

EN

NO

RM

ALLY

C

LOS

ED

Figure 2-2 Basic Electrical Symbols

2-4


Figure 2-3 Pilot and Copilot Subpanels

PILOT'S SUBPANEL (LJ-1063 THRU LJ-1352)

COPILOT'S SUBPANEL (C90A AND B)

PILOT'S SUBPANEL (LJ-1353 AND SUBSEQUENT)


2-5

OVERHEAD METER PANEL (LJ-1534 AND AFTER)

OVERHEAD METER PANEL (LJ-1353 TO LJ-1533)

OVERHEAD METER PANEL (PRIOR TO LJ-1353)

ON

OFF

Figure 2-4 Overhead Meter Panel

2-6


Revision .01

LEFT GENERATOR BUS

(Gen No. 1)Flap MotorFlap Control & IndicatorL Generator Bus (Bus Tie & Meter Indication)Flight Instrument LightsAvionics & Engine Instrument LightsProp Synchrophaser (Opt)L Main Anti-ice (Ice Vane)L Chip DetectorL Engine Fuel Control HeaterR Bleed Air ControlCigarette LighterL Landing LightL Fuel Vent HeatNo. 2 Avionics BusL Generator Control Panel (1)L Generator Field & Sense (1)L Generator LoadmeterPilot Windshield Anti-ice (1)R Standby Anti-ice (Ice Vane)Vent BlowerRotating Beacon LightsTail Flood Lights (Opt)R Firewall ValveR Fuel Boost PumpCrossfeedNo. 1 Inverter Power Control (1)No. 1 Inverter Power Select (1)

(Avionics)*Comm 2ADF 1Copilot AudioAutopilotTransponder 2DME 2VLF/OMEGAWeather RadarPaging AmplifierCheck List (Radar)Data NavVNAVEFIS

CENTER BUS

AVIONICS ANNUNCIATORGENERATORRESETPNEUMATIC SURFACEDEICEWINDSHIELD WIPERTAXI LIGHTICING LIGHTNO. 1 INVERTER POWERCONTROLNO. 2 INVERTER POWERCONTROLPROP DEICE POWERPROP DEICE CONTROLAIR CONDITIONERNORMAL HEAT (ELECTRIC)MAX HEAT (ELECTRIC)LANDING GEAR MOTOR (1)NO. 1 & NO. 2 INVERTERPOWER

RIGHT GENERATOR BUS

(Gen No. 2)Trim Tab (Opt)R Generator Bus (Bus Tie & Meter Indication)Overhead, Subpanel & Pedestal LightsSide Panel LightsCabin Reading Lights & Sign ChimeR Main Anti-ice (Ice Vane)R Chip DetectorR Engine Fuel Control HeaterRudder BoostElectric Toilet (Opt)R Landing LightR Fuel Vent HeatNo. 3 Avionics BusR Generator Control Panel (1)R Generator Field & Sense (1)R Generator Loadmeter (1)Copilot Windshield Anti-ice (1)L Standby Anti-ice (Ice Vane)Furnishing - (Refreshment Bar)R Landing LightsStrobe Lights (Opt)R Pitot HeatStall Warning HeatL Firewall ValveL Fuel Boost PumpCrossfeedNo. 2 Inverter Power Control (1)No. 2 Inverter Power Select (1)

(Avionics)*Nav 2Glideslope 2Transponder 1DME 1Compass 2Flight Director 2ADF 2Marker Beacon 2HFRadar AltimeterRadio TelephoneRMI 1StereoInterphoneFlight Path AdvisoryAir Data EncoderGPSIntegrated Hazard AvoidanceSystemMultifunction Display

Figure 2-5 Electrical System Buses and Feeders (1 of 2)


2-7

Voltage on each bus may be monitored on thevoltmeter (located in the overhead panel) byselecting the desired bus using the VOLTMETERBUS SELECT switch, adjacent to the voltmeter.The electrical system provides maximum protec-tion against loss of electrical power should a

ground fault occur. High current (Hall effect)sensors, bus tie relays and current limiters areprovided to isolate a fault from its power source.The electrical system bus arrangement isdesigned to provide multiple power sources forall circuits.

Landing Gear Warning HornAnnunciator PowerStall Warning SystemL Generator Overheat (CAA)L Starter ControlL Ignitor PowerFire DetectionL Fire Detection (CAA)L Oil Temperature

& Oil PressureL Oil Pressure Warning (CAA)Autofeather System (Opt)L Fuel Flow IndicatorL Pitot HeatLanding Gear ControlPilot Turn & SlipNavigation LightsPilot Encoder & AltimeterAvionics Master ControlInstrument Indirect LightsCabin Fluorescent LightsTriple-Fed Bus

(Bus Tie & Meter Indication)L Firewall Fuel ValveL Boost PumpL Transfer PumpL Fuel Quantity IndicatorL Fuel Pressure WarningCrossfeed Fuel ValveLanding Gear

Position IndicatorAnnunciator Indicator

R Generator Overheat (CAA)R Starter ControlR Ignitor PowerR Fire Detection (CAA)R Oil Temp & Oil PressR Oil Pressure Warning (CAA)Prop Governor TestR Fuel Flow IndicatorL Bleed Air ControlCabin Pressure Loss (CAA)Bus Tie ControlNo. 1 Avionics BusR Firewall Fuel ValveR Boost PumpR Fuel Quantity IndicatorR Fuel Pressure Warning

(Avionics)*Comm 1Nav 1Glideslope 1Radio RelaysCompass 1Flight Director 1Pilot AudioMarker Beacon 1Servo AltimeterRNAVRMI 2

HOT BATTERY BUS

L Engine Fire Extinguisher (1)R Engine Fire Extinguisher (1)RNAV Memory (1)

(Ground Comm Power)Entrance & Aft Dome Lights (1)Stereo (1)Battery Relay Control (1)L Fuel Boost Pump (1)R Fuel Boost Pump (1)Fuel Crossfeed (1)

(1) The circuit breaker in thiscircuit is not accessible tothe pilot in flight.

TRIPLE-FED OR BATTERY BUS

*Optional avionics busing. Check avionics circuit breaker panel orwiring diagram for specific busing configuration.

Figure 2-5 Electrical System Buses and Feeders (2 of 2)

2-8


DC POWER DISTRIBUTION

The DC power distribution system is commonlycalled a “triple-fed” system. In normal operation,all buses are automatically tied into a single loopsystem in which all sources collectively supplypower through individual protective devices.

Three in-flight DC power sources are available:

●

One 24-volt, 34-ampere-hour nickel-cad-mium battery (or one 24-volt, 42-amperehour, lead acid battery for LJ-1534 andlater)

●

Two 28-volt, 250-ampere starter/generators

When the battery switch is turned ON, the bat-tery relay and the battery bus tie relays close(Figure 2-16). Battery power is routed throughthe battery relay to the triple-fed bus, and throughthe battery bus tie relay to the center bus and toboth starter relays. Neither generator bus is pow-ered since the generator bus ties are normallyopen, however, battery power is available to per-mit starting either engine.

After either engine has been started and the gen-erator switch has been moved to RESET, thegenerator control unit (GCU) will bring the gen-erator up to voltage. Releasing the spring-loadedswitch to the center ON position closes the gen-erator line contactor, thereby powering thegenerator bus, and closing both generator tiesautomatically. This action distributes powerthrough the 250-amp current limiters and thegenerator bus tie relays. Generator output willthen be routed through the center bus to permitbattery charging. In addition, the opposite gener-ator bus and triple-fed bus will be powered by thegenerator, supplying 28-VDC power to the fiveprimary airplane buses (Figure 2-19) When bothgenerators are operating, each generator directlyfeeds its respective generator bus.

The generator buses, hot battery bus, and batteryare tied together by the center bus. The triple-fedbus is powered by the battery and each generator

bus through 60-amp limiters and through diodesproviding fault isolation protection between thepower sources.

BUS TIE SYSTEM

The electrical system is protected from exces-sively high current flow by the bus tie system.Three current sensors, consisting of Hall effectdevices and solid-state circuitry, are used to sensecurrent flow through the portion of the circuitbeing monitored. Two bus tie sensors and theirrelays are located between the generator busesand the center bus, and a third is between the bat-tery and the center bus.

With no power applied to the aircraft electricalsystem, all three bus tie relays are open. Whenthe BAT switch is turned ON, hot battery busvoltage energizes the coil circuit of the batterybus tie relay, thereby closing it. This action hasno effect on the generator bus ties.

A similar action occurs when a generator orexternal power is brought on-line. When eithergenerator is brought on-line, voltage from thegenerator control panel energizes the coil circuitof both generator bus tie relays. This switchesvoltage from the L and R GEN TIE OPENannunciators to the relays, causing the annuncia-tors to extinguish and the bus tie relays to close.When external power is brought on-line, the onlydifference is the source of generator bus tie coilvoltage, which is the small pin of the externalpower receptacle. Neither generator or externalpower affect the battery bus tie circuitry unlessthe battery switch is also turned ON.

Activation of an internal, solid-state switchwithin the sensor by a current of at least 275 ±5amperes will open the coil circuit of the relay,causing it to deenergize and open the associatedbus tie relay. The coil circuit of the bus tie relayis latched open to prevent the bus tie relay fromclosing. Deenergizing the bus tie relay will illu-mina te the appropr ia te BUS TIE OPENannunciator. When the bus tie relay has beenopened by excessively high current flow through


2-9

Figure 2-6 Right Side and Fuel Management Circuit Breaker Panels (LJ-1063 Thru LJ-1360, LJ-1362

2-10


(LJ-1361, LJ-1363 and After)Figure 2-6A Right Side and Fuel Management Circuit Breaker Panels


2-11

the Hall effect sensor (i.e. a bus fault), it can onlybe reset by momentarily activating the BUSSENSE switch on the pilot’s left subpanel toRESET. The Hall effect sensors are unidirec-tional. They only sense overcurrent in thedirection of the arrow on the symbol.

Two switches located on the pilot’s left subpanelcontrol the bus tie system. One switch, placardedBUS SENSE - TEST - RESET, is spring loadedto the center (NORM) position. Momentarilyactivating it to TEST connects bus voltage to allthree current sensor test circuits (Figure 2-23).This voltage simulates the condition resultingfrom a high current through each bus tie relay.The solid state switches of each sensor are thusactivated to de-energize (open) their respectiverelays, thereby opening the bus tie relays andactivating the annunciator readouts. Once acti-vated, the test circuitry latches the bus ties open,preventing their automatic closing.

Current sensor reaction time is approximately0.010 seconds for the generator current sensorsand 0.012 seconds for the battery current sen-sor. Once activated, the relays latch open, andreaction time for the system is limited to reac-tion time for the relays. Therefore, onlymomentary activation of the TEST switch isrequired. Prolonged activation of this switchwill damage or destroy the sensor modules andshould be avoided.

Momentary activation of the switch to RESETpowers the coil of the bus tie relays, unlatchingthe test circuits and, permitting the bus ties toenergize (close). Voltage is transferred from theannunciator readouts to the coils, closing the bustie relays. Since high-current sensing is latchedout when the switch is in RESET, only momen-tary activation is desirable. This preventsaccidental welding of the bus tie relay contactsand/or opening a 250-amp current limiter by abus ground fault.

The second switch on the pilot’s left subpanelcontrols the bus tie system and is placarded GENTIES - MAN CLOSE - NORM - OPEN. Thisswitch must be lifted (lever-lock) to move it fromcenter to OPEN. This switch is spring loaded toMAN CLOSE.

Only the generator bus tie relays may be manu-ally opened or closed with this switch. Manuallyclosing the generator bus tie relays will connectthe generator buses to the center bus and powerto the entire system (Figure 2-17). Momentarilyplacing the switch in CLOSE applies bus voltageto the coil of the generator bus tie relays, com-pletes a latching circuit, activates the MAN TIESCLOSE annunciator and closes the bus tie relays.The latching circuit is completed through thenormally closed contacts of the control relay forthe generator line contactors. A generator bus tierelay cannot be manually closed if a fault openedthe tie; the BUS SENSE switch must be momen-tarily activated to RESET, which resets the tie.

When the generator ties are closed, the GENTIES switch can open the generator bus ties ascertain normal/abnormal procedures may dictate.When the GEN TIES switch is positioned toOPEN, the ground is removed from the relay cir-cuit which allows the relay to spring open.

BUS ISOLATION

Bus isolation is one of the features of the multi-bus electrical system. The two generator busesand the center bus are protected by high-currentsensing (Hall effect) devices. In case of excessivecurrent draw on one bus, the sensors will isolatethe affected bus by opening its bus tie, allowingthe other buses to continue operating as a system.During cross-generator engine starts, the highcurrent sensors and current limiters are bypassedby cross-start relays to allow the required highcurrent flow to pass from the power sources tothe starter generator without causing the bus tiesto open. Battery starts are routed through the bat-tery bus tie, which is desensitized for starting.

A 250-amp current limiter (slow to open fuse) isalso located in the circuitry between the centerbus and each of the generator buses. Since theHall effect devices sense high current in only onedirection, the current limiters provide protectionin the opposite direction. If an overcurrent situa-tion causes a current limiter to open, it also willcause bus isolation.

2-12


The current protection for the triple-fed bus isprovided exclusively by 60-amp current limiters.Triple-fed bus isolation will occur only if allthree of these limiters open.

For typical examples of bus isolation, refer toFigures 2-25 (generator bus), 2-26 (center bus),and 2-27 (triple-fed bus).

LOAD SHEDDING

Load shedding is another highly beneficial fea-ture of the multi-bus electrical system. Theelectrical system will automatically removeexcess loads (generator buses), when the powersource is reduced to battery only. When both gen-erators are off line, the generator bus ties openand the generator bus loads are “shed” (Figure 2-24). The battery will continue to power the cen-ter, triple-fed, and hot-battery buses. If necessary,power to the generator buses can be restored byclosing the generator ties manually with the GENTIES switch (Figure 2-17). When load sheddingoccurs in flight, land as soon as practical, unlessthe situation can be remedied and at least onegenerator brought back on-line.

WARNING

Closing the generator bus ties manu-ally in flight with a loss of bothgenerators will cause the battery todischarge at a faster rate. If it becomesnecessary to close the generator ties inthis situation, they should be openedas soon as possible since batterypower should be conserved. Withoutan operable generator, the battery can-not be recharged in flight. Land assoon as practical.

BATTERY

The nickel-cadmium (Ni-Cad) battery is locatedin the right wing center section in an air-cooledbox (Figure 2-7). The battery relay, charge moni-tor shunt, and air-cooling thermostat are mountedin the battery compartment immediately forwardof the battery. Power to the main electrical buses

is routed from the battery via the battery relayand battery bus tie, which are controlled by theBAT-ON-OFF switch on the pilot’s left subpanel.

The hot battery bus provides power directly to afew aircraft systems (Figure 2-15). These sys-tems may be operated without turning the batteryswitch ON. Care should be taken, however, toinsure that utilization of these systems is minimalwhen the generators are inoperative and/or theaircraft is secured to prevent excessive dischargeof the battery.

The lead acid battery box is not air cooled anddoes not include the charge monitor shunt, or theair-cooling thermostat. A battery charge monitorsystem (airplanes prior to LJ-1534) advises thepilot of battery charge rate. The system is dis-abled on aircraft that have been converted to alead-acid battery. A charge of more than7 amperes, for six or more seconds, will triggerthe yellow BATTERY CHARGE annunciator,indicating excessive charge rate.

Figure 2-7 Battery Installation


2-13

The BATTERY CHARGE annunciator mayoccasionally illuminate for short intervals whenheavy electrical draw items are cycled. For exam-ple, following a battery-powered engine start, thebattery recharge current is very high and causesillumination of the BATTERY CHARGE annun-ciator, thus providing an automatic self-test ofthe battery monitor system.

As the battery approaches a full charge and thecharge current decreases to a satisfactory level.the annunciator will extinguish. This will nor-mally occur within a few minutes after an enginestart, but may require a longer time if the batteryhas a low state of charge, low charge voltage percell (20-cell battery), or low battery temperature.

STARTER/GENERATORS

The starter/generators are dual-purpose, engine-driven units (Figure 2-8). The same unit is usedas a starter to drive the engine during engine startand as a generator to provide electrical powerwhen driven by the engine. A series starter wind-ing is used during starter operation and a shuntfield winding is used during generator operation.The generator shunt field winding is disabledwhen the series starter winding is activated by thestart switch. The regulated output of the genera-tor is 28.25 ±0.25 volts with a maximumcontinuous load of 250 amperes.

In addition to the starter/generators, the generatorsystem consists of control switches, generator con-trol units (GCU), line contactors and loadmeters.

Figure 2-8 Starter/Generator Installation

2-14


Starter power to each individual starter/generatoris provided by the battery, or by the operatinggenerator for cross-starts. The start cycle is con-trolled by a three-position switch, one for eachengine, placarded: IGNITION AND ENGINESTART - LEFT - RIGHT - ON - OFF -STARTER ONLY, located on the pilot’s left sub-panel (Figure 2-3).

Selecting a start switch to either the STARTERonly position or ON activates the starter and dis-ables the respective generator. The starter drivesthe compressor section of the engine through theaccessory gearbox.

During engine starts, the battery is connected tothe starter/generator by the starter relay. With oneengine running and its generator on the line, theopposite engine can by started with power fromthe battery and operating generator through thestarter relay and the cross-start relay. This iscalled a cross-start. Normally one engine isstarted on battery power alone, and the secondengine is cross-started.

During a cross generator start, (Figure 2-20) theoperating generator control panel closes thecross-start relay, bypassing the generator bus,current limiter and bus tie relay. This assures the250-amp current limiter will not open due totransient surges, since the generator would nor-mally provide the current required for the start. Inaddition, while a starter is selected the bus tiesensors are disabled to prevent them from open-ing their respective bus tie relays.

CAUTION

Do not exceed the starter motoroperating time limits of 40 secondsON, 1 minute off, 40 seconds ON, 1minute off, 40 seconds ON, then 30minutes off.

DC Generation

The generator phase of operation is controlled bythe generator switches, located in the pilot’s leftsubpanel, next to the BAT switch under theMASTER SWITCH gang bar (Figure 2-3). Theswitches provide OFF, ON, and RESET capabili-ties. The generating system is self-exciting anddoes not require electrical power from the air-craft electrical system for operation.

Generator operation is controlled through twogenerator control units (GCU) mounted belowthe center aisle floor, that make constant voltageavailable to the buses during variations inengine speed and electrical load requirements.The generators are manually connected to theGCUs by GEN 1 and GEN 2 control switcheslocated on the pilot’s left subpanel. The load oneach generator is indicated by the respective leftand right loadmeters located on the overheadpanel (Figure 2-3).

The generator control units are designed to con-trol the generators and the load shared within2.5 percent.

The generator control units (GCU) provide thefollowing functions:

●

Voltage regulation and line contactorcontrol

●

Overvoltage and overexcitation protection

●

Paralleling/load sharing

●

Reverse-current protection

●

Cross-start relay activation


2-15

Voltage Regulation and Line Contactor Control

The generators are normally regulated to28.25 ±.25 VDC. When the generator controlswitch is held to RESET, generator residual volt-age is applied through the GCU to the generatorshunt field causing the generator output voltageto rise. When the switch is released to ON, the28-volt regulator circuit takes over and beginscontrolling the generator shunt field in order tomaintain a constant output voltage. The voltageregulator circuit varies shunt field excitation asrequired to maintain a constant 28-volt outputfrom the generator for all rated conditions of gen-erator speed, load, and temperature.

When the generator switch is released to ONgenerator voltage is applied to the GCU to enablethe line contactor control circuit. The GCU com-pares the generator output voltage with aircraftbus voltage. If the generator output voltage iswithin 0.5 volts of the aircraft bus voltage, theGCU sends a signal to the line contactor whichcloses and connects the generator to the aircraftbus (Figure 2-21) and closes both generator tiesto connect the center bus and the generator buses.This allows the generator to recharge the aircraftbattery and power all aircraft electrical loads.

During single-generator operation, the GCUopens the line contactor and isolates the inopera-tive generator from its bus.

Overvoltage and Overexcitation Protection

The GCU provides overvoltage protection toprevent excessive generator voltage from beingapplied to the aircraft equipment. If a generatoroutput exceeds the maximum allowable 32-volts, the overexcitation circuits of the GCU willdetect which generator is producing excessivevoltage output and attempting to absorb all theaircraft electrical loads. The GCU overexcitationcircuit will then disconnect the generator fromthe electrical system.

Paralleling/Load Sharing

The paralleling circuit averages the output ofboth generators to equalize load levels. The par-alleling circuits of both GCUs become operativewhen both generators are on the line. The paral-leling circuits sense the interpole windingvoltages of both generators to provide an indica-tion of the load on each generator.

The voltage regulator circuits are then biased upor down as required to increase or decrease gen-erator loads until both generators share the loadequally. The GCUs are designed to balance loadsto within 2.5 percent.

Reverse-Current Protection

Reverse-current protection is provided by theGCU. When a generator becomes underexcitedor cannot maintain bus voltage, i.e., low genera-tor speed during engine shutdown, it will begin todraw current (reverse current) from the aircraftelectrical system. The GCU senses the reversecurrent by monitoring the generator interpolevoltage and opens the line contactor to protectthe generator.

Cross-Start Relay Activation

During cross-start, the operating generator helpsto start the second engine. The cross-start relayon the operating generator circuit closes to allowstarting current to bypass the generator bus, cur-rent limiter, and bus tie relay. The current flowsthrough the center bus, to the Hall effect sensoron the opposite generator bus.

During start, the Hall effect sensors are disabled,so no bus isolation takes place. The current isrouted to the starter physically between the Halleffect sensor and the bus tie relay, so if the bus tieopened, it wouldn’t effect engine start. The cur-rent is then made available to the start relay forengine start.

2-16


AC POWER DISTRIBUTION

AC power for the avionics equipment and theAC-powered engine instruments is supplied bytwo inverters (Figure 2-9). Either one may beused at the pilot’s discretion through the inverterselector switch. Each inverter provides two levelsof power: 115 volts, 400 Hz for the avionicsequipment and 26 volts, 400 Hz for the applica-ble engine instruments, and some avionics.Output of the standard inverter is rated at250 volt-amps. An optional inverter is rated at300 volt-amps.

The inverters are installed in the wing, immedi-ately outboard of each nacelle. Inverter operationis controlled by the INVERTER NO. 1 - OFF -NO. 2 select switch on the pilot’s left subpanel.Selection of either inverter actuates the inverterpower relay installed nearby to supply it with DCpower. An inverter select relay provides the nec-essary switching to permit the operating inverterto supply 26 VAC avionics and instrument power,and 115 VAC avionics and test jack power. Theinverter select relay is energized when the num-

ber one inverter is selected. It is de-energizedwhen the inverter switch selects either OFFor NO. 2.

Dual sources of DC input power are provided foreach inverter. The power select relay for eachinverter is automatically selected to provideinverter power from the adjacent generator bus,or from the center bus if the generator bus is de-energized (Figures 2-10, 2-11, and 2-12). Whenbattery power is applied to the center bus prior toengine start (Figure 2-16), inverter power isrouted through a circuit breaker and the nor-mally-closed contacts of an inverter power-selectrelay to the power relay of each inverter. Whenthe generator buses are powered (Figure 2-21),voltage is also routed through a circuit breaker onthe copilot’s circuit breaker panel to the coil ofeach inverter power-select relay, causing theinverter to be powered by its generator bus.

During normal operation, an inverter bus selectrelay is energized and power is supplied from thegenerator bus. Should a fault occur that wouldinterrupt power to that bus, the bus select relay

INVERTER

CENTER BUS(NO 1 INV FEEDER)

40

5

7.540

10

5 5

5 5

5 5

5

5

1011526

115 vac 400 Hz

26 vac 400 Hz

11526

1 2

LH GEN BUS RH GEN BUS


40

7.5 40

ANN. IND.

TESTJACK

SELRELAY

LHNO. 1INV

RHNO. 2INV

SELSW

NO. 1 INVCONTROL

NO. 2 INVCONTROL

Figure 2-9 Simplified Inverter Schematic


2-17

INVERTER

28VDC26VAC115VAC


40

5

7.540

10

5 5

5 5

5 5

5

5

1011526

115 vac 400 Hz

26 vac 400 Hz

11526

1 2



40

7.5 40

ANN. IND.

TESTJACK

SELRELAY

LHNO. 1INV

RHNO. 2INV

SELSW

NO. 1 INVCONTROL

NO. 2 INVCONTROL

Figure 2-10 Inverter Schematic—Condition 1

INVERTER

28VDC26VAC115VAC


40

5

7.540

10

5 5

5 5

5 5

5

5

1011526

115 vac 400 Hz

26 vac 400 Hz

11526

1 2



40

7.5 40

ANN. IND.

TESTJACK

SELRELAY

LHNO. 1INV

RHNO. 2INV

SELSW

NO. 1 INVCONTROL

NO. 2 INVCONTROL


2-18


would de-energize and inverter input powerwould be taken from the center bus of the air-plane, precluding the possibility of loss of aninverter due to failure of the generator bus.

Inadequate inverter output power is indicatedby the illumination of the INVERTER annun-ciator. This could happen due to loss of inputpower, or an inverter failure. Other indicationsof inverter loss would be erratic behavior, orloss of AC powered instruments (torque gage),or AC avionics.

During inverter power up (after start and taxi)both inverters should be checked for thefollowing:

●

Using the AC volt/frequency meter

❥

115 VAC

❥

400 Hz

❥

Inverter annunciator light out

●

When cycling inverters, check theAC volt/frequency meter drops to zeroand the annunciator light comes onwhen the switch is in the center orOFF position

EXTERNAL POWER

The external power receptacle, under the rightwing outboard of the nacelle, connects an exter-nal power unit to the electrical system when theairplane is parked. The power receptacle isdesigned for a standard three prong AN plug.

When external power is connected, a relay in theexternal power sensor will close only if the polar-ity of the voltage being supplied to the externalpower receptacle is correct (Figure 2-28).

Whenever an external power plug is connected tothe receptacle and the BAT switch is ON, the yel-low EXT PWR annunciator will illuminate,whether or not the external power unit is ON. Ifthe EXT PWR annunciator is flashing – and theexternal power unit is connected – then one of

INVERTER

28VDC26VAC115VAC


40

5

7.540

10

5 5

5 5

5 5

5

5

1011526

115 vac 400 Hz

26 vac 400 Hz

11526

1 2



40

7.5 40

ANN. IND.

TESTJACK

SELRELAY

LHNO. 1INV

RHNO. 2INV

SELSW

NO. 1 INVCONTROL

NO. 2 INVCONTROL



2-19

three conditions exists: EXT PWR Switch isOFF, EXT PWR voltage is low, or EXT PWRvoltage is too high.

External power voltage can be monitored anytime, even before the EXT PWR switch on thepilot’s left subpanel is switched ON, by turningthe VOLTMETER BUS SELECT switch in theoverhead panel (Figure 2-3) to the EXT PWRposition and reading the voltage on the voltmeter.

A high-voltage sensor will lock out the externalpower re lay i f ex te rna l power i s above31 ±0.5 volts DC.

When the EXT PWR - ON - OFF - RESETswitch is switched ON, the external power relaycloses. As external power enters the aircraft. theleft and right generator bus tie relays close, per-mitting power to reach all buses. Consequently,the entire electrical system can be operated.

Observe the following precautions when using anexternal power source:

CAUTION

THE RECOMMENDED MINIMUMINDICATED BATTERY VOLTAGEPRIOR TO CONNECTING EXTER-NAL POWER IS 23 VOLTS.HOWEVER, NEVER CONNECT ANEXTERNAL POWER SOURCE TOTHE AIRPLANE UNLESS A BAT-TERY INDICATING A CHARGE OFAT LEAST 20 VOLTS IS IN THEAIRCRAFT. If the battery voltage isless than 20 volts, the battery must berecharged, or replaced with a batteryindicating at least 20 volts, before con-necting external power.

Only use an external power source fit-t ed wi th an AN- type p lug . Theauxiliary power unit must be regulatedat 28.25 volts DC and be capable ofsupplying at least 1000 amperes for atleast 1 second (300 amperes maximumcontinuous) at a minimum of 16 voltsDC during the start cycle.

Voltage is required to energize the avi-onics master power relays to removethe power from the avionics equip-ment. Therefore, never apply externalpower to the airplane without firstapplying battery voltage.

The bat tery may be damaged i fexposed to voltages higher than30 volts for extended periods of time.

To preclude damage to the externalpower unit, disconnect external powerfrom the airplane before applying gen-erator power to the electrical buses.

Refer to the “Normal Procedures” section ofthe

POH

for procedural detai ls of usingexternal power.

AVIONICS MASTER POWER

The avionics systems installed on each air-plane usually consist of individual nav/comunits, each having its own ON - OFF switch.Avionics packages will vary on different air-plane installations. Due to the large number ofindividual receivers and transmitters, a Beechavionics master switch placarded AVIONICSMASTER POWER is installed on the pilot’sleft subpanel. An Avionics Master PowerSchematic diagram is shown in Figure 2-13.Refer to the Avionics chapter of this trainingmanual for details of the avionics system.

CIRCUIT BREAKERS

Both AC and DC power are distributed to thevarious aircraft systems via two separate circuitbreaker panels which protect most of the com-ponents in the airplane. The smaller one islocated below the fuel management panel, to theleft of the pilot (Figure 2-6). The large panel islocated to the right of the copilot’s position.Each of the circuit breakers has its amperagerating printed on it.

2-20


The small circuit breaker panel, on the lower por-tion of the fuel panel, contains the circuitbreakers for the fuel system. (On LJ-1361, LJ-1363 and after, engine instrument circuit break-ers are also included here.) (See Figure 2-6A)

The large circuit breaker panel is located on thecopilot’s side of the cockpit. This panel containsthe breakers for the remaining electrical systems,which include engine-related systems, all avion-ics components, the environmental system,lights, annunciator warning systems, and other

systems. The circuit breakers for the electricaldistribution system are also located on this panel.

Procedures for tripped circuit breakers, andother related electrical system warnings, can befound in the “Emergency” section of the

Pilot’sOperating Handbook. If a non-essential circuitbreaker on either of the two circuit breaker pan-els trips while in flight, do not reset it. Resettinga tripped breaker can cause further damage tothe component or system.

Figure 2-13 Avionics Master Power Schematic


If an essential system circuit breaker trips, how-ever, wait 30 seconds and then reset it. If it failsto reset, DO NOT attempt to reset it again. Takecorrective action according to the procedures inthe “Emergency” section of your POH.

If all the avionics equipment drops off-line butdoes not trip the circuit breaker, the trouble maybe in the AVIONICS MASTER switch. Theswitch can be bypassed, and your radiosreturned to service, by pulling the AVIONICSMASTER circuit breaker on the copilot’s cir-cuit breaker panel.

The various power distribution configurations forthe electrical system are as follow:

● Power Distribution—Battery OFF(Figure 2-15)

● Power Distribution—Battery ON(Figure 2-16)

● Power Distribution—Battery ON(Generator Ties Manually Closed)(Figure 2-17)

● Power Distribution—Right Engine Start(Generator Ties Normal)(Figure 2-18)

● Power Distribution—Right GeneratorON (Figure 2-19)

● Power Distribution—Left Engine Cross-start (Right Engine Running)(Figure 2-20)

● Power Distribution—Both GeneratorsON (Figure 2-21)

● Power Distribution—Both GeneratorsON (Generator Ties Open) (Figure 2-22)

● Bus Sense Test—Both Generators ON(Figure 2-23)

● Both Generators Failed—Load Shedding(Figure 2-24)

● Right Generator Bus Short—Bus Isola-tion (Figure 2-25)

● Center Bus Short—Bus Isolation(Figure 2-26)

● Triple-Fed Bus Short—Bus Isolation(Figure 2-27)

● Power Distribution—External Power(External Power and Battery SwitchesON) (Figure 2-28)


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D

TRIPLE-FED / BATTERY BUS

CENTER BUSLEFT GEN BUS RIGHT GEN BUSH

ED

275 275275250 250

27

5

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE

L GEN TIE OPEN R GEN TIE OPEN R DC GENBAT TIE OPEN

Figure 2-14 Power Distribution Schematic


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

27

5

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-15 Power Distribution—Battery OFF


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-16 Power Distribution—Battery ON


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-17 Power Distribution—Battery ON (Generator Ties Manually Closed)

2-26 FOR TRAINING PURPOSES ONLY Revision .01

GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-18 Power Distribution—Right Engine Start (Generator Ties Normal)


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-19 Power Distribution—Right Generator ON


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-20 Power Distribution—Left Engine Cross-Start (Right Engine Running)


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-21 Power Distribution—Both Generators ON


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

27

5

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-22 Power Distribution—Both Generators ON (Generator Ties Open)


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

27

5

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-23 Bus Sense Test—Both Generators ON


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-24 Both Generators Failed—Load Shedding


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

27

5

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-25 Right Generator Bus Short—Bus Isolation

2-34 FOR TRAINING PURPOSES ONLY Revision .01

GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

27

5

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-26 Center Bus Short—Bus Isolation


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

BATTERYRELAY

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-27 Triple-Fed Bus Short—Bus Isolation


GEN TIES MAN CLOSE

NORM

OPEN

BUS SENSERESET

TEST

LEFT GEN BUS

HOT BATTERY BUS

275

HE

D



ED

275 275275250 250

275

H E D

BATTERY

BATT. BUSTIE

LEFT GENBUS TIE

RIGHT GENBUS TIE

RCS LCS

RSRLSR

GENSW.

GPU 60

60

60

L GENCONTACTOR

GCU

GEN

SW.

L GEN

CONTACTOR

GCU

GENSW.

R GENCONTACTOR

GCU

BATTERY CHARGE

L DC GEN

MAN TIES CLOSE


Figure 2-28 Power Distribution—External Power (External Power and Battery Switches ON)


3-i

CHAPTER 3

CONTENTS

Page

LIGHTING

INTRODUCTION ..................................................................................................................

3-1

DESCRIPTION.......................................................................................................................

3-1

Cockpit Lighting..............................................................................................................

3-1

Cabin Lighting.................................................................................................................

3-2

Exterior Lighting .............................................................................................................

3-4

Circuit Breakers...............................................................................................................

3-4


3-iii

Figure Title Page

ILLUSTRATIONS

3-1

Overhead Lighting Control Panel ............................................................................

3-2

3-2

Cabin Lighting Controls ..........................................................................................

3-3

3-3

Threshold Light Switch ...........................................................................................

3-3

3-4

Exterior Light Controls............................................................................................

3-4

3-5

Light System Circuit Breakers.................................................................................

3-5


3-1

CHAPTER 3LIGHTING

INTRODUCTION

The aircraft lighting system consists of cockpit-controlled interior and exterior lights. Interiorlights are in the cockpit and passenger cabin and consists of navigation lights, entry and exitthreshold lights, and baggage area lights. Exterior lighting consists of navigation lights, rotatingbeacons, strobe lights, landing and taxi lights, ice lights, and recognition lights.

DESCRIPTION

The Lighting chapter of the training manual pre-sents a description and discussion of the airplanelighting system and components. The locatio-nand purpose of switches, indicators, lights, andcircuit breakers are described.

COCKPIT LIGHTING

An overhead light control panel, easily accessibleto both pilot and copilot, incorporates a func-tional arrangement of all lighting systems in thecockpit (Figure 3-1). Each light group has itsown rheostat switch placarded BRT - OFF.

��

��

� ��

EXIT

3-2


The MASTER PANEL LIGHTS - ON/OFFswitch is the master switch for: PILOT & COPI-LOT FLIGHT INSTR, PILOT & COPILOTGYRO INSTR, ENGINE INSTR, AVIONICSPANEL, OVHD, PED & SUBPANEL, and SIDEPANEL. The indirect instrument lighting andmap (overhead) lights are controlled by rheostatswitches mounted on the overhead panel.

CABIN LIGHTING

A three-position switch on the copilot’s left sub-panel light control panel, placarded CABIN -START/BRIGHT - DIM - OFF on the C90A andCABIN - BRIGHT - DIM - OFF on the C90B,

controls the indirect fluorescent cabin lights (Fig-ure 3-2). A switch to the right of the interior lightswitch activates the cabin NO SMOKING/FAS-TEN SEAT BELT signs and accompanyingchimes. This three-position switch is placardedNO SMK & FSB - OFF - FSB.

A hot-wired threshold light is mounted on the leftside of the entryway at floor level. Optionalairstair door lights mounted under each step maybe installed. These lights share the same controls;a slide type switch (Figure 3-3) mounted adjacentto the threshold light, and a microswitchmounted in the door lock. Whenever the slideswitch is in the ON position and the door is open,the lights will come on.

Figure 3-1 Overhead Lighting Control Panel


3-3

To turn the lights OFF, either use the thresholdlight switch, or fully close and lock the cabindoor. The microswitch in the door lock will turnoff the lights when the threshold switch is left on.

The lights will not go out if the door is simplylatched, the door handle must be in the fullylocked position.

When the battery master switch is on, the indi-vidual reading lights along the top of the cabinmay be turned on or off by the passengers withthe pushbutton switch adjacent to each light.

The light in the baggage compartment may beturned on or off by the adjacent push-buttonswitch regardless of the position of the batterymaster switch. This baggage compartment lightis connected to the hot battery buss.

Figure 3-2 Cabin Lighting Controls

Figure 3-3 Threshold Light Switch

3-4


EXTERIOR LIGHTING

Switches for the landing lights, taxi lights,wing ice lights, navigation lights, recognitionlights, rotating beacons, and wingtip and tailflood lights are located on the pilot’s subpanelFigure 3-4. They are appropriately placarded asto their function.

Tail floodlights, if installed, are incorporated intothe horizontal stabilizers and are designed to

illuminate both sides of the vertical stabilizer. Aswitch for these lights, placarded LIGHTS -TAIL FLOOD - OFF, is located on the pilot’ssubpanel (Figure 3-2).

CIRCUIT BREAKERS

Lighting system circuit breakers are shown inFigure 3-5.

Figure 3-4 Exterior Light Controls


3-5

Figure 3-5 Light System Circuit Breakers


4-i

CHAPTER 4

CONTENTS

Page

MASTER WARNING SYSTEM

INTRODUCTION ...................................................................................................................

4-1

GENERAL ...............................................................................................................................

4-1

ANNUNCIATOR SYSTEM....................................................................................................

4-3

Fault Warning Flasher ......................................................................................................

4-3

Dimming...........................................................................................................................

4-5

Testing and Lamp Replacement .......................................................................................

4-5

ANNUNCIATOR PANEL DESCRIPTION ...........................................................................

4-6

4-ii



4-iii

Figure Title Page

ILLUSTRATIONS

4-1

Annunciator System..................................................................................................

4-2

4-2

Master Caution and Fault Warning Flashers ............................................................

4-3

4-3

Warning, Caution, and Advisory Annunciators .......................................................

4-4

4-4

Lamp Replacement ...................................................................................................

4-5

Revision .02


4-v

Table Title Page

TABLES

4-1

WARNING Annunciators .........................................................................................

4-6

4-2

CAUTION Annunciators ..........................................................................................

4-7

4-3

ADVISORY Annunciators........................................................................................

4-9


4-1

CHAPTER 4MASTER WARNING SYSTEM

INTRODUCTION

Warning and caution indicators can be the first indication of trouble or malfunction in somesystem or component of the airplane. Crewmembers should have complete familiarity withthese indicators and the related action necessary to correct the problem or cope with the situa-tion until a safe landing can be made. In the case of an on-ground indication, the problemshould be corrected before flight.

GENERAL

This chapter presents a description and discus-sion of the warning, caution, and advisoryannunciator panel.

The annunciator panel is described in detail,including each annunciator, its purpose, and theassociated cause for illumination.

TEST

4-2


Figure 4-1 Annunciator System


4-3

ANNUNCIATOR SYSTEM

The annunciator system (Figure 4-1) consists ofan annunciator panel centrally located in theglareshield, a PRESS-TO-TEST switch, and aFAULT WARNING flasher (Figure 4-2). The redFAULT WARNING flasher (and yellow MASTERCAUTION flasher [LJ-1353 and after]) is locatedin the glareshield in front of the pilot, and thePRESS-TO-TEST switch is located immediatelyto the right of the annunciator panel. The annunci-ators are of the word-readout type. Whenever afault condition covered by the annunciator systemoccurs, a signal is generated, and the appropriateannunciator is illuminated

.

Whenever an annunciator-covered conditionoccurs that requires the pilot’s attention but nothis immediate reaction, the appropriate yellowcaution annunciator (Figure 4-3) in the annuncia-tor panel illuminates (as well as the MASTERCAUTION flasher on LJ-1353 and after).

The annunciator panel also contains green advi-sory annunciators. There are no fault warningflashers associated with advisory annunciators.

An illuminated caution annunciator on theannunciator panel will remain on until the faultcondition is corrected, at which time it will extin-guish. An annunciator can be extinguished onlyby correcting the condition indicated on the illu-minated lens.

The illumination of a green annunciator light willnot trigger the fault warning system, but a redannunciator will actuate the FAULT WARNINGflasher. Yellow annunciators will actuate the yel-low MASTER CAUTION flasher..

FAULT WARNING FLASHER

If the fault requires the immediate attention andreaction of the pilot, the appropriate red warningannunciator (Figure 4-3) in the annunciator panelilluminates, and the FAULT WARNING flasherbegins flashing.

Figure 4-2 Master Caution and Fault Warning Flashers

4-4


Any illuminated red lens in the annunciator panelwill remain on until the fault is corrected. TheFAULT WARNING flasher can be extinguishedby depressing the face of the FAULT WARNINGflasher, even if the fault is not corrected. In such acase, the FAULT WARNING flasher will again

be activated if an additional warning annunciatorilluminates. When a warning fault is corrected,the affected warning annunciator will extinguish,but the FAULT WARNING flasher will continueflashing until it is depressed.

NOTE: CHIP DETECT - Lights red on the C90A DC GEN - Lights Red (Prior to LJ-1353 and after) NO FUEL XFR - Lights Red (Prior to LJ-1353) OIL PRESS -Optional prior to LJ-1353

L FUEL PRESS

L DC GEN

L IGNITION ON

L OIL PRESS

L NO FUEL XFR

R IGNITION ON

RVS NOT READY

L AUTOFEATHER

L CHIP DETECT

R AUTOFEATHER

L ENG ICE FAIL

L ENG ANTI-ICE

R ENG ICE FAIL

INVERTER A/P FAIL

MAN TIES CLOSE

A/P TRIM FAIL

L GEN TIE OPEN

FUEL CROSSFEED

CABIN ALT HI

BAT TIE OPEN

HYD FLUID LO

CABIN DOOR

R GEN TIE OPEN

BATTERY CHARGE

R ENG ICE FAIL

R ENG ANTI-ICE

BAGGAGE DOOR

PITCH TRIM OFF

EXT POWER

R CHIP DETECT

LDG/TAXI LIGHT

R OIL PRESS

R NO FUEL XFR

R FUEL PRESS

R DC GEN

Figure 4-3 Warning, Caution, and Advisory Annunciators


4-5

DIMMING

The warning, caution, and advisory annunciatorsfeature both a bright and a dim mode of illumina-tion intensity. The dim mode will be selectedautomatically whenever all of the following con-ditions are met:

●

A generator is on line.

●

The

OVERHEAD FLOODLIGHT is OFF.

●

The MASTER PANEL LIGHTS switchis ON.

●

The PILOT FLIGHT LIGHTS are ON.

●

The ambient light level in the cockpit (assensed by a photoelectric cell located inthe overhead light control panel) is belowa preset value.

Unless all these conditions are met, the brightmode will be selected automatically. The FAULTWARNING flasher does not have a dim mode. Awarning or caution annunciator will cause thedim mode to be bright.

TESTING AND LAMP REPLACEMENT

The lamps in the annunciator system should betested before every flight and any time the integ-rity of a lamp is in question. Depressing thePRESS-TO-TEST button, located to the right ofthe annunciator panel in the glareshield, illumi-nates all the annunciator lights and the FAULTWARNING flasher. Any lamp that fails to illumi-nate when tested should be replaced.

The annunciator panel style allows each annunci-ator to be removed from the panel (Figure 4-4).Each readout annunciator contains two lamps. Toreplace any annunciator lamp, first depress thecenter of the annunciator with your finger.Release your finger, and the annunciator will popout slightly. Pull the annunciator from the panel,and remove the lamp from the rear of the annun-ciator. Replace the failed lamp with a spare lampcontained in an unused annunciator. Depress theannunciator until it locks in place.

Figure 4-4 Lamp Replacement

4-6


ANNUNCIATOR PANEL DESCRIPTION

Tables 4-1, 4-2, and 4-3 list all the warning, cau-tion, and advisory annunciators on the King Air

C90A/B. The cause for illumination is includedbeside each annunciator.

Table 4-1 WARNING ANNUNCIATORS

NOMENCLATURE CAUSE FOR ILLUMINATION

Low fuel pressure on left side; check boost pump, crossfeed.

(LJ-1353 and after)

The inverter selected is inoperative, or both inverters are off.

Autopilot is disconnected by switching other than pilot’s discon-nect button.

Improper trim or no trim command from autopilot.

Cabin altitude exceeds10,000 feet (12,500 feet on LJ-1353 and later) pressure altitude.

Cabin door is open or not secure.

Nose baggage door is not secure (Prior to LJ-1531).

(LJ-1353 and after)

Low fuel pressure on right side; check boost pump, crossfeed.

Fire in left engine compartment.

Fire in right engine compartment.

* Optional equipment

L FUEL PRESS

L OIL PRESS

INVERTER

A/P FAIL

A/P TRIM FAIL

CABIN ALT HI

CABIN DOOR

BAGGAGE DOOR*

R OIL PRESS

R FUEL PRESS

L ENG FIRE*

R ENG FIRE*


4-7

Table 4-2 CAUTION ANNUNCIATORS


Left generator is off line (Red prior to LJ-1353 and after).

Left wing tank is empty or transfer pump failed.

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

Metal contamination is detected in left engine oil, probable engine shutdown (red for C90A; yellow for C90B).

Left engine anti-ice vanes in transit or inoperative.

Right engine anti-ice vanes in transit or inoperative.

Left generator bus is isolated from the center bus.

Battery is isolated from the generator buses and center bus.

Right generator bus is isolated from the center bus.

Crossfeed valve is receiving power.

Hydraulic fluid in the hydraulic fluid reservoir is low.

Charge rate on the battery exceeds 7 amps for 6 seconds(Airplanes prior to LJ-1534).

Pitch trim deenergized by a trim disconnect switch on the control wheel with the system power switch on the pedestal turned on.

Metal contamination is detected in right engine oil, probable engine shutdown (red for C90A; yellow for C90B).

L DC GEN

L NO FUEL XFR

RVS NOT READY

L CHIP DETECT

L ENG ICE FAIL

R ENG ICE FAIL

L GEN TIE OPEN

BAT TIE OPEN

R GEN TIE OPEN

FUEL CROSSFEED

HYD FLUID LO

BATTERY CHARGE

PITCH TRIM OFF

R CHIP DETECT

4-8



Right wing tank is empty or transfer pump failed.

Right generator is off line (Yellow for LJ-1353 and after).

External power connector is plugged in.

R NO FUEL XFR

R DC GEN

EXT PWR

Table 4-2 CAUTION ANNUNCIATORS (Cont)


4-9

Table 4-3 ADVISORY ANNUNCIATORS


System is armed and left engine torque is below 400 ft-lb, or the left ignition and engine start switch is ON.

System is armed and right engine torque is below 400 ft-lb, or the right ignition and engine start switch is ON.

Left autofeather is armed with power levers advanced above 90% N

1

position, or autofeather test switch is in test.

Right autofeather is armed with power levers advanced above 90% N

1

position, or autofeather test switch is in test.

Left engine anti-ice vanes are in position for icing conditions.

Right engine anti-ice vanes are in position for icing conditions.

Manually closed generator bus ties.

Landing lights or taxi light is on with landing gear UP.

L IGNITION ON

R IGNITION ON

L AUTOFEATHER*

R AUTOFEATHER*

L ENG ANTI-ICE

R ENG ANTI-ICE

MAN TIES CLOSE

LDG/TAXI LIGHT

*Optional Equipment

4-10



5-i

CHAPTER 5

CONTENTS

Page

FUEL SYSTEM

INTRODUCTION ..................................................................................................................

5-1

DESCRIPTION.......................................................................................................................

5-1

Fuel System .....................................................................................................................

5-1

Fuel Tank System............................................................................................................

5-2

Boost Pumps....................................................................................................................

5-2

Fuel Transfer Pumps .......................................................................................................

5-5

Fuel Capacity...................................................................................................................

5-6

Fuel Tank Vents ..............................................................................................................

5-7

FUEL SYSTEM OPERATION ..............................................................................................

5-8

Firewall Shutoff Valves.................................................................................................

5-10

Crossfeed Operation ......................................................................................................

5-10

Fuel Drain Purge System...............................................................................................

5-12

FUEL GAGING SYSTEM ...................................................................................................

5-12

Components and Operation ...........................................................................................

5-13

FUEL DRAINS.....................................................................................................................

5-14

FUEL HANDLING PRACTICES ........................................................................................

5-14

Fuel Grades and Additives ............................................................................................

5-17

Filling the Tanks............................................................................................................

5-18

Draining the Fuel System ..............................................................................................

5-19


5-iii

Figure Title Page

ILLUSTRATIONS

5-1

Fuel System Schematic Diagram.............................................................................

5-3

5-2

Fuel Tank System ....................................................................................................

5-4

5-3

Fuel Transfer Pump Switch .....................................................................................

5-6

5-4

Fuel Control Panel ...................................................................................................

5-6

5-5

Fuel Vent System.....................................................................................................

5-7

5-6

Fuel Flow Diagram ..................................................................................................

5-8

5-7

Firewall Shutoff Valve ..........................................................................................

5-10

5-8

Firewall Shutoff Valve Switches ...........................................................................

5-10

5-9

Crossfeed Schematic..............................................................................................

5-11

5-10

Fuel Drain Purge System Schematic .....................................................................

5-12

5-11

Fuel Quantity Indication System ...........................................................................

5-12

5-12

Fuel Probe..............................................................................................................

5-13

5-13

Fuel Drains.............................................................................................................

5-14

5-14

Fuel Drain Locations .............................................................................................

5-15

5-15

Fuel Temperature Graph........................................................................................

5-16


5-1

CHAPTER 5FUEL SYSTEM

INTRODUCTION

A complete understanding of the fuel system is essential to competent and confident operationof the aircraft. Management of fuel and fuel system components is a major everyday concern ofthe pilot. This section gives the pilot the information he needs for safe, efficient fuelmanagement.

DESCRIPTION

The Fuel System section of the training manualpresents a description and discussion of the fuelsystem. The physical layout of the fuel cells andfuel system are described in this section. Correctuse of the boost pumps, transfer pumps, cross-feed, and firewall shutoff valves are discussed.Fuel drains, their location, and type are describedwith correct procedure for taking and inspecting

samples of fuel. Approved fuels and tank fillingsequence are included.

FUEL SYSTEM

The Beechcraft King Air fuel system is designedto simplify flight procedures in the cockpit, and

��

��

��

0

2

4 6

8

10

MAINFUEL

LBS X 100

5-2


provide easy access on the ground (Figure 5-1).There are two separate wing fuel systems, onefor each engine, connected by a valve-controlledcrossfeed system. Each fuel system consists of anacelle tank and four interconnected wing tanks,electrical boost and transfer pumps and an elec-trically operated crossfeed valve. Total usablefuel capacity is 384 gallons.

Three modes of operation are available, each ofwhich is described briefly.

1. Normal operation—Each engine receivesfuel from its corresponding fuel cells andboost pump. The boost pump is required toprovide fuel under pressure to the enginedriven high pressure pump.

2. Automatic crossfeed operation—In the eventof a boost pump failure, boost pressure isobtained by supplying fuel to both engines,through the crossfeed valve, from one boostpump. A drop in output pressure from thefailed pump is sensed by a pressure switch,which automatically opens the crossfeedvalve when the pressure drops below about10 psi, and illuminates the low fuel pressureannunciator. The fuel pressure annunciatorwill then extinguish as pressure is restored bythe boost pump on the opposite engine.

3. Suction feed—This mode of operation maybe employed after a boost pump has failed,and allows the use of fuel from tanks on theside with the failed pump. Suction feed oper-ation is obtained by moving the crossfeedvalve control switch from the AUTO positionto the CLOSED position. Vacuum created bythe engine-driven fuel pump draws fuel fromthe nacelle fuel tank. Suction feed is limitedto ten hours cumulative between engine-driven fuel pump overhauls.

FUEL TANK SYSTEM

The fuel system (Figure 5-2) in each wing con-sists of one wing leading-edge bladder-type tank(40 gallons), two outboard-wing panel bladder-type tanks (23 gallons and 25 gallons), one centersection bladder-type tank (44 gallons), and the

nacelle tank (61 gallons). The total usable fuelcapacity of each wing fuel system is 192 gallons.The outboard wing tanks supply the center sec-tion and nacelle tanks by gravity flow. Since thecenter section tank is lower than the other wingtanks and the nacelle tank, the fuel is transferredto the nacelle tank by the fuel transfer pump inthe low point of the center section tank. Fuel foreach engine is pumped directly from its nacellefuel tank by an electric boost pump. Each systemhas two filler cap openings; one in the top of thenacelle tank and one mid-wing in the leadingedge tank. An anti-siphon valve is installed ateach filler port to prevent the loss of fuel or col-lapse of fuel-tank bladder in the event the fillercap is improperly secured.

There is a check valve between the nacelle tankand the wing tank. Fuel can flow only into thenacelle tank, not back into the wing tank. If a fullfuel load is needed, fill the nacelle tank first, thenfill the wing tank.

The heated fuel vent and the NASA integral ramscoop vent work together to prevent the bladdersfrom collapsing as fuel is drawn out of them.

Each nacelle tank is connected to the engine onthe opposite side by a crossfeed line for single-engine or failed boost pump operation. Crossfeedoperation is automatic depending on the boostpump selected in the feeding nacelle tank. Thissystem makes it possible for fuel in either wingsystem to be available to either engine, or bothengines simultaneously.

BOOST PUMPS

Each system has a submerged boost pump in thenacelle tank. This pump supplies a pressure ofabout 30 psi to the engine-driven fuel pump. Theboost pumps are submerged, rotary, vane-typeimpeller pumps, and are electrically-driven. A10-amp circuit breaker for each boost pump islocated on the fuel panel. Two red FUEL PRESSannunciators are associated with the boostpumps. When illuminated, there is low fuel pres-sure on the side indicated. Check the boostpumps prior to flight.


5-3

Figure 5-1 Fuel System Schematic Diagram

5-4


With crossfeed in AUTO, a boost pump failurewill be denoted by the momentary illuminationof the FUEL PRESS annunciator and fault warn-ing flasher, then the illumination of the FUELCROSSFEED annunciator. To identify the failedboost pump, momentarily place the crossfeed inthe CLOSED position. The FUEL PRESS annun-ciator on the side of the failed boost pump willilluminate. Place the crossfeed switch in theOPEN position. The FUEL PRESS annunciatorwill then extinguish.

In the event of a boost pump failure during anyphase of flight, the system will begin to crossfeedautomatically. If the boost pump fails , the cross-feed switch may be closed and the flightcontinued, relying on the engine-driven highpressure pump. In some instances the pilot may

elect to continue the flight with the remainingpump and the crossfeed system in operation.

CAUTION

Operation with the FUEL PRESSannunciator on is limited to 10 hours,after which the engine-driven highpressure pump must be overhauled orreplaced. When operating with Avia-tion Gasoline base fuels, operation onthe engine-driven high pressure pumpalone is permitted up to 8,000 feet fora period not to exceed 10 hours. Oper-ation above 8,000 feet requires boostor operation of crossfeed.

The following Fuel Management Limitations,listed in the Limitations section of the C90A andC90B

POH

, pertain to fuel system boost pumps.

Figure 5-2 Fuel Tank System

Revision .01


5-5

Both boost pumps must be operable prior to takeoff.

Operation is limited to 8000 feet when operatingon aviation gasoline with boost pumps inoperative.

Operation with the FUEL PRESS annunciator onis limited to 10 hours between main engine-driven fuel pump overhaul or replacement.

FUEL TRANSFER PUMPS

Fuel level in the nacelle tank is automaticallymaintained at near full capacity during normaloperation by a fuel transfer system, whenever thefuel level in the nacelle tank drops by approxi-mately 10 gallons. Submerged, electrically-driven,impeller pumps located in the wing center sectiontanks provide the motive force for fuel transferfrom wing tanks to nacelle tanks. The transferpumps are controlled by float-operated switcheson the nacelle tank fuel quantity transmitters.

Fuel is transferred automatically when theTRANSFER PUMP switches are placed inAUTO, unless the nacelle tanks are full. As theengines burn fuel from the nacelle tanks (61 gal-lon capacity each tank), fuel from the wing tanksis transferred into the nacelle tanks each time thenacelle tank levels drop approximately 10 gal-lons. The nacelle tanks will fill until the fuelreaches the upper transfer limit and a float switchturns the TRANSFER PUMP off.

A pressure switch, located in the fuel transferline, will automatically turn off the transfer pumpif a preset pressure is not obtained within approx-imately 30 seconds after the pump is turned on,or if the transfer pump pressure drops below apreset pressure due to empty wing tanks or pumpfailure. For example, when 132 gallons of fuel(each side) are used from the wing tanks (132gallons usable each side), the pressure sensingswitch reacts to a pressure drop in the fuel trans-fer line as the wing tanks are exhausted of fuel.After-30 seconds, the transfer pump shuts off andthe respective yellow (red on prior to LJ1353)NO FUEL XFR annunciator on the anuunciatorpanel illuminates.

The NO FUEL XFR annunciators will illumi-nate for the reasons mentioned: no pressureafter 30 second time delay due to empty wingtanks or transfer pump failure. The NO FUELXFR annunciator also functions as an operationindicator for the transfer pump during preflight.A TRANSFER TEST swi tch (p lacardedENGINE L and ENGINE R) is provided to ver-ify the operation of each pump when its nacelletank is full. Holding the Transfer Test switch inthe test position (either L or R) will activate thetransfer pump and pressure sensor. In the testmode, the 30-second delay is by-passed, result-ing in immediate indications. The NO FUELXFR annunciator will momentarily illuminateand the Fault Warning Flasher will also beginflashing. The NO FUEL XFR annunciator willextinguish when fuel pressure to the sensorreaches a minimum pressure of 2.5 psi. If thetransfer pump is operating, use of the transfertest will not be possible.

The fuel transfer system may be monitored byperiodically checking the nacelle tank quantityagainst the total tank quantity.

If the NO FUEL XFR does not illuminate and thetransfer test indicates a working pump, the flowswitches may be suspect. Using the transfer testwill begin the fill-up cycle, however, fuel quan-tity in the nacelle will drop below the lower levelwithout activating the transfer pump. Proceed bymoving the transfer pump switch (Figure 5-3) tothe OVERRIDE position. In this mode, the trans-fer pump will run continuously until the transferpump switch is returned to the OFF position.When the nacelle tank becomes full, excess fuelwill be returned to the center section wing tankthrough the vent line.

Illumination of the NO FUEL XFR annunciatormay indicate a normal or abnormal situation.During normal operation, when the fuel in thewing tanks is exhausted, the NO FUEL XFRannunciator indicates that the wing tanks areempty and the fuel transfer switch should beturned off.

5-6


If the transfer pump fails to operate during flight,gravity feed will perform the transfer. When thenacelle tank level drops to approximately150 pounds, or approximately 22 gallons, thegravity port in the nacelle tank opens and gravityflow from the wing tank starts. All wing fuel,except 28 gallons from the center section tank,will transfer during gravity feed.

FUEL CAPACITY

The fuel quantity system is a capacitance gag-ing system with one quantity indicator per wing

(Figure 5-4). A toggle switch selector allows thepilot to check total system or just the nacelletank quantity. The system has a total capacity of387 gallons, and a maximum usable fuel quan-tity of 384 gallons. The fuel quantity gages andthe engine fuel flow indicators read in poundstimes 100. At 6.7 pounds per gallon, 2572.8pounds of usable fuel are available in the sys-tem, 1286.4 pounds per side.

There is no structural limitation for which a Max-imum Zero Fuel Weight must be set.

OPEN

CLOSED

FIREWALLSHUTOFF

VALVE

LEFT

RIGHT

FUEL SYSTEMS

TRANS. PUMPOVERRIDE

TRANS. PUMPOVERRIDE

OFF OFF

SEE MANUAL FORFUEL CAPACITY

BOOST PUMPON BOOST PUMP

ON

OFF OFF

TRANSFER TEST

FUEL QUANITYTOTAL

NACELLE

ENGINE ENGINE

OFF

AUTO AUTO

CROSSFEEDOPEN

CLOSE

AUTO

LEFT

64

122

0 14

108

QTY

FUEL

LBS X 100

RIGHT

64

122

0 14

108

QTYLBS X 100

FUEL

CLOSED

ITT PROPTACH

TURBINETACH

FUELFLOW

OILPRESS

OILTEMP

5 5 5 5 5 5 5

5 5 5 5 5 5 5

TORQUE

FIREWALLVALVE

TRANSPUMP

QTYIND

PRESSWARN

5 10 5 5 5 5

BOOSTPUMP

CROSSFEED

QTYIND

TRANSPUMP

BOOSTPUMP

FIREWALLVALVE

5 5 5 10 5

PRESSWARN

RIGHTLEFT

OPEN

ENGINE INSTRUMENT

FIREWALLSHUTOFF

VALVE

MAIN TANKONLY

MAIN TANKONLY

TRANS. PUMPOVERRIDE

OFF

AUTO

TRANS. PUMPOVERRIDE

OFF

AUTO

Figure 5-3 Fuel Transfer Pump Switch

OPEN

CLOSED

FIREWALLSHUTOFF

VALVE

LEFT

RIGHT

FUEL SYSTEMS

TRANS. PUMPOVERRIDE

TRANS. PUMPOVERRIDE

OFF OFF



ON

OFF OFF

TRANSFER TEST

FUEL QUANITYTOTAL

NACELLE

ENGINE ENGINE

OFF

AUTO AUTO

CROSSFEEDOPEN

CLOSE

AUTO

LEFT

64

122

0 14

108

QTY

FUEL

LBS X 100

RIGHT

64

122

0 14

108

QTYLBS X 100

FUEL

CLOSED

ITT PROPTACH

TURBINETACH

FUELFLOW

OILPRESS

OILTEMP

5 5 5 5 5 5 5

5 5 5 5 5 5 5

TORQUE

FIREWALLVALVE

TRANSPUMP

QTYIND

PRESSWARN

5 10 5 5 5 5

BOOSTPUMP

CROSSFEED

QTYIND

TRANSPUMP

BOOSTPUMP

FIREWALLVALVE

5 5 5 10 5

PRESSWARN

RIGHTLEFT

OPEN

ENGINE INSTRUMENT

FIREWALLSHUTOFF

VALVE

MAIN TANKONLY

MAIN TANKONLY

LEFT

64

122

0 14

108

QTY

FUEL

LBS X 100

MAIN TANKONLY

Figure 5-4 Fuel Control Panel


5-7

FUEL TANK VENTS

The fuel system is vented through a recessed ramscoop vent, coupled to a heated external vent,located on the underside of the wing, adjacent tothe nacelle (Figure 5-5). One vent is recessed toprevent icing. The external vent is heated to pre-vent icing. Each vent serves as a backup for theother should one or the other become plugged.

In each wing fuel system, the wing panel tanks, theleading edge tank, the center section tank, and thenacelle tank are all crossvented with one another.

The line from the vent valve in the outboard wingpanel fuel tank is routed forward along the lead-ing edge of the wing, inboard to the nacelle, andaft through a check valve to the heated ram vent.Another line tees off from the heated vent lineand extends to a recessed or ram scoop vent. Theheated vent is described in the Anti-Ice Sectionof this manual. A suction relief valve is installedin the line from the float-operated vent valve tothe siphon break line.

Figure 5-5 Fuel Vent System

5-8


FUEL SYSTEM OPERATION

Fuel flow from each wing tank system andnacelle tank is automatic without pilot action(Figure 5-6). The wing tanks gravity feed into thecenter section tank through a line extending fromthe aft inboard wing tank to be outboard side ofthe center section tank. A flapper-type checkvalve in the end of the gravity feed line preventsany backflow of fuel into the wing tanks.

The fuel pressure required to operate the engineis provided by an engine-driven fuel pumpmounted in conjunction with the fuel control uniton the accessory case. Fuel is pumped to the highpressure fuel pump by an electrically-drivenboost pump submerged in the nacelle tank.

The supply line from the nacelle tank is routedfrom the outboard side of the nacelle tank, for-

ward to the engine-driven fuel pump through amotored firewall shutoff valve installed in thefuel line immediately behind the engine firewall.

The firewall shutoff valve for each engine fuelsystem is actuated by its respective FW SHUT-OFF VALVE switch on the pilot’s fuel controlpanel . When the FW SHUTOFF VALVEswitch is closed, its respective firewall shutoffvalve closes to shut off the flow of fuel to theengine. From the firewall shutoff valve, fuel isrouted to the fuel strainer filter and drain onthe lower center of the engine firewall, the fuelpressure switch, the fuel flow indicator trans-mitter, the fuel heater, and then to the engine-driven fuel pump and engine fuel control unit.The 20 micron filter incorporates a bypassvalve to permit fuel flow in case of pluggingand a drain valve used to drain the filter priorto each flight. A pressure switch mounteddirectly above the filter senses boost pump fuel

Figure 5-6 Fuel Flow Diagram


5-9

pressure at the filter. At a pressure, about 10psi, the switch closes and actuates the redFUEL PRESS light in the annunciator panel.

CAUTION

Operation with the FUEL PRESS lightON is limited to 10 hours betweenoverhaul or replacement of the engine-driven fuel pump. Such operation isrestricted to 10 hours at altitudes not toexceed 8000 feet when aviation gaso-line is being used. Windmilling time isnot equivalent to operation of theengine at high power with respect tothe effects of cavitation on fuel pumpcomponents; consequently, windmill-ing time is not to be included in the10-hour limit on engine operationwithout a boost pump.

The red FUEL PRESS light will go out at about10 psi of increasing fuel pressure. From the fuelstrainer and filter, fuel is routed through the fuelflow transmitter mounted on the firewall, inboardof the pressure switch. Fuel from the transmitteris routed through the fuel heater, which utilizesheat from the engine oil to warm the fuel. Thefuel is then routed to the fuel control unit thatmonitors the flow of fuel to the engine fuel noz-zles. A heater boot is also installed on thegovernor control line of each engine. Each airline heater is protected by a 7.5 ampere, push-pull circuit breaker mounted in the circuitbreaker panel beside the copilot. The heaters arecontrolled by switches installed on the pedestaland activated by the condition levers.

The engine-driven fuel pump is mounted on theaccessory case of the engine in conjunction with

the fuel control unit. This pump is protectedagainst fuel contamination by an internal, 200-mesh strainer. The primary fuel boost pump is anelectrically-driven pump located in the bottom ofeach nacelle tank. The electrically-driven boostpump is capable of supplying fuel to the engine-driven fuel pump at the minimum pressurerequirements of the engine manufacturer.

CAUTION

Should the boost pumps fail, suctionfeed operation may be employed; how-ever, suc t ion feed opera t ion i srestricted to 10 hours total t imebetween fuel pump overhaul periods.If the engine-driven pump is operatedon suction feed beyond the 10-hourlimit, overhaul or replacement of thepump is necessary.

The electrically-driven boost pump also providesthe pressure required for the crossfeed of fuelfrom one side of the aircraft to the other.

The electrical power with which the boost pumpsare operated is controlled by lever-lock toggleswitches on the fuel control panel. One source ofpower to the boost pumps is supplied from thetriple-fed bus that supplies the circuit breakers.This circuit is protected by two 10-ampere circuitbreakers located on the fuel panel. Power fromthis circuit is available only when the masterswitch is on.

The other source of power to the boost pumps isdirectly from the battery through the batteryemergency bus. During shutdown, both boostpump switches and crossfeed must be turned offto prevent discharge of the battery.

5-10


FIREWALL SHUTOFF VALVES

The firewall shutoff valves (Figure 5-7), locatedbetween the engine-driven fuel pump and thenacelle tank, are controlled by guarded switchesin the cockpit (Figure 5-8). There is one switchon each side of the fuel system circuit breakerpanel on the fuel panel. These switches have twopositions. The OPEN position allows uninter-rupted fuel flow to the engine. The CLOSEposition cuts off all fuel to the engine. When thered guard closes, it forces the switch into theopen position and protects it in the open position.

Each firewall shutoff valve receives electricpower through its own 5-amp breaker on the fuelpanel which brings electric power from the tri-ple-fed bus as well as the generator bus. Thissource of power is available only when the bat-tery and/or generator switches are on. The onlypilot action necessary to ensure main fuel systemoperation is to have the firewall shutoff valves inthe OPEN position.

CROSSFEED OPERATION

Crossfeeding fuel is authorized only in the eventof engine failure or electric boost pump failure.

Each nacelle tank is connected to the engine inthe opposite wing by a crossfeed line routed fromthe side of the nacelle, aft to the center section,and across to the side of the opposite nacelle. Thecrossfeed line is controlled by a valve (Figure5-9). With the crossfeed valve OPEN, one systemcan supply fuel to the other engine. The systemuses the electric boost pump in the nacelle tank.This pump supplies the pressure to transfer fuelas well as fuel boost to one or both engines. Withone engine inoperative, the crossfeed systemallows fuel from the inoperative side to be sup-plied to the operating engine.Figure 5-7 Firewall Shutoff Valve

OPEN

CLOSED

FIREWALLSHUTOFF

VALVE

LEFT

RIGHT

FUEL SYSTEMS

TRANS. PUMPOVERRIDE

TRANS. PUMPOVERRIDE

OFF OFF



ON

OFF OFF

TRANSFER TEST

FUEL QUANITYTOTAL

NACELLE

ENGINE ENGINE

OFF

AUTO AUTO

CROSSFEEDOPEN

CLOSE

AUTO

LEFT

64

122

0 14

108

QTY

FUEL

LBS X 100

RIGHT

64

122

0 14

108

QTYLBS X 100

FUEL

CLOSED

ITT PROPTACH

TURBINETACH

FUELFLOW

OILPRESS

OILTEMP

5 5 5 5 5 5 5

5 5 5 5 5 5 5

TORQUE

FIREWALLVALVE

TRANSPUMP

QTYIND

PRESSWARN

5 10 5 5 5 5

BOOSTPUMP

CROSSFEED

QTYIND

TRANSPUMP

BOOSTPUMP

FIREWALLVALVE

5 5 5 10 5

PRESSWARN

RIGHTLEFT

OPEN

ENGINE INSTRUMENT

FIREWALLSHUTOFF

VALVE

MAIN TANKONLY

MAIN TANKONLY

OPEN

CLOSED

FIREWALLSHUTOFF

VALVE

CLOSED

OPEN

FIREWALLSHUTOFF

VALVE

Figure 5-8 Firewall Shutoff Valve Switches


5-11

The crossfeed system is controlled by a three-position switch placarded: CROSSFEED OPEN,AUTO, and CLOSED. The valve can be manu-ally opened or closed, but under normal flightconditions it is left in the AUTO position. In theAUTO position, the fuel pressure switches areconnected into the crossfeed control circuit.

In the event of a boost pump failure, causing adrop in fuel pressure, these switches open thecrossfeed valve allowing the remaining boostpump to supply fuel to both engines.

In the event of a boost pump failure during take-off , the sys tem wi l l beg in to c ross feedautomatically allowing the pilot to complete thetakeoff without an increase in workload at a cru-cial time. After the takeoff is completed, or if theboost pump fails after takeoff, the crossfeedswitch may be closed and the flight continuedrelying on the engine-driven high pressure pumpwithout boosted pressure. In some instances, the

pilot may elect to continue the flight with theremaining boost pump and the crossfeed systemin operation.

When the crossfeed switch on the fuel controlpanel is actuated, power is drawn from a5-ampere circuit breaker on the fuel control panelto the solenoid that opens the crossfeed valve.The crossfeed is also powered through the hotbattery bus through a 5-amp fuse.

When the crossfeed valve is receiving power, theyellow FUEL CROSSFEED light on the annun-ciator panel will illuminate. The crossfeed willnot transfer fuel from one wing to another; itsfunction is to supply fuel from one side to theopposite engine during a boost pump failure oran engine-out condition. If the boost pumps onboth sides are operating and the crossfeed valveis open, fuel will be supplied to the engines in thenormal manner because the pressure on each sideof the crossfeed valve should be equal.

Figure 5-9 Crossfeed Schematic

5-12


FUEL DRAIN PURGE SYSTEM

The fuel purge system (Figure 5-10) is designedto assure that any residual fuel in the fuel mani-folds is consumed during engine shutdown.During engine starting, fuel manifold pressurecloses the fuel manifold poppet valve, allowingP

3

air to pressurize the purge tank. Duringengine operation, engine compressor air (P

3

air)is routed through a filter and check valve andmaintains pressurization of the small purgetank. Upon engine shutdown, fuel manifoldpressure subsides, thus allowing the engine fuelmanifold poppet valve to open. The pressuredifferential between the purge tank and fuelmanifold causes air to be discharged from thepurge tank, forcing residual fuel out of theengine fuel manifold lines, through the nozzles,and into the combustion chamber. As the fuel isburned, a momentary surge in (N

l

) gas generatorrpm should be observed. The entire operation isautomatic and requires no input from the crew.

FUEL GAGING SYSTEM

The airplane is equipped with a capacitance-typefuel quantity indication system (Figure 5-11). Itautomatically compensates for fuel temperature-

Figure 5-11 Fuel Quantity Indication System

Figure 5-10 Fuel Drain Purge System Schematic


5-13

density variations. The left fuel quantity indicator,on the fuel control panel, indicates the amount offuel remaining in the left-side fuel system tankswhen the FUEL QUANTITY select switch is inthe TOTAL (upper) position, and the amount offuel remaining in the left-side nacelle fuel tankwhen the FUEL QUANTITY select switch is inthe NACELLE (lower) position. The right fuelquantity indicator indicates the same informationfor the right-side fuel systems, depending upon theposition of the FUEL QUANTITY switch. Thegages are marked in pounds.

The fuel quantity indicating system is a capaci-tance type that is compensated for specificgravity and reads in pounds on a linear scale. Anelectronic circuit in the system processes the sig-nals from the fuel quantity (capacitance) probes(Figure 5-12) in the various fuel cells for anaccurate readout by the fuel quantity indicators.A selector switch, located between the fuel quan-tity indicators in the fuel panel beside the pilot,may be set in either the TOTAL or NACELLEpositions to determine whether the gages indicatethe pounds of fuel in the nacelle and wing fuelcells of the fuel system, or the pounds of fuel inonly the nacelle fuel cell.

COMPONENTS AND OPERATION

Each side of the airplane has an independent gag-ing system consis t ing of a fuel quant i ty(capacitance) probe in the nacelle fuel cell, one inthe aft-inboard fuel cell, two in the leading-edgefuel cell, and one in the center-section fuel cell.

When the fuel selector switch is left in itsTOTAL position, power is supplied from a5-ampere circuit breaker (on the fuel panel)through the fuel quantity indicator to all of thecapacitance probes in the fuel system. When thefuel selector switch is placed in the NACELLEposition, power is then supplied through the fuelquantity indicator to the capacitance probe in thenacelle fuel cell only.

Fuel density and electrical dielectric constantlyvary with respect to temperature, fuel type, andfuel batch. The capacitance gaging system isdesigned to sense and compensate for these vari-ables. The fuel quantity probe is simply avariable capacitor comprised of two concentrictubes. The inner tube is profiled by changing thediameter as a function of height so that the capac-itance between the inner and outer tube isproportional to the tank volume. The tubes serveas fixed electrodes and the fuel of the tank in thespace between the tubes acts as the dielectric ofthe fuel quantity probe.

The capacitance of the fuel quantity probe varieswith respect to the change in the dielectric thatresults from the ratio of fuel-to-air in the fuelcell. As the fuel level between the inner and outertubes rises, air with a dielectric constant of one isreplaced by fuel with a dielectric constant ofapproximately two, thus increasing the capaci-tance of the fuel quantity probe. This variation inthe volume of fuel contained in the fuel cell pro-duces a capacitance variation that actuates thefuel quantity indicator.

Figure 5-12 Fuel Probe

5-14


FUEL DRAINS

During each preflight, the fuel sumps on thetanks, pumps and filters or strainers should bedrained to check for fuel contamination.There are four sump drains and one filterdrain or strainer drain in each wing (Figures5-13 and 5-14).

The leading edge tank sump has a drain on theunderside of the outboard wing just forward ofthe main spar. The flush drain valve for thefirewall fuel strainer drain is accessible on theunderside of the engine cowling. The boostpump sump drain is at the bottom center of thenacelle, just forward of the wheel well. Thewheel well sump drain is inside the wheel wellon the gravity feed line. The drain for thetransfer pump sump is just outboard of thewing root, forward of the flap.

When draining the flush-mounted drains, do notturn the draining tool. Turning or twisting of thedraining tool will unseat the O-ring seal andcause a leak.

The flush valve attached to the base of the fuelstrainer can be opened or closed with a coin, ascrew driver, or a fuel drain tool making it possi-ble to drain fuel from the fuel strainer forpreflight check.

Since jet fuel and water are of similar densities,water does not settle out of jet fuel as easily asfrom aviation gasoline. For this reason, the air-plane must sit perfectly still, with no fuel beingadded, for approximately three hours prior todraining the sumps if water is to be removed.Although turbine engines are not so critical asreciprocating engines regarding water ingestion,water should still be removed periodically to pre-vent formations of fungus and contaminationinduced inaccuracies in the fuel gaging system.

FUEL HANDLING PRACTICES

Takeoff is prohibited when the fuel-quantity indi-cator needles are in the yellow arc, with theselector in the total position, or when there is lessthan 265 pounds of fuel in each wing system.

Figure 5-13 Fuel Drains


5-15

The King Air C90A and C90B require that bothboost pumps must be operable prior to takeoff.

All hydrocarbon fuels contain some dissolvedand some suspended water. The quantity of watercontained in the fuel depends on temperature andthe type of fuel. Kerosene, with its higher aro-matic content, tends to absorb and suspend morewater than aviation gasoline. In addition to water,it will suspend rust, lint and other foreign materi-als longer. Given sufficient time, these suspendedcontaminants will settle to the bottom of the tank.

The settling time for kerosene is five times that ofaviation gasoline; therefore, jet fuels requiregood fuel-handling practices to assure that theairplane is serviced with clean fuel. If recom-

mended ground procedures are carefullyfollowed, solid contaminants will settle and freewater can be reduced to 30 parts per million(ppm), a value that is currently accepted by themajor airlines.

Since most suspended matter can be removedfrom the fuel by sufficient settling time andproper filtration, it is not a major problem. Dis-solved water has been found to be the major fuelcontamination problem. Its effects are multipliedin aircraft operating primarily in humid regionsand warm climates.

Dissolved water cannot be filtered from the fuelby micronic-type filters, but can be released bylowering the fuel temperature, which will occur

Figure 5-14 Fuel Drain Locations

5-16


in flight. For example, a kerosene fuel may con-tain 65 ppm (8 fluid ounces per 1000 gallons) ofdissolved water at 80º F. When the fuel tempera-ture is lowered to 15º F, only about 25 ppm willremain in solution. The difference of 40 ppm willhave been released as supercooled water dropletswhich need only a piece of solid contaminant oran impact shock to convert them to ice crystals.

Tests indicate that these water droplets will notsettle during flight and are pumped freelythrough the system. If they become ice crystals inthe tank, they will not settle since the specificgravity of ice is approximately equal to that ofkerosene. The 40 ppm of suspended water seemslike a very small quantity, but when added to sus-pended water in the fuel at the time of delivery, itis sufficient to ice a filter. While the critical fueltemperature range is from 0 to –20º F, which pro-duces severe system icing, water droplets canfreeze at any temperature below 32º F.

Even if the fuel does not contain water or youhave drained the water out, there is still thepossibility of fuel icing at very low tempera-

tures. The oil-to-fuel heat exchanger is used toheat the fuel prior to entering the fuel controlunit. Since no temperature measurement isavailable for fuel prior to the heat exchanger,the temperature must be assumed to be thesame as the outside air temperature.

The graph in the Limitations section of the


is used as a guide inpreflight planning, based on known or forecastconditions, to determine operating temperatureswhere icing at the fuel control unit could occur.Enter the graph with the known or forecast Out-side Air Temperature and plot vertically to thegiven pressure altitude. In this example (Figure5-15), Outside Air Temperature equals minusthirty degrees Celsius and pressure altitudeequals 5000 feet. Next, plot horizontally to deter-mine the minimum oil temperature required toprevent icing. In this example, the minimum oiltemperature required is 38 degrees Celsius. If theplot should indicate that oil temperature versusOutside Air Temperature is such that ice forma-tion could occur during takeoff or in flight, anti-icing additive must be mixed with the fuel.

Figure 5-15 Fuel Temperature Graph


5-17

The King Air maintains a constant oil tempera-ture, however, this temperature varies from oneaircraft to another. For most aircraft the oil tem-perature will be between 50 and 60 degreesCelsius. Compare the minimum oil temperatureobtained from this graph with the oil tempera-ture achieved by each particular airplaneinvolved. When required, only anti-icing addi-tive conforming to Specification MIL-I-27686is approved. The anti-icing additive should beadded during fueling.

Water in jet fuel also creates an environmentfavorable to the growth of a microbiological“sludge” in the settlement areas of the fuel cells.This sludge, plus other contaminants in the fuel,can cause corrosion of metal parts in the fuel sys-tem as well as clogging of the fuel filters.Although this airplane uses bladder-type fuelcells, and all metal parts (except the boost pumpsand transfer pumps) are mounted above the set-tlement areas, the possibility of filter cloggingand corrosive attacks on fuel pumps exists if con-taminated fuels are consistently used.

Fuel biocide-fungicide “BIOBORJF” in concen-trations noted in the

POH

may he used in the fuel.BIOBORJF may be used as the only fuel additiveor it may be used with the anti-icing additive con-forming to MIL-I-27686 specification. Usedtogether, the additives have no detrimental effecton the fuel system components.

The primary means of fuel contamination controlby the owner/ operator is “good housekeeping.”This applies not only to fuel supply, but to keep-ing the aircraft system clean. The following is alist of steps that may be taken to recognize andprevent contamination problems.

1. Know your supplier. It is impractical toassume that fuel free from contaminants willalways be available, but it is feasible to exer-cise caution and be watchful for signs of fuelcontamination.

2. Assure, as much as possible, that the fuelobtained has been properly stored, that it isfiltered as it is pumped to the truck, and againas it is pumped from the truck to the aircraft.

3. Perform filter inspections to determine ifsludge is present.

4. Maintain good housekeeping by periodicallyflushing the fuel tanks and systems. The fre-quency of flushing will be determined by theclimate and the presence of sludge.

5. Aviation gas is an emergency fuel. The150 hours maximum operation on aviationgasoline per a “Time Between Overhaul”should be observed.

6. Use only clean fuel-servicing equipment.

7. After refueling, allow a settling period of atleast four hours whenever possible, thendrain a small amount of fuel from each drain.

CAUTION

Remove spilled fuel from the ramparea immediately to prevent the con-taminated surface from causing tiredamage.

When fueling the King Air C90A or C90B, thenacelle fuel tanks should be filled first before anyfuel is put in the wing tank system to insure thatthe wing tanks are completely full.

FUEL GRADES AND ADDITIVES

Aviation Kerosene Grades Jet A, Jet A-1, Jet B,JP-4, JP-5, and JP-8 may be mixed in any ratio.Aviation Gasoline Grades 80 (80/87), 100LL,100 (100/130), and 115/145 are emergency fuelsand may be mixed with the recommended fuelsin any ratio; however, use of the lowest octanerating available is suggested. Operation on avia-tion gasoline shall be limited to 150 hours perengine during each Time Between Overhaul(TBO) period.

If the King Air C90A or C90B is fueled with avi-ation gasoline, some operational limitations,which are listed in the

POH

, must be observed.Maximum operation with aviation gasoline islimited to 150 hours between engine overhauls.

5-18


Use of aviation gas is limited to 150 hours due tolead deposits which form on the turbine wheelsduring aviation gas consumption, and whichcause power degradation. Since the aviation gaswill probably be mixed with jet fuel already inthe tanks, it is important to record the number ofgallons of aviation gas taken aboard for eachengine. Determine the average fuel consumptionfor each hour of operation. If, for example, anengine has an average fuel consumption of 40gallons per hour, each time 40 gallons of aviationgasoline are added, one hour of the 150 hour lim-itation is being used. In other words, using the 40gph consumption rate as an example, the engineis allowed 6000 gallons of aviation gasolinebetween overhauls.

If the tanks have been serviced with aviationgas, flights are limited to 8,000 feet pressurealtitude or below if either boost pump is inoper-ative. Because it is less dense, aviation gasdelivery is much more critical than jet fueldelivery. Aviation gas feeds well under pressurefeed but does not feed well on suction feed -par-ticularly at high altitudes. For this reason, analternate means of pressure feed must be avail-able for aviation gas at high altitude. Thisalternate means is crossfeed from the oppositeside. Thus, a crossfeed capability is required forclimbs above 8,000 feet pressure altitude. Theselimitations are found in the Limitations sectionof your


.

The POH lists two approved fuel additives for theKing Air C90A and C90B. Any anti-icing addi-tive conforming to Specification MIL-L-127686is approved as is the fuel biocide-fungicideBIOBORJF. Each additive may be used as theonly fuel additive or they may be used together. Ithas been determined that, used together, the addi-tives have no detrimental effect on the fuelsystem components.

Additive concentrations and blending proceduresare found in the

King Air 90 MaintenanceManual

.

The FUEL BRANDS AND TYPE DESIGNA-TIONS chart in the Handling, Service &Maintenance section of the

POH

gives the fuelrefiner’s brand names, along with the corre-sponding designations established by theAmerican Petroleum Institute (APT) and theAmerican Society of Testing Material (ASTM).The brand names are listed for ready referenceand are not specifically recommended by BeechAircraft Corporation. Any product conforming tothe recommended specification may be used.

FILLING THE TANKS

When filling the aircraft fuel tanks, alwaysobserve the following:

1. Make sure the aircraft is statically groundedto the servicing unit and to the ramp.

2. Service the nacelle tank on each side first.The nacelle tank filler caps are located at thetop of each nacelle. The wing tank filler capsare located in the top of the wing, outboard ofthe nacelles.

NOTE

Servicing the nacelle tanks first pre-vents fuel transfer through the gravityfeed interconnect lines from the wingtanks into the nacelle tanks duringfueling. If wing tanks are filled first,fuel will transfer from them into thenacelle tank leaving the wing tanksonly partially filled. Be sure thenacelle tanks are completely full afterservicing the fuel system to assureproper automatic fuel transfer duringflight operation.


5-19

3. Allow a four-hour settling period wheneverpossible, then drain a small amount of fuelfrom each drain point. Check fuel at eachdrain point for contamination.

DRAINING THE FUEL SYSTEM

Open each fuel drain daily to drain off any wateror other contamination collected in the lowplaces. Along with the drain on the firewallmounted fuel filter, there are four other drains:the nacelle tank fuel-pump drain, center-sectiontank transfer-pump drain, wheelwell drain, andthe inboard end of the outboard-wing tank drain.

The fuel pump and tank drains are accessiblefrom the underside of the airplane.

NOTE

The firewall shutoff valve has to beelectrically opened to drain large quan-tities of fuel from the firewall fuel-filter drain.

Fuel may be drained from the tanks by gravityflow through the center-section transfer-pumpdrains into suitable containers. Fuel may also bypumped out of the tanks utilizing an externalpump and suction hoses inserted into the filleropenings. For the fastest means of draining thesystem see the procedures in the Beechcraft KingAir 90 Series Maintenance Manual.


6-1

The material normally covered in this chapter is notapplicable to this airplane.


7-i

CHAPTER 7

CONTENTS

Page

POWERPLANT

INTRODUCTION ..................................................................................................................

7-1

GENERAL ..............................................................................................................................

7-1

ENGINES................................................................................................................................

7-3

General ............................................................................................................................

7-3

Turboprop Engine Ratings ..............................................................................................

7-4

Engine Terms ..................................................................................................................

7-5

Free-Turbine Reverse-Flow Principle .............................................................................

7-5

Engine Airflow ................................................................................................................

7-6

Engine Stations................................................................................................................

7-9

Engine Modular Concept.................................................................................................

7-9

Compressor Bleed Valve...............................................................................................

7-10

Igniters...........................................................................................................................

7-11

Accessory Section ........................................................................................................

7-11

Lubrication System........................................................................................................

7-13

Engine Fuel System.......................................................................................................

7-16

Fuel Control Unit...........................................................................................................

7-18

Fuel Pressure Indicators ................................................................................................

7-19

Fuel Flow Indicators......................................................................................................

7-20

Anti-icing Fuel Additive ...............................................................................................

7-20

Engine Power Control ...................................................................................................

7-20

ITT and Torquemeters...................................................................................................

7-20

7-ii


ITT Gage.......................................................................................................................

7-20

Torquemeter..................................................................................................................

7-21

Gas Generator Tachometer (N

1

) ...................................................................................

7-22

Control Pedestal ............................................................................................................

7-22

Engine Limitations........................................................................................................

7-23

Starter Operating Time Limits ......................................................................................

7-25

Data Collection Form....................................................................................................

7-27

PROPELLERS .....................................................................................................................

7-27

General..........................................................................................................................

7-27

Propeller System ...........................................................................................................

7-28

McCauley and Hartzell Four-Blade Propellers.............................................................

7-28

Blade Angle ..................................................................................................................

7-28

Primary Governor .........................................................................................................

7-28

Low Pitch Stop..............................................................................................................

7-34

Beta and Reverse Control .............................................................................................

7-36

Overspeed Governor .....................................................................................................

7-38

Fuel Topping Governor ................................................................................................

7-39

Power Levers ................................................................................................................

7-39

Propeller Control Levers...............................................................................................

7-40

Autofeather System.......................................................................................................

7-41

Propeller Synchrophaser System ..................................................................................

7-41

Propeller Synchroscope ................................................................................................

7-45


7-iii

Figure Title Page

ILLUSTRATIONS

7-1

Powerplant Installation ............................................................................................

7-2

7-2

Engine Installation ...................................................................................................

7-3

7-3

PT6A-21 Specifications...........................................................................................

7-4

7-4

Free Turbine.............................................................................................................

7-5

7-5

Engine Cutaway.......................................................................................................

7-6

7-6

Engine Stations ........................................................................................................

7-6

7-7

Engine Orientation...................................................................................................

7-7

7-8

Engine Gas Flow......................................................................................................

7-7

7-9

Power and Compressor Sections .............................................................................

7-8

7-10

Engine Construction ................................................................................................

7-8

7-11

Typical Engine Modular Construction ....................................................................

7-9

7-12

Compressor Bleed Valve .......................................................................................

7-10

7-13

Engine Start and Ignition Switches........................................................................

7-11

7-14

Typical PT6A Engine ............................................................................................

7-12

7-15

Engine Lubrications Diagram................................................................................

7-14

7-16

Engine Oil Dipstick ...............................................................................................

7-15

7-17

Magnetic Chip Detector.........................................................................................

7-16

7-18

Simplified Fuel System Diagram...........................................................................

7-17

7-19

Simplified Fuel Control System ............................................................................

7-18

7-20

Fuel Pressure Annunciators ...................................................................................

7-19

7-21

Fuel Flow Indicator................................................................................................

7-20

7-22

Control Levers .......................................................................................................

7-21

7-23

Engine Instrument Markings .................................................................................

7-21

7-iv


7-24

Control Pedestal ....................................................................................................

7-22

7-25

Engine Limits Chart ..............................................................................................

7-24

7-26

Overtorque Limits Chart .......................................................................................

7-25

7-27

Overtemperature Limits (Starting) ........................................................................

7-25

7-28

Overtemperature Limits (Except Starting)............................................................

7-26

7-29

View through Exhaust Duct ..................................................................................

7-26

7-30

In-Flight Engine Data Log ....................................................................................

7-27

7-31

Propellers...............................................................................................................

7-29

7-32

Propeller Tiedown Boot Installed .........................................................................

7-30

7-33

Primary Governor Diagram...................................................................................

7-30

7-34

Blade Angle Diagram............................................................................................

7-31

7-35

Propeller Onspeed Diagram ..................................................................................

7-33

7-36

Propeller Overspeed Diagram ...............................................................................

7-33

7-37

Propeller Underspeed Diagram .............................................................................

7-34

7-38

Low Pitch Stop Diagram.......................................................................................

7-35

7-39

Beta Range and Reverse Diagram.........................................................................

7-37

7-40

Overspeed Governor Diagram ..............................................................................

7-39

7-41

Power Levers.........................................................................................................

7-40

7-42

Propeller Control Levers .......................................................................................

7-40

7-43

Autofeather System Diagram—Left Engine Failed and Feathering .....................

7-42

7-44

Autofeather System Diagram—Armed.................................................................

7-43

7-45

Autofeather Test Diagram.....................................................................................

7-44

7-46

Propeller Synchrophaser .......................................................................................

7-45


7-1

CHAPTER 7POWERPLANT

INTRODUCTION

In-depth knowledge of the powerplants is essential to good power management by the pilot.Knowing and operating within safe parameters of the powerplant and propeller system extendsengine life and ensures safety. This chapter describes the basic sections of the engine and itsoperational limits and preflight checks.

In-depth knowledge of the propeller system is also essential to proper operation of the enginepower system. Operating within safe parameters of the powerplant and propeller systemsextends engine life and ensures safety. This chapter also describes the propeller system and itsoperational limits and preflight checks.

GENERAL

The Engines section of this chapter presents adescription and discussion of the Pratt and WhitneyPT6A turboprop engines. The engines used onthese airplanes will be described in sufficient detailfor flight crewmembers to understand normal oper-ational practices and limitations. The purpose of

this section is to give the participants a sufficientunderstanding of the engine so that they will befamiliar with normal and emergency procedures.

The Propellers section of this chapter presents adescription and discussion of the propeller sys-tem. Location and use of propeller controls,principle of operation, reversing, and featheringare included.

��

#1 DCGEN

7-2


Figure 7-1 Powerplant Installation


7-3

ENGINES

GENERAL

The powerplants chosen by Beech designers forthe King Airs are Pratt and Whitney Series PT6Afree-turbine turboprop engines (Figures 7-1 and7-2) . The King Air C90A and C90B usePT6A-21 engines. The PT6A-21 engine is flat-rated to 550 shaft horsepower.

The engines are equipped with conventionalthree-blade (C90A) or four-blade (C90B), full-feathering, reversing, variable-pitch propellersmounted on the output shaft of the engine reduc-tion gearbox. The propeller pitch and speed arecontrolled by engine oil pressure through single-action, engine-driven propeller governors. Thepropellers will feather automatically when theengines are shut down on the ground, and willunfeather when the engines are started.

Figure 7-2 Engine Installation

7-4


When reference is made to the right or left side ofthe airplane or engine, it is always looking fromthe rear to the front.

TURBOPROP ENGINE RATINGS

In turboprop engines, power is measured inEquivalent Shaft Horse Power (ESHP) andShaft Horse Power (SHP). SHP is determined

by propeller rpm and torque applied to turn thepropeller shaft. The hot exhaust gases alsodevelop some kinetic energy as they leave theengine, similar to a turbojet engine. This jetthrust amounts to about 10% of the total enginepower. ESHP is the term applied to total powerdelivered, including the jet thrust. Turbopropengine specifications usually show both ESHPand SHP, along with limiting ambient tempera-tures. The engine specifications in Figure 7-3

Figure 7-3 PT6A-21 Specifications


7-5

Primary Governor Diagram show the engine rat-ings and temperatures.

ENGINE TERMS

To properly understand the operation of thePT6A series engines, there are several basicterms you should know:

●

N

1

or N

G

—Gas generator rpm is percentof turbine speed

●

N

2

or Np—Propeller rpm

●

N

F

—Power turbine rpm (not indicated onengine instruments)

●

P

3

—Air pressure at station three (thesource of bleed air)

●

ITT or T

5

—Interstage Turbine Tempera-ture in degrees of temperature at station 5

Review and remember these terms. They will beused often to describe PT6A engines.

FREE-TURBINE REVERSE-FLOW PRINCIPLE

The Pratt and Whitney PT6 family of enginesconsists basically of free-turbine, reverse-flowengines driving a propeller through planetarygearing (Figures 7-4, 7-5, 7-6, and 7-7). Theterm “free-turbine” refers to the design of theturbine sections of the engine. There are twoturbine sections: one, called the compressor tur-bine, which drives the engine compressor andaccessories; and the other, consisting of a singlepower turbine, which drives the power sectionand propeller. The power turbine section has nophysical connection to the compressor turbineat all. These turbines are mounted on separateshafts and are driven in opposite directions bythe gas flow across them. The term “reverseflow” refers to airflow through the engine. Inletair enters the compressor at the aft end of the

Figure 7-4 Free Turbine

7-6


engine, moves forward through the combustionsection and the turbines, and is exhausted at thefront of the engine.

ENGINE AIRFLOW

Inlet air enters the engine through an annular ple-num chamber, formed by the compressor inletcase, where it is directed forward to the compres-sor (Figures 7-8, 7-9, and 7-10). The compressorconsists of three axial stages combined with a

single centrifugal stage, assembled as anintegral unit.

A row of stator vanes, located between eachstage of compression, diffuses the air, raises itsstatic pressure, and directs it to the next stage ofcompression. The compressed air passes throughdiffuser tubes, which turn the air through 90° indirection and convert velocity to static pressure.The diffused air then passes through straighten-ing vanes to the annulus surrounding thecombustion chamber liner.

Figure 7-5 Engine Cutaway

Figure 7-6 Engine Stations


7-7

The combustion chamber liner has varying sizeperforations which allow entry of compressordelivery air. Approximately 25% of the air mixeswith fuel to support combustion. The remaining75% centers the flame in the combustion cham-ber and provides internal cooling for the engine.As it enters the combustion area and mixes withfuel, the flow of air changes direction 180°. Thefuel/air mixture is ignited, and the resultant

expanding gases are directed to the turbines. Thelocation of the liner eliminates the need for along shaft between the compressor and the com-pressor turbine, thus reducing the overall lengthand weight of the engine.

During normal operation, fuel is injected intothe combustion chamber liner through 14 sim-plex nozzles, which are supplied by a dual

Figure 7-7 Engine Orientation

Figure 7-8 Engine Gas Flow

7-8


manifold consisting of primary and secondarytransfer tubes and adapters. During starting, thefuel/air mixture is ignited by two spark igniterswhich protrude into the liner. After starting, theigniters are turned off, since combustion is self-sustaining. The resultant gases expand from theliner, reverse direction in the exit duct zone, and

pass through the compressor turbine inlet guidevanes to the single-stage compressor turbine.The guide vanes ensure that the expandinggases impinge on the turbine blades at the cor-rect angle, with minimum loss of energy. Theexpanding gases are then directed forward todrive the power turbine section.

Figure 7-9 Power and Compressor Sections

Figure 7-10 Engine Construction


7-9

The single-stage power turbine, consisting of aninlet guide vane and turbine, drives the propellershaft through a reduction gearbox.

The compressor and power turbines are located inthe approximate center of the engine, with theirrespective shafts extending in opposite directions.This feature simplifies the installation and inspec-tion procedures. The exhaust gas from the powerturbine is directed through an annular exhaust ple-num to atmosphere through twin opposed exhaustports provided in the exhaust duct.

ENGINE STATIONS

To identify points in the engine, it is commonpractice to establish engine station numbers atvarious points (Figure 7-6). To refer to pressureor temperature at a specific point in the engineairflow path, the appropriate station number isused, such as P

3

for the Station 3 pressure or T

5

for the gas temperature at Station 5. For instance,temperature of the airflow is measured between

the compressor turbine and the power turbine atEngine Station Number 5. This is called Inter-stage Turbine Temperature (ITT) or T

5

. Bleed airis taken off the engine after the centrifugal com-pressor s t age and p r io r to en te r ing thecombustion chamber. This air, commonlyreferred to as P

3

air, is used for cabin heat, pres-surization, and the pneumatic system.

ENGINE MODULAR CONCEPT

With the modular free-turbine design, the engineis basically divided into two modules: a gas gen-erator section and a power section (Figure 7-11).The gas generator section includes the compres-sor and the combustion section. Its job is to drawair into the engine, add energy to it in the form ofburning fuel, and produce the gases necessary todrive the compressor and power turbines.

The power section’s job is to convert the gas flowfrom the gas generator section into mechanicalaction to drive the propeller. This is done through

Figure 7-11 Typical Engine Modular Construction

7-10


an integral planetary gearbox, which converts thehigh speed and low torque of the power turbine tothe low speed and high torque required at thepropeller. The reduction ratio from power turbineshaft rpm to propeller rpm is approximately 15:1.

COMPRESSOR BLEED VALVE

At low N

1

rpm, the axial compressors producemore compressed air than the centrifugal com-pressor can effectively handle (accept). Acompressor bleed valve compensates for thisexcess airflow at low rpm by opening, to relievethis pressure. As compressor speed increases,the valve closes proportionally until, at 80% N

1

,the valve is fully closed (Figure 7-12). Thispressure relief helps prevent compressor stall ofthe centrifugal stage.

The compressor bleed valve is a pneumatic pis-ton which references the pressure differentialbetween the axial and centrifugal stages. Look-ing forward, the valve is located at the 6 o’clockposition. The function of this valve is to preventcompressor stalls and surges in the low N

1

rpmrange (75 to 80% N

1

).

At low N

1

rpm, the valve is in the open position.At takeoff and cruise N

1

rpm, above approxi-mately 80%, the bleed valve will be closed. Ifthe compressor bleed valve sticks closed, acompressor stall will result. If the valve sticksopen, the ITT would be noticably higher as thepower lever is advanced above 80% N

1

.

Figure 7-12 Compressor Bleed Valve


7-11

IGNITERS

The engine start switches are located on thepilot’s left subpanel (Figure 7-13). This subpanelcontains the IGNITION AND ENGINE STARTswitches and ENG AUTO IGNITION switches.

The IGNITION AND ENGINE START switcheshave three positions: ON, OFF, and STARTERONLY. The ON position is lever-locked and acti-vates both the starter and igniters. The STARTERONLY position is a momentary hold-down posi-tion of the spring-loaded-to-center OFF position.It provides for motoring only to clear the engineof unburned fuel. With the switch in this position,there is no ignition.

The combustion chamber has two spark-typeigniters to provide positive ignition duringengine start. While the engine is equipped withtwo igniters, it will start with only one. Thesystem is designed so that if one igniter is openor shorted, the remaining igniter will continueto function. Once the engine is started, theigniters are de-energized, since the combustionis self-sustaining.

The ignition system features an automaticbackup function for emergencies. This backupsystem is called “autoignition.” The ENG AUTOIGNITION switches should be moved to theARM position just prior to takeoff. If enginetorque falls below approximately 400 ft-lb, theigniter will automatically energize, attempting torestart the engine. The IGNITION ON annuncia-tor will be illuminated.

The spark ignition provides the engine with anignition system capable of quick light-ups over awide temperature range. The system consists ofan airframe-mounted ignition exciter, two indi-vidual high-tension cable assemblies, and twospark igniters. It is energized from the aircraftnominal 28-VDC supply and will operate in the9- to 30-volt range. The igniter control box pro-duces up to 3,500 volts. The ignition exciter isenergized only during the engine startingsequence and emergencies to initiate combustionin the combustion chamber.

ACCESSORY SECTION

Most of the engine-driven accessories, except thepropeller governors and propeller tach generator,are mounted on the accessory gearbox located atthe rear of the engine (Figure 7-14). The accesso-ries are driven from the compressor shaft througha coupling shaft.

Figure 7-13 Engine Start and Ignition

Switches

7-12


Figure 7-14 Typical PT6A Engine


7-13

The lubricating and scavenge oil pumps aremounted inside the accessory gearbox, with theexception of the two scavenge pumps which areexternally mounted.

The starter-generator, high-pressure fuelpump, N

1

tachometer generator, and otheroptional accessories are mounted on pads onthe rear of the accessory drive case. There areseven such mounting pads, each with its owndifferent gear ratio.

LUBRICATION SYSTEM

The PT6A engine lubrication system has a dualfunction (Figure 7-15). Its primary function isto cool and lubricate the engine bearings andbushings. Its second function is to provide oil tothe propeller governor and propeller reversingcontrol system.

The main oil tank houses a gear-type engine-driven pressure pump, oil pressure regulator, andoil filter. The engine oil tank is an integral part ofthe compressor inlet case and is located in frontof the accessory gearbox.

The oil tank is provided with a filler neck andintegral quantity dipstick housing. The cap anddipstick are secured to the filler neck, whichpasses through the gearbox housing and acces-sory diaphragm and into the tank. The markingson the dipstick indicate the number of U.S. quartsof oil less than full (Figure 7-16).

The engine oil system has a total capacity of 3.5U.S. gallons, including the 2.3-gallon oil tank.Maximum oil consumption is one quart every10 hours of operation. Normal oil consumptionmay be as little as 1 quart per 50 hours of operation.

The dipstick will indicate 1 to 2 1/2 quartsbelow full when the oil level is normal. Do notoverfill. When adding oil between oil changes,do not mix types or brands of oil due to the pos-sibility of chemical incompatibility and loss oflubricating qualities.

A placard inside the engine cover shows thebrand and type of oil used in that particularengine. Although the preflight checklist calls forchecking the oil level, which is required, the besttime to check oil quantity is shortly after shut-down, since oil levels are most accuratelyindicated at that time.

Oil level checks during preflight may requiremotoring the engine for a brief time for an accu-rate level reading. Each engine tends to seek itsown oil level. The pilot should monitor the oillevel to ensure proper operation.

As pressure oil leaves the tank, it passes throughthe pressure and temperature-sensing bulbsmounted on or near the rear accessory case. Theoil then proceeds to the various bearing compart-ments and nose case through an external oiltransfer line below the engine. Scavenge oilreturns from the nose case and the bearing com-partments to the gear-type oil scavenge pumps inthe accessory case through external oil transferlines, and through the external oil cooler belowthe engine.

The oil cooler is thermostatically controlled tomaintain the desired oil temperature. Anotherexternally mounted unit, the oil-fuel heatexchanger, uses hot engine oil to heat fuel beforeit enters the engine fuel system. When gas gener-a to r speeds a re above 72% N

1

, and o i ltemperatures are between 60 and 70º C, normaloil pressure is between 80 and 100 psi.

Magnetic Chip Detector

A magnetic chip detector is installed in the bot-tom of each engine nose gearbox (Figure 7-17).This detector will activate a yellow light on theannunciator panel, L CHIP DETECT or R CHIPDETECT, to alert the pilot of oil contamination.

C90B aircraft, engine parameters should be mon-i tored for abnormal indicat ions. I f suchindications are observed, appropriate check listaction should be taken.

C90A aircraft are equipped with red “CHIPDETECT” annunciator panel lights. A steady“CHIP

7-14

FOR

TRAIN

ING

PU

RPOSES

ON

LY

Figure 7-15 Engine Lubrications Diagram


7-15

Figure 7-16 Engine Oil Dipstick

7-16


DETECT” light requires the engine be shut downto prevent serious internal damage.

When a CHIP DETECT annunciator lightcomes on and s tays on, t imely act ion isrequired to prevent serious damage to theinternal engine components. The chip detectorindicates the presence of ferrous particles inthe propeller gearbox.

ENGINE FUEL SYSTEM

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 compli-cated. The engine fuel control system consistsof the main components shown in the block dia-gram (Figure 7-18). They are the electric low-pressure boost pump, oil-to-fuel heat exchanger,high-pressure fuel pump, fuel control unit, fuel

cutoff valve, flow divider, and dual fuel mani-fold with 14 simplex nozzles.

The PT6A-21 engine uses an electric low-pres-sure boost pump to supply a 30-psi head pressureto the high-pressure engine-driven fuel pump.This head pressure prevents fuel cavitation at thehigh-pressure pump. The fuel is also used forcooling and lubricating the pump. The oil-to-fuelheat exchanger uses warm engine oil to maintaina desired fuel temperature at the fuel pump inletto prevent icing at the pump filter. This is donewith automatic temperature sensors and requiresno action by the pilot.

Fuel enters the engine fuel system through theoil-to-fuel heat exchanger, and then flows into thehigh-pressure engine-driven fuel pump and oninto the fuel control unit (FCU).

The high-pressure fuel pump is an engine-drivengear-type pump with an inlet and outlet filter. Flowrates and pressures will vary with gas generator

Figure 7-17 Magnetic Chip Detector


7-17

(N

1

) rpm. Its primary purpose is to provide suffi-cient pressure at the fuel nozzles for a proper spraypattern during all modes of engine operation. Thehigh-pressure pump supplies fuel at approximately800 psi to the fuel side of the FCU.

Two valves included in the FCU ensure consis-tent and cool engine starts. When the ignition orstart system is energized, the purge valve is elec-trically opened to clear the FCU of vapors andbubbles. The excess fuel flows back to the nacellefuel tanks. The spill valve, referenced to atmo-spheric pressure, adjusts the fuel flow for coolerhigh-altitude starts.

Between the FCU fuel valve and the engine com-bustion chamber, the minimum pressurizing

valve in the FCU remains closed during startinguntil fuel pressure builds sufficiently to maintaina proper spray pattern in the combustion cham-ber. About 80 psi is required to open theminimum pressurizing valve. If the high pressurefuel pump should fail, the valve would close, andthe engine would flame out.

The fuel cutoff valve is located downstream fromthe minimum pressurizing valve in the FCU. Thisvalve is controlled by the condition lever, eitheropen or closed. There is no intermediate positionof this valve. For starting, fuel flows initiallythrough the flow divider to the 10 primary fuelnozzles in the combustion chamber. As theengine accelerates through approximately 40%N

1

, fuel pressure is sufficient to open the flow

Figure 7-18 Simplified Fuel System Diagram

7-18


divider to the 4 secondary fuel nozzles. At thistime all 14 nozzles are delivering atomized fuelto the combustion chamber. This progressivesequence of primary and secondary fuel nozzleoperation provides cooler starts. During enginestarting, there is a noticable increase in ITT whenthe secondary fuel nozzles are activated.

During engine shutdown, any fuel left in the mani-fold is forced out through the fuel nozzles and intothe combustion chamber by purge tank pressure.As the fuel is burned, a momentary increase in N

1rpm may be observed. The entire operation isautomatic and requires no input from the crew.

Fuel Control UnitThe fuel control unit (Figure 7-19), which isreferred to as the FCU, has multiple functions,

but its primary purpose is to meter proper fuelamounts to the fuel nozzles in all modes ofengine operation.

FCU operation will be simplified and describedbriefly here. For detailed description and opera-tion, refer to the Pratt & Whitney MaintenanceManual which applies to this engine.

The condi t ion lever se lec ts id le speedsbetween LOW IDLE ( 51% to 58% N1) toHIGH IDLE (70% N1), while the power leverselects speeds between idle and maximum,101.5% N1. These control levers influence theN1 governor and control N1 speed. The gover-nor uses pneumatic air (P3) pressure to controlengine speed. The governor controls the airpressure in the fuel control unit by varying theP3 leak rate.

Figure 7-19 Simplified Fuel Control System

TO FUELTOPPING

GOVERNORTO GRAVITYFEED LINE

FUELPURGE P3

P3INLET

FUEL SUPPLY

FLOW DIVIDER and DUMP VALVE

MINIMUMFLOWSTOP

ENGINE DRIVENFUEL PUMP

MINIMUMPRESSURIZING VALVE

FUEL CUT-OFFVALVE

CONDITION LEVER

POWER LEVER

PURGE VALVE

N1GOVERNOR


The P3 air chamber and fuel chamber are sepa-rated by a diaphragm, which has a needle valvemounted on it which is called the metering valve.As the diaphragm is influenced by varyingair/fuel pressures, the metering valve is reposi-tioned to achieve the desired fuel flow. The N1governor controls fuel flow by allowing some P3pressure to be leaked off at varying rates,depending on the desired fuel flow.

In an underspeed condition, the N1 governor actsto increase P3 air pressure. This repositions themetering valve, allowing more fuel to enter thecombustion chamber, increaseing N1.

In an overspeed condition, the N1 governorallows the P3 pressure to be reduced in theFCU, which repositions the metering valvereducing the fuel flow into the combustionchamber, decreasing N1.

Should the P3 air pressure be lost, due to a mal-function, the metering valve will be positioned tothe minimum flow stop. Minimum flow powerwould be approximately 48% N1. The powerlever and condition lever would then have noeffect on engine speed.

Fuel Pressure IndicatorsIn the event of an electric boost pump failure, therespective FUEL PRESS annunciator (Figure7-20) will illuminate and the master warninglight will flash. The FUEL PRESS light illumi-nates when outlet pressure at the boost pumpdecreases below about 10 psi. If the crossfeedswitch is in the AUTO position, the automaticcrossfeed feature will open the valve extinguish-ing the annunciator.

In the event of an engine-driven fuel pump (high-pressure) failure, the engine will flame out.

CAUTION

Engine operation with the FUELPRESS light on is limited to tenhours between overhaul or replace-ment of the engine-driven high-pressure fuel pump.

L FUEL PRESS

L DC GEN

L IGNITION ON

L OIL PRESS

L NO FUEL XFR

R IGNITION ON

RVS NOT READY

L AUTOFEATHER

R CHIP DETECT

LDG/TAXI LIGHT

R OIL PRESS

R NO FUEL XFR

R FUEL PRESS

R DC GEN

Figure 7-20 Fuel Pressure Annunciators


Fuel Flow IndicatorsFuel flow information is sensed by a transmitterin the engine fuel supply line, between the boostpump and the engine-driven high-pressure pump,and indicated on the fuel flow gage on the instru-ment panel (Figure 7-21). The gage indicates fuelflow in pounds-per-hour units times 100. There-fore when the needle indicates 2 on the dial, fuelflow is 200 pounds per hour. The fuel flow gagesare DC-powered.

Anti-icing Fuel AdditiveEngine oil is used to heat the fuel prior to enter-ing the FCU. Since no temperature measurementis available for the fuel at this point, it must beassumed to be the same as the Outside Air Tem-perature. The Minimum Oil Temperature chart issupplied for use as a guide in preflight planning,based on known or forecast operating conditions,to indicate operating temperatures where icing atthe FCU could occur. If the plot should indicatethat oil temperature versus OAT is such that iceformation could occur during takeoff or in flight,anti-icing additive per MIL-1-27686 should bemixed with the fuel at refueling to ensure safeoperation. Refer to the King Air MaintenanceManual for procedures to follow when blendinganti-icing additive with the airplane fuel.

Anti-icing additive conforming to SpecificationMIL-1-27686 is the only approved fuel additive.

ENGINE POWER CONTROLThe propeller lever adjusts the propeller governorto the desired propeller speed (Figure 7-22). Thepropeller will maintain the set speed by varyingthe blade angle. Torque is controlled by thepower lever acting on the N1 governor. When thepower lever is advanced, the N1 governor causesthe FCU to increase fuel flow, resulting in anincrease in engine speed.

ITT AND TORQUEMETERSPower management is relatively simple, with twoprimary operating limitations. The engines aretemperature and torque limited. During operationrequiring maximum engine performance, enginetorque and ITT operating parameters are affectedby ambient temperature and altitude: at cold tem-perature or low altitude, torque limits power; athot temperature or high altitude, ITT limitspower. Whichever limit is reached first, deter-mines the power available.

ITT GAGEThe ITT gage monitors the interstage turbinetemperature at station 5 (Figure 7-23). ITT is aprime limiting indicator of the amount of poweravailable from the engine under varying ambienttemperature and altitude conditions. The normaloperating range, indicated by the green arc on thegage, is 400 to 695º C. These limits also apply tomaximum continuous power. The maximumstarting temperature of 1,090º C is indicated bythe dashed red line on the instrument, or a reddiamond on LJ-1361, 1363 and after. This start-ing limit of 1,090º C is limited to two seconds.The ITT gages are self-energizing and do notrequire electrical power (LJ-1361, LJ-1363, andafter are DC-powered). The engines will be dam-aged if limiting temperatures indicated on theITT gage are exceeded.

Figure 7-21 Fuel Flow Indicator


TORQUEMETERThe torquemeter, which is marked in ft-lb, con-stantly measures rotational force applied to thepropeller shaft (Figure 7-23). The maximum per-missible sustained torque is 1,315 ft-lb, the redradial at the top of the green arc on the instru-ment. A transient torque limit of 1,500 ft-lb istime-limited to two seconds. Cruise torques varywith altitude and temperature.

Torque is measured by a hydromechanicaltorquemeter in the first stage of the reductiongearcase. Rotational force on the first-stage ringgear allows oil pressure to change in thetorquemeter chamber. The difference betweenthe torquemeter chamber pressure and reductiongear internal pressure accurately indicates the

Figure 7-22 Control Levers

7

6

9 10

12

245

8 ITTSTART

˚C X 100

16

14

6

4

2

12 FTLB X 100

0

810

TORQUE 0

TURBINE

%RPM

110

100

9080 50

40

30

20

0

70 60

.0

Figure 7-23 Engine Instrument Markings


torque being produced at the propeller shaft.The torque transmitter measures this torque andsends an AC signal to the instrument on theinstrument panel (DC signal on LJ-1361, LJ-1363, and after).

GAS GENERATOR TACHOMETER (N1)

The N1 gas generator tachometer measures therotational speed of the compressor shaft, in per-cent of rpm, based on 37,500 rpm at 100%(Figure 7-23). The face of this instrument con-sists of two dials: a smaller dial labeled from 0 to9, and a larger dial labeled from 0 to 100. Thesmaller dial is calibrated in 1% increments, andthe larger dial in 10% increments. Between 30and 100% on the larger dial, the increments arein gradations of 2%.

The N1 indicator is self-generating (LJ-1361, LJ-1363, and after are DC-powered). The tachome-ter generator sensing unit, located in the engineaccessory section, is geared down to supply N1speed information to the instrument panel to indi-cate the percent of N1 revolutions.

Maximum continuous gas generator speed is lim-ited to 38,100 rpm, which is 101.5% on the N1indicator. A transient speed up to 102.6%, 38,500rpm, is time-limited to 2 seconds, to provide abuffer for surges during engine acceleration.

CONTROL PEDESTALThe control pedestal extends between pilot andcopilot (Figure 7-24). The three sets of controllevers are left to right: the power levers, propellerlevers, and the condition levers.

Power LeversThe power levers (Figure 7-22) control enginepower, from idle to maximum power, by opera-tion of the N1 governor in the fuel control unit.Increasing N1 rpm results in increased enginepower. The power levers have three controlranges: flight, Beta, and reverse. The bottom ofthe flight range is at IDLE. When the levers are

lifted over the IDLE detent and pulled back,they control engine power through the Beta andreverse ranges. A selectable ground fine (orzero thrust) power lever gate position is pro-vided on the C90B.

Condition LeversThe condition levers have multiple positions:FUEL CUTOFF and LO IDLE through HI IDLE(Figure 7-22). At the FUEL CUTOFF position,fuel flow to its respective engine is cut off.

At LO IDLE, engine gas generator speed (N1)is a minimum of 51% on the C90A or 58% onthe C90B; at HI IDLE it is 70%. The levers canbe set anywhere between LOW IDLE andHIGH IDLE

Propeller LeversThe propeller levers are conventional in settingthe propeller rpm for takeoff, climb and cruise(Figure 7-22). The normal governing range is1,800 to 2,200 rpm. This airplane is equippedwith both manual and automatic propeller feath-ering systems. To feather a propeller manually,pull the propeller lever back past the friction

YAWENG

APENG

SR

I/2Ø

DN

L

UP

R

HDG NAV APPR B/C CLIMB

ALT ALT SEL VS IAS DSC

CABINPRESSDUMP

RUDDERBOOST

ELEVTRIM

TEST OFF

PRESS

OFF

NAV DATA

GSPTTG

ET

COURSETIMER

EFIS

POWER TEST

PREACT

XFR

S/S

DH

MAP MAP

ARC ARCHSI

WX

TST

CRSSEL

WARNING DE-PRESSURIZE CABINBEFORE LANDING

CABINALT

CABINALT

RATE

1000FT

MIN

MAX

UP

DN

0

5

CAUTIONREVERSE

ONLY WITHENGINESRUNNING

LIFT

LIFTIDLE

GDFINE

ELEVATOR

TRIM

REVERSE

FRICTIONLOCK

AILERON TABLEFT RIGHT

FRICTIONLOCK

PROP

CONDITION

UP

FEATHER

RUDDER TABLEFT RIGHT

F L A P

CMPST

NORMAL

DN

10

2

34

56

78

910

-1

01

10

2022

18

26

16

12 14

24

ACFT ALT1000 FT

FUELCUTOFF

TEST

+

5

UP

10

APPROACH

DOWN

PUSH PUSHHDG CRS

SYN C

DIR E CT

2 41.5

DOWN

60

80

5

31

5

31

5

3

1

5

310

GO AROUND

GEAR HORNSILENCE

TRIM HDG NAVARM DR APPRARM B/C VNAV 1/2 0

YAW DIS AP

YAWDIS ALT ALTARM VS GSARM IAS GA DSC CLM APDIS

Figure 7-24 Control Pedestal


detent into the red and white striped section ofthe quadrant. To unfeather, push the lever for-ward of the detent into the governing range. Thepropellers go to feathered position when theengines shut down because of the loss of oil pres-sure in the propeller dome.

Control Lever OperationThe engines are controlled from the cockpit byusing the propeller, power, and condition levers.Both the power and condition levers are con-nected to the N1 governing section of the FCU.Either lever will reset the FCU to maintain a newN1 rpm. For starting, the power levers are at theIDLE position, and the condition levers aremoved to the LO IDLE position to open the fuelcutoff valves and set the governor at LO IDLE.The condition levers are continuously variablefrom LO IDLE to HI IDLE. This variable operat-ing speed with power levers at IDLE enhancesengine cooling by maintaining a steady airflowthrough the engines. With the condition levers atLO IDLE, the power levers will select N1 rpmfrom LOW IDLE to 101.5%, the maximum fortakeoff. However, if the condition levers are at HIIDLE, the power levers can select N1 rpm onlyfrom 70 to 101.5%.

Moving the power or condition levers mostdirectly affects N1 rpm. As the power or condi-tion levers are advanced, ITT, torque, and fuelflow increases. These indicators are by-productsof the N1 speed maintained by the FCU. With thepower levers in a fixed position, N1 remains con-stant even in a climb or descent. However, ITT,torque, and fuel flow will vary with altitude,ambient air temperature, and propeller setting.

ENGINE LIMITATIONSAirplane and engine limits are described in the“Limitations” section of the POH (Figure 7-25).These limitations have been approved by theFederal Aviation Administration, and must beobserved in the operation of the Beechcraft KingAir C90A and C90B. The Engine Operating Lim-its chart gives the major operating limits. ThePower Plant Instrument Markings chart lists theminimum, normal, and maximum limits.

During engine start, temperature is the most criti-cal limit. The ITT starting limit of 1,090º C,represented on the ITT gage by a dashed red line,is limited to two seconds. During any start, if theindicator needle approaches the limit, the startshould be aborted before the needle passes thedashed red line. For this reason, it is helpful dur-ing starts to keep the condition lever out of theLO IDLE detent so that the lever can be quicklypulled back to FUEL CUTOFF.

Monitor oil pressure and oil temperature. Duringthe start, oil pressure should come up to the mini-mum red line at 40 psi quickly, but should notexceed the maximum at 100 psi. During normaloperation the oil temperature and pressure gagesshould be in the green arc normal operatingrange. The green arc extends from 80 to 100 psi.

Oil pressure between 40 and 80 psi is undesirable;it should be tolerated only for completion of theflight, and then only at a reduced power setting.

Oil pressure below 40 psi is unsafe; it requiresthat either the engine be shut down or that a land-ing be made as soon as possible, using minimumpower required to sustain flight.

For increased service life of engine oil, an oiltemperature between 74 and 80º C is recom-mended. A minimum oil temperature of 55º C isrecommended for oil-to-fuel heater operation attakeoff power. Oil temperature limits are –40 and+99º C. During extremely cold starts, oil pressuremay reach 200 psi. Refer to the Engine Limitschart in the POH for minimum oil temperatureoperation limitations.

During ground operations, ITT temperatures arecritical. With the condition levers at LO IDLE,high ITT can be corrected by reducing the DCgenerator and other N1 loads, then increasing theN1 rpm by advancing the condition levers to HIIDLE. The air conditioner, for example, draws aheavy load on both engines, and may have to betemporarily turned off. At approximately 70% N1rpm, the HI IDLE condition lever position willnormally reduce the ITT. At any N1 below 70%,there is an idle ITT restriction of 660º C maxi-mum. If an ITT above 660º C is observed when


running N1 below 70%, the generator loadshould be reduced and the N1 speed increasedbefore re-introducing a load on the engines.

At N1 speeds of 70% or more, the 660º Crestriction is removed, as airflow through theengine is sufficient.

In the climb, torque will decrease and ITT mayincrease slightly. The cruise climb and recom-mended normal cruise ITT limit is not placardedon the indicator. At altitude, the PerformanceChart numbers may not be attainable due to alti-tude and temperature variations.

Transient limits provide buffers for surges dur-ing engine acceleration. Torque and ITT have

an allowable excursion duration of two sec-onds. A momentary peak of 1,500 ft-lb and825º C is allowed for torque and ITT respec-tively during acceleration.

The Overtorque Limits Chart (Figure 7-26) showsactions required if torque limits are exceededunder all conditions. If the torque limits areexceeded for more than a few minutes, the gear-box can be damaged. The chart shows the specificlimits and action required if they are exceeded.

The Overtemperature Limits charts (Figures 7-27and 7-28) show the specific actions required ifITT limits are exceeded during Starting Condi-tions and All Conditions Except Starting. Forarea A (Figure 7-28 Overtemperature Limits

Figure 7-25 Engine Limits Chart


(Except Starting)), determine and correct thecause of overtemperature. If it was during a start,have the engine visually inspected through theexhaust duct (Figure 7-29), then record the actionin the engine logbook.

Overtemperature in area B will require that a hotsection inspection be performed. During a hotsection inspection, the components forward ofthe combustion chamber are examined andreplaced. Parts may be repaired or replaced asnecessary. In area C overtemperatures mayrequire that the engine be returned for overhaul.Exceeding ITT limits in this area for more than afew seconds may cause extensive engine damage.

STARTER OPERATING TIME LIMITSThe engine starters are time-limited during thestarting cycle if for any reason multiple starts arerequired in quick sequence. The starter is limitedto 40 seconds ON then 60 seconds OFF for cool-ing before the next sequence of 40 seconds ON,60 seconds OFF. After the third cycle of 40 sec-onds ON, the starter must stay OFF for 30minutes. If these limits are not observed, over-heating may damage the starter.

Figure 7-26 Overtorque Limits Chart

Figure 7-27 Overtemperature Limits (Starting)


Trend MonitoringDuring normal operations, gas turbine enginesare capable of producing rated power forextended periods of time. Engine operatingparameters, such as output torque, interstage tur-bine temperature, compressor speed, and fuelflow for individual engines are predictable underspecific ambient conditions. On PT6A engines,these predictable characteristics may be takenadvantage of by establishing and recording indi-vidual engine performance parameters. Theseparameters can then be compared periodically topredicted values to provide day-to-day visualconfirmation of engine efficiency.

The Engine Condition Trend Monitoring System,recommended by Pratt and Whitney, is a process ofperiodically recording engine instrument read-ings

Figure 7-28 Overtemperature Limits (Except Starting)

Figure 7-29 View through Exhaust Duct


such as torque, interstage turbine temperature, com-pressor speed, and fuel flow, correcting the readingsfor altitude, outside air temperature, and airspeed, ifapplicable, and then comparing them to a set of typ-ical engine characteristics. Such comparisonsproduce a set of deviations in interstage turbinetemperature, compressor speed, and fuel flow.

DATA COLLECTION FORMThe trend monitoring procedure used specifiesthat flight data be recorded on each flight day,every five flight hours, or other flight period.Select a flight with long established cruise, per-ferably at a representative altitude and airspeed.With engine power established and stabilized fora minimum of five minutes, record the followingdata on a form similar to the in-flight engine datalog shown in (Figure 7-30):

Indicated airspeed (IAS) ....................... In knots

Outside air temperature (OAT) .................. In º C

Pressure altitude (ALT)............................. In feet

Propeller speed (NP) ................................. In rpm

Torque (TQ).................................. In foot-pounds

Gas generator speed (NG or N1 ) ......In %NG or N1

Interturbine temperature (ITT)................... In º C

Fuel Flow (FF) .......................................... In pph

PROPELLERS

GENERALThis section describes the propellers and theassociated system. Location and use of propellercontrols, principle of operation, reversing, adnfeathering are included in this discussion.

DATE OAT PRESS IAS PROP TORQUE N1 ITT FUEL DELTA* DELTA* DELTA* OIL OIL ELECT(°) ALT (KTS) SPEED (FT/LBS) (%) FLOW NG ITT FF TEMP PRESS LOAD

LEFT

RIGHT

LEFT

RIGHT

LEFT

RIGHT

LEFT

RIGHT

LEFT

RIGHT

LEFT

RIGHT

LEFT

RIGHT

LEFT

RIGHT

Figure 7-30 In-Flight Engine Data Log


PROPELLER SYSTEMThis section on the operation and testing of thepropeller system on the Beechcraft King AirC90A and C90B is directed at increasing thepilot’s understanding of the theory of operationof a constant-speed, full-feathering, reversingpropeller, and helping him better understand thepropeller system checks conducted as outlined inthe Before Takeoff (Runup) checklist in thePilot’s Operating Handbook.

Each engine is equipped with a conventionalthree-blade (C90A) or four-blade (C90B),full-feathering, constant-speed, counter-weighted, reversing, variable-pitch propellermounted on the output shaft of the reductiongearbox (Figure 7-31).

The propeller pitch is controlled by engine oilpressure boosted through a governor pump inte-gral within the propeller governor. Centrifugalcounterweights and feathering springs move thepropeller blades toward high pitch and into thefeathered position. Without oil pressure to coun-teract the counterweights and feathering springs,the propeller blades would move into feather. Anoil pump, which is part of the propeller governor,boosts engine oil pressure to move the propellerto low pitch and reverse. The propeller feathersafter engine shutdown.

Propeller tiedown boots (Figure 7-32) are pro-vided to prevent windmilling at zero oil pressurewhen the airplane is parked.

Low pitch propeller position is determined by theprimary low pitch stop, which is a mechanicallyactuated hydraulic stop. Beta and reverse bladeangles are controlled by the power levers in theBeta and reverse range.

Two governors, a primary governor and a backupoverspeed governor, control the propeller rpm.The propeller control lever adjusts the governor’ssetting (1,800 to 2,200 rpm). The overspeed gov-ernor will limit the propeller to 2,288 rpm shouldthe primary governor malfunction. However, ifthe propeller exceeds 6% above the selected rpmof the primary governor, usually the fuel toppinggovernor will limit the rpm by reducing engine

power. In the reverse range, the fuel topping gov-ernor is reset to limit the propeller rpm to 95% ofselected rpm.

MCCAULEY AND HARTZELL FOUR-BLADE PROPELLERSThe C90B is equipped with Hartzell on LJ-1542and after (McCauley on C90B prior to LJ-1542),90-inch- diameter, four-blade, full-reversing,dynamically balanced propellers. The mainadvantages of the four-blade propellers are thatthey have lower tip speeds (and thus generate lessnoise), create less airframe vibration, and providegenerous propeller tip-to-ground clearance.Dynamic vibration absorbers mounted inside thecockpit and cabin (a total of 26 absorbers) areused in conjunction with the four-blade propel-lers to reduce noise and vibration even more.

BLADE ANGLEBlade angle is the angle between the chord of thepropeller and the propeller’s plane of rotation.Blade angle is different near the hub than it isnear the tip, due to the normal twist which isincorporated in a blade to increase its efficiency.In the propellers used on the C90A and C90BKing Air, the blade angle is measured at thechord 30 inches out from the propeller’s center.This position is referred to as the “30-inch sta-tion.” All blade angles given in this section areapproximate (Figure 7-34).

PRIMARY GOVERNORThe primary governor (Figure 7-33) is needed toconvert a variable-pitch propeller into a constant-speed propeller. It does this by changing bladeangle to maintain the propeller speed the operatorhas selected. The primary governor can maintainany selected propeller speed from approximately1,800 rpm to 2,200 rpm.

Suppose an airplane is in normal cruising flightwith the propeller turning 1,900 rpm. If the pilottrims the airplane down into a descent withoutchanging power, the airspeed will increase. Thisdecreases the angle of attack of the propeller


Figure 71-31 Propellers


blades, causing less drag on the propeller, thusbeginning to increase its rpm. Since this propellerhas a variable-pitch capabilities and is equippedwith a governor set at 1,900 rpm, the governorwill sense this “overspeed” condition andincreases blade angle to a higher pitch. Thehigher pitch increases the blade’s angle of attack,slowing it back to 1,900 rpm, or “onspeed.”

Likewise, if the airplane moves from cruise toclimb airspeeds without a power change, the pro-peller rpm tends to decrease, but the governorresponds to this “underspeed” condition bydecreasing blade angle to a lower pitch, and therpm returns to its original value. Thus the gover-nor gives “constant-speed” characteristics to thevariable-pitch propeller.

Power changes, as well as airspeed changes,cause the propeller to momentarily experi-ence overspeed or underspeed conditions, butagain the governor reacts to maintain theonspeed condition.

There are times, however, when the primary gov-ernor is incapable of maintaining selected rpm.For example, imagine an airplane approaching toland with its governor set at 1,900 rpm. As powerand airspeed are both reduced, underspeed condi-tions exist which cause the governor to decrease

Figure 7-33 Primary Governor Diagram

Figure 7-32 Propeller Tiedown Boot

Installed


Figure 7-34 Blade Angle Diagram


blade angle to restore the onspeed condition. Ifblade angle could decrease all the way, to 0º orreverse, the propeller would create so much dragon the airplane that the aircraft control would bedramatically reduced. The propeller, acting as alarge disc, would blank the airflow around the tailsurfaces, and a rapid nosedown pitch changewould result.

To prevent these unwanted aerobatics, somedevice must be provided to stop the governorfrom selecting blade angles that are too low forsafety. As the blade angle is decreased by thegovernor, eventually the low pitch stop isreached, and now the blade angle becomes fixedand cannot continue to a lower pitch. The gover-nor is therefore incapable of restoring theonspeed condition, and propeller rpm falls belowthe selected governor rpm setting.

PRIMARY GOVERNOR OPERATIONThe propeller levers adjust the primary propellergovernor between 1,800 rpm and 2, 200 rpm. Theprimary propeller governor, mounted at the top ofthe engine reduction gearbox, has two functions:it can select any constant propeller rpm withinthe range of 1,800 to 2,200, and it can alsofeather the propeller. The primary propeller gov-ernor adjusts propeller rpm by controlling the oilsupply to the propeller dome.

An integral part of the primary propeller gover-nor is the governor pump. This pump is driven bythe N2 shaft and raises the engine oil pressurefrom normal to approximately 375 psi. Thegreater the oil pressure sent to the propellerdome, the lower the propeller pitch. The oil pres-sure is always trying to maintain a low pitch;however, the feathering springs and centrifugalcounterweights are trying to send the propellerinto the feathered position. Propeller control is abalancing act of opposing forces. A transfergland is located on the propeller shaft. This trans-fer gland allows the oil to enter and exit thepropeller dome area. Thus, the transfer gland isalways replenishing the oils supply to the propel-ler pitch mechanism with fresh warm oil.

The primary propeller governor uses a set of rotat-ing flyweights that are geared to the propeller

shaft. The flyweights act as a comparison to adesired reference speed of how fast the propeller isturning. These flyweights are connected to a free-floating pilot valve. The slower the flyweights areturning in relation to the desired reference speed,the lower the position of the pilot valve. If the pro-peller and the flyweights turn faster, the additionalcentrifugal force makes the pilot valve rise insidethe governor. The pilot valve position determineshow much oil pressure is being sent to the propel-ler pitch mechanism. Here are a few examples.

If a propeller rpm of 1,900 is selected and thepropeller is actually turning at 1,900, the fly-weights are in their center or “onspeed”condition (Figure 7-34). The pilot valve is in themiddle position. This maintains a constant oilpressure to the propeller pitch mechanism, whichcreates a constant pitch and a constant rpm.

If the airplane enters a descent, without anychange to the cockpit controls, there will be atendency for the airspeed to increase and thepropeller to turn faster (Figure 7-36). The fly-weights will, in turn, rotate faster. The additionalcentrifugal force will make the pilot valve rise.Notice that oil can now escape via the pilotvalve. Lower oil pressure will result in a higherpitch and a reduction of propeller rpm. The pro-peller will then return to its original rpm setting.The flyweights will then slow down, and thepilot valve will return to the equilibrium positionto maintain the selected propeller rpm.

If the airplane enters a climb without any changein the cockpit controls, the airspeed will decreaseand the propeller will tend to slow (Figure 7-37).The flyweights in the propeller governor willslow down, because of a loss in centrifugal force,and the pilot valve will lower. This will allowmore oil pressure to the propeller pitch mecha-nism. High oil pressure will result in a lowerpitch. This in turn will cause an increase in pro-peller rpm. The propeller will increase to itsoriginal rpm setting, the flyweights will thenspeed up, and the pilot valve will return to itsequilibrium or “onspeed” position, such astorque, interstage turbine temperature, compres-sor speed, and fuel flow, correcting the heldconstant by changing the propeller blade angles.The cockpit propeller lever adjusts where the


Figure 7-35 Propeller Onspeed Diagram

Figure 7-36 Propeller Overspeed Diagram


equilibrium or “onspeed” condition will occur.The pilot can select any constant propeller rpmfrom 1,800 to 2,200 rpm. Normally 2,200 is usedfor takeoff and 2,000 rpm for climb. Cruise rpmis 1,900 rpm.

LOW PITCH STOPIt is easy for the pilot to determine when the pro-peller blade angle is at the low pitch stop.Assuming the propeller is not feathered or in theprocess of being feathered, whenever the propel-ler rpm is below the selected governor rpm, thepropeller blade angle is at the low pitch stop.

This assumes that momentary periods of under-speed are not being considered. Rather, thepropeller rpm is below and staying below theselected governor rpm.

For example, if the propeller control is set at1,900 rpm but the propeller is turning at lessthan 1,900 rpm, the blade angle is at the lowpitch stop.

On many types of airplanes, the low pitch stopis simply at the low pitch limit of travel, deter-mined by the propeller’s construction. But witha reversing propeller, the extreme travel in thelow pitch direction is past 0º, into reverse ornegative blade angles (Figure 7-38). Conse-quently, the low pitch stop on this propellermust be designed in such a way that it can berepositioned when reversing is desired.

The low pitch stop is created by mechanical link-age sensing the blade angle. The linkage causes avalve to close, which stops the flow of oil pres-sure coming into the propeller dome. Since thispressure causes low pitch and reversing, once itis blocked, a low pitch stop has been created. Thelow pitch stop is commonly referred to as the“Beta” valve. Furthermore, the valve is spring-loaded to cause the propeller to feather in theevent of mechanical loss of Beta valve control.

The position of the low pitch stop is controlledfrom the cockpit by the power lever. Wheneverthe power lever is at IDLE or above, this stop is

Figure 7-37 Propeller Underspeed Diagram


set at approximately 15º for the C90A or approx-imately 12º for the C90B. But bringing the powerlever aft of IDLE progressively repositions thestop to lesser blade angles.

Before reversing can take place, the propellermust be on the low pitch stop. As the propellersreach approximately 15º for the C90A or approx-imately 12º for the C90B, the Beta valve isrepositioned, creating the low pitch stop. The pri-mary governor is sensing an underspeed and isdirecting oil pressure into the propeller dome.The Beta valve is controlling oil flow into the pri-mary governor, and is defining the low pitch stopthrough oil pressure.

When blade angles less than approximately 15ºfor the C90A or approximately 12º for the C90Bare requested, the linkage pulls the Beta valveactuator, readjusting the propeller blade angle asthe Beta valve allows more oil into the propellerdome. The slip ring moves with the prop domeand will define the low pitch stop at a lower, ornegative, blade angle. If blade angles less than

approximately 15º for the C90A or approxi-mately 12º for the C90B are requested before thepropeller blades are on the low pitch stop, the slipring will not move, and the reversing cable andlinkage may be damaged.

The region from 15º to –11º (C90A) or 12º to–10º (C90B) blade angle is referred to as the Betarange. On the C90A, the range from 15º to –5º,the engine’s compressor speed (N1) remains atthe value it had when the power lever was atIDLE (low idle to high idle) based on conditionlever position. From –5º to –11º blade angle, theN1 speed progressively increases to a maximumvalue at –11º blade angle of approximately 85%_+3%. This region, designated by red and whitestripes on the power lever gate, is referred to asthe “Beta Plus Power” range or Reverse, andends at maximum reverse.

On the C90B, the Ground Fine range extendsfrom +12º to +3º, and the engine’s compressorspeed (N1) remains at the value it had when thepower lever was at IDLE (low idle to high idle)

Figure 7-38 Low Pitch Stop Diagram


based on condition lever position. From +3º to–10º blade angle, the N1 speed progressivelyincreases to a maximum value at –10º bladeangle of approximately +85% _+3%.

Low Pitch Stop OperationDuring non-reversing operations, the low pitchstop prevents the propeller blades from reducingthe airflow over the empennage of the aircraft.

The low pitch stop uses a mechanical linkage tohydraulically control propeller blade angle. Asthe propeller blades reduce angle throughapproximately 20º of pitch, the flange mountedon the propeller dome contacts the nuts locatedon the rods mounted on the slip ring. The propel-ler dome moves the slip ring forward, which inturn activates the Beta valve, which controls oilpressure into the propeller dome.

Riding in the slip ring is linkage which connectsthe Beta valve with the slip ring, and the powerlevers via a cable. As the slip ring moves, the link-age pivots about the end with the cable attached toit, with the Beta valve in the middle. For reversing,the pilot repositions the linkage with the powerlevers, which resets the low pitch stop.

When the Beta valve is controlling blade angle,oil pressure supplied from the governor oil pumpis supplying pressure through the Beta valve tothe propeller dome. The Beta valve modulatesthe amount of pressure entering the propellerdome, controlling the blade angle. The primarygovernor must be in the underspeed condition,allowing all of the pressure flowing from the Betavalve into the propeller dome. If the underspeedcondition did not exist when lower blade anglesare requested, the Beta valve could not fully con-trol the propeller blade angle, and the slip ringwould not move without help from the propellerblades. Since the propeller blades only contactthe slip ring when the blades are at the low pitchstop, the request for lower blade angles when thepropellers are not on the low pitch stop will resultin damage to the control cable, as it cannot effectthese changes alone.

BETA AND REVERSE CONTROLThe geometry of the power lever linkage throughthe cam box is such that power lever incrementsfrom idle to full forward thrust have no effect onthe position of the Beta valve. When the powerlever is moved from idle into the reverse range, itpositions the Beta valve to direct governor oilpressure to the propeller piston, decreasing bladeangle through zero into a negative range. Thetravel of the propeller servo piston is fed back tothe Beta valve to null its position and, in effect,provide infinite negative blade angles all the wayto maximum 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 controlfor ground handling.

Beta and Reverse Control OperationWhen the blade angle reaches approximately 20º,the flange extending from the dome makes con-tact with the Beta nuts (Figure 7-39). As thepropeller pitch angle continues to decrease, eachflange on the propeller dome pushes the nut andthe attached Beta rod forward. As the rod movesforward, it pulls the slip ring forward. In turn, aBeta valve inside the governor is pulled into theoil pressure cutoff position. The linkage is set tocontrol the oil pressure supply to the dome whenthe blade angle reaches low pitch stop.

If this system were fixed at the low pitch stop, thepropeller could not be reset throughout the Betarange. However, the low pitch stop can beadjusted to allow access to the Beta and reverserange on the ground. The hydraulic low pitchstop can be reset to allow the propeller to operatein the Beta and reverse ranges while the aircraftis on the ground and the engines are operating.

When the power levers are lifted up and over theidle detent into the Beta range, the Beta valve isrepositioned. As the Beta arm moves back, the


Figure 7-39 Beta Range and Reverse Diagram


Beta valve is opened, re-establishing oil flow tothe propeller dome. This allows the propellerblade to move to a flatter pitch. As the propellerblades move to a flatter pitch, the propellerdome and slip ring continue forward, eventuallymoving the Beta valve back into position to stoppropeller blades. In summary, the position ofthe low pitch stop is controlled by the powerlevers. When the power levers are set at idle orabove, the stop is set at approximately 15º onthe C90A or approximately 12º on the C90B.When the power levers are moved aft of idle,however, the low pitch stop is repositioned tolesser blade angles.

The propeller can be feathered by moving thepropeller lever full aft past the detent into thefeather range. The feathering action raises thepilot valve to the full up position. The oil pres-sure is released from the propeller pitchmechanism and the propeller feathers. In thistype of turbine engine, the propeller shaft and N1shaft are not connected. Thus, the propeller canbe feathered with the engine running at idlepower. Without an autofeather system, in flight,the propeller will maintain rpm unless it is manu-ally feathered when the engine is shut down.

There are situations where the propeller primarygovernor cannot maintain the selected propellerrpm, such as final approach where power and air-speed are being reduced. With the progressivereduction of power and airspeed on final, the pro-peller and rotating counterweights will tend to goto the underspeed condition. In the underspeedcondition the pilot valve will open, increasing oilpressure to the dome, and the propeller pitch willdecrease as power and airspeed are reduced.Since the reversible propeller is capable ofdecreasing past 0º into negative or reverse bladeangles, the low pitch stop prevents the bladeangle from decreasing beyond a predeterminedvalue. When the propeller governor becomesincapable of maintaining the onspeed condition,the propeller rpm will fall below the selectedgovernor rpm setting.

Assuming the propeller is not feathered, when-ever the propeller rpm is below the selectedgovernor setting, the propeller blade angle is atthe low pitch stop. The low pitch stop mechanismis created by linkage that references the actualblade angle.

Moving the power lever within the Beta range onthe C90A or the ground fire range on the C90Badjusts propeller pitch. Moving the power leverswithin the reverse range adjusts propeller pitchand N1, up to the maximum N1 in reverse of88%. Attempting to pull the power levers inreverse with the propellers in feather will causedamage to the reversing linkage of the powerlever. Also, pulling the power levers into thereverse position on the ground with the enginesshut down will damage the reversing system.

OVERSPEED GOVERNORThe overspeed governor provides protectionagainst excessive propeller speed in the event ofprimary governor malfunction. Since the PT6’spropeller is driven by a free turbine (independentof the engine’s), overspeed could occur if the pri-mary governor were to fail.

The operating point of the overspeed governor isset at 2,288 rpm. If an overspeeding propeller’sspeed reached 2,288 rpm, the overspeed governorwould control the oil pressure and pitch to pre-vent the rpm from continuing its rise. From apilot’s point of view, a propeller tachometer sta-bilized at approximately 2,288 would indicatefailure of the primary governor and proper opera-tion of the overspeed governor. The overspeedgovernor can be reset to approximately 2,000rpm for test purposes.

Overspeed Governor OperationIf the primary propeller governor failed, an over-speed condition could occur. However, severalsafety devices in the systems come into play in theevent of a primary governor failure. A hydraulic


overspeed governor (Figure 7-40) is located on theleft side of the propeller reduction gearbox. It hasa set of flyweights and a pilot valve similar tothose of the primary governor. If a runaway pro-peller’s speed were to reach 2,288 rpm, theoverspeed governor flyweights would make itspilot valve rise. This would decrease the oil pres-sure at the propeller dome. The blade angle wouldincrease as necessary to prevent the rpm from con-tinuing its rise. Testing of the overspeed governorat approximately 2,000 rpm is accomplished dur-ing runup by using the propeller governor testswitch on the pilot’s left subpanel.

FUEL TOPPING GOVERNORThe fuel topping governor can also control anoverspeed condition and is set at 6% above theprimary governor’s selected speed. In an over-speed condition, the fuel topping governor willlimit propeller rpm by decreasing pneumaticpressure to the fuel control unit, reducing fuelflow and engine speed as means of controlling

propeller rpm. In reverse, the fuel topping gover-nor is reset to 95% of selected rpm to insure thatthe propeller will not reach the selected rpm. Thefuel topping governor will only prevent an over-speed if the primary governor’s flyweight’s arestill operational.

POWER LEVERSThe power levers (Figure 7-41) are located on thepower lever quadrant (first two levers on the leftside) on the center pedestal. They are mechani-cally interconnected through a cam box to the fuelcontrol unit, the Beta valve and follow-up mech-anism, and the fuel topping (NP) governor. Thepower lever quadrant permits movement of thepower lever from idle to maximum thrust and inthe Beta/reverse range from idle to maximumreverse. A gate in the power lever quadrant at theIDLE position prevents inadvertent movement ofthe lever into the Beta/reverse range. The pilotmust lift the power levers up and over this gateto select Beta or reverse. On the C90B, there

Figure 7-40 Overspeed Governor Diagram


is a second gate labeled “GROUND FINE,” todistinguish between GROUND FINE andREVERSE.

The function of the power levers is to establish agas generator rpm through the gas generator gov-

ernor (NG) and a fuel flow that will produce andmaintain the selected N1 rpm. In the Beta orGROUND FINE range, the power levers are usedto change the propeller blade angle, thus chang-ing propeller thrust.

In the REVERSE range, the power lever:

● Selects a blade angle proportionate to theaft travel of the lever

● Selects an N1 that will sustain theselected reverse power

● Resets the fuel topping governor from itsnormal setting of 106% to approximately95% of the primary governor setting

Propeller Control LeversPropeller rpm, within the primary governor rangeof 1,800 to 2,200 rpm, is set by the position of thepropeller control levers (Figure 7-42). Theselevers, one for each propeller, are located betweenthe power levers and the condition levers on thecenter pedestal quadrant. The full forward positionsets the primary governor at 2,200 rpm. In the fullaft position at the feathering detent, the primarygovernor is set at 1,750 rpm. Intermediate propel-ler rpm positions can be selected by moving the

Figure 7-41 Power Levers

Figure 7-42 Propeller Control Levers


propeller levers to the corresponding position, toselect the desired rpm as indicated on the propellertachometer. These tachometers read directly inrevolutions per minute.

A detent at the low rpm position prevents inad-vertent movement of the propeller lever into thefeather position, indicated by the red and whitestripes across the lever slots in the quadrant. Atthe full feather position, the levers position thegovernor pilot valve to dump oil pressure fromthe propeller hub, and allow the counterweightsand springs to position the propeller blades to thefeather position.

AUTOFEATHER SYSTEMThe automatic feathering system provides ameans of immediately dumping oil pressure fromthe propeller hub, thus enabling the featheringspring and counterweights to start the featheringaction of the blades in the event of an engine fail-ure (Figure 7-43). Although the system is armedby a swi tch on the subpanel , p lacarded“AUTOFEATHER” and “ARM–OFF–TEST,” thecompletion of the arming phase occurs whenboth power levers are advanced above 90% N1, atwhich time both the right and left indicator lightson the annunciator panel indicate a fully armedsystem (Figure 7-44). The annunciator panellights are green, placarded “L AUTOFEATHER”and “R AUTOFEATHER.” The system willremain inoperative as long as either power leveris retarded below 90% N1 position. The system isdesigned for use only during takeoff, climb, andmissed approach and should be turned off whenestablishing cruise. With the system armed, iftorquemeter oil pressure on either engine dropsbelow a prescribed setting, the oil is dumpedfrom the servo, the feathering spring starts theblades toward feather, and the autofeather systemof the other engine is disarmed. Disarming of theautofeather portion of the operative engine is fur-ther indicated when the annunciator indicatorlight for that engine extinguishes. AutofeatherSystem Test

The autofeather test is accomplished with thepower below 90% N1. Therefore, the autofeatherswitch must be held to TEST so that the power

lever switches are bypassed to complete theautofeather circuit (Figure 7-45).

PROPELLER SYNCHROPHASER SYSTEMA Type II synchrophaser system is installed inthe King Air C90A and C90B. The propeller syn-chrophaser automatically matches the rpm of thetwo propellers and maintains the blades of onepropeller at a predetermined relative positionwith the blades of the other propeller. The pur-pose of the system is to reduce propeller beat andcabin noise from unsynchronized propellers.

Synchrophaser OperationThe Type II synchrophaser system (Figure 7-46)is an electronic system, certificated for takeoffand landing. It is not a master-slave system, andit functions to match the rpm of both propellersand establish a blade phase relationship betweenthe left and right propellers to reduce cabin noiseto a minimum.

The system cannot reduce rpm of either propellerbelow the datum selected by the propeller controllever. Therefore, there is no indicating annuncia-tor light associated with the Type II system.

To prevent either propeller from losing excessiverpm if the other propeller is feathered while thesynchrophaser is on, the synchrophaser has a lim-ited range of authority from the manual governorsetting. In no case will the rpm fall below thatselected by the propeller control lever. Normalgovernor operation is unchanged, but the syn-chrophaser will continuously monitor propellerrpm and reset either governor as required. Propel-ler rpm and position is sensed by a magnetic pick-up mounted adjacent to each propeller spinnerbulkhead. This magnetic pick-up will transmitelectrical pulses once per revolution to a controlbox installed forward of the pedestal.

The control box converts any pulse rate differ-ences into correction commands, which aretransmitted to coils mounted close to the fly-weights of each primary governor. By varying


Figure 7-43 Autofeather System Diagram—Left Engine Failed and Feathering


Figure 7-44 Autofeather System Diagram—Armed


Figure 7-45 Autofeather Test Diagram


the coil voltage, the governor speed settings arebiased until the prop rpm’s exactly match. A tog-gle switch installed adjacent to the synchroscopeturns the system on. In the synchrophaser OFFposition, the governors operate at the manualspeed settings selected by the pilot. To operatethe synchrophaser system, synchronize the pro-pellers manually or establish a maximum of 20rpm difference between the engines, then turn thesynchrophaser on. The system may be on fortakeoff and landing.

To change rpm with the system on, adjust bothpropeller controls at the same time. If the syn-chrophaser is on but does not adjust the proprpm to match, the system has reached the end ofits range. Increasing the setting of the slowprop, or reducing the setting of the fast prop,

will bring the speeds within the limited syn-chrophaser range. I f preferred, turn thesynchrophaser switch off, resynchronize manu-ally, and turn the synchrophaser on.

Propeller SynchroscopeA propeller synchroscope is located to the left ofthe oil pressure/temperature indicators and givesthe status of propeller synchronization. The faceof the synchroscope has a black and white crosspattern which can spin either left or right. If theright propeller rpm is greater than the left, theface turns clockwise or right. With the left pro-peller rpm greater than the right, the face turnscounterclockwise or left. No rotation of the faceindicates that both propellers are synchronized.

Figure 7-46 Propeller Synchrophaser


8-i

CHAPTER 8

CONTENTS

Page

FIRE PROTECTION

INTRODUCTION ..................................................................................................................

8-1

GENERAL ..............................................................................................................................

8-1

Fire Detection System .....................................................................................................

8-1

Fire Extinguishing System ..............................................................................................

8-2

8-ii



8-iii

Figure Title Page

ILLUSTRATIONS

8-1

Fire Detection System..............................................................................................

8-3

8-2

Fire Extinguishing System.......................................................................................

8-4

8-3

Fire Extinguisher Cylinder Pressure Gage ..............................................................

8-5


8-1

CHAPTER 8FIRE PROTECTION

INTRODUCTION

The aircraft fire protection system consists of engine fire detection and fire extinguishingsystems. Cockpit controls and indicators monitor and operate the system.

GENERAL

The fire protection chapter of the training manualpresents a description and discussion of the air-plane fire protection system and components.The location and purpose of switches and indica-tors are described.

FIRE DETECTION SYSTEM

The fire detection system (Figure 8-1) isdesigned to provide immediate warning in theevent of fire in either engine compartment.

The detection system is operable whenever thegenerator buses are active.

The system consists of the following: threephotoconductive cells for each engine; a con-trol amplifier for each engine; two red warninglights on the warning annunciator panel, oneplacarded L ENG FIRE, the other R ENGFIRE; a test switch on the copilot’s left sub-panel; and a circuit breaker placarded FIREDET on the right side panel.

��

��

��

��

��

��

��

��

��

��

FIRE PULL

FIREWARN

8-2


The six photoconductive-cell flame detectors aresensitive to infrared radiation. They are posi-tioned in each engine compartment so as toreceive both direct and reflected infrared rays,thus monitoring the entire compartment withonly three photocells. Temperature level and rateof temperature rise are not controlling factors inthe sensing method.

Conductivity through the photocell varies indirect proportion to the intensity of the infraredradiation striking the cell. As conductivityincreases, the amount of current from the electri-cal system flowing through the flame detectorincreases proportionally. To prevent stray lightrays from signaling a false alarm, a relay in thecontrol amplifier closes only when the signalstrength reaches a preset alarm level. When therelay closes, the appropriate left or right warningannunciators illuminate. When the fire has beenextinguished, the cell output voltage drops belowthe alarm level and the relay in the control ampli-fier opens. No manual resetting is required toreactivate the fire detection system.

Fire Detection Test System

The rotary switch on the copilot’s left subpanel,placarded TEST SWITCH-FIRE DET, has fourpositions: OFF - 3 - 2 - 1. (If the optional enginefire extinguishing system is installed, the switchis placarded TEST SWITCH - FIRE DET &FIRE EXT and the left side of the test switch willinclude LEFT - EXT - RIGHT positions.)

The three test positions for the fire detector sys-tem are located on the right side of the switch.When the switch is rotated from OFF (down) toany one of these three positions, the output volt-age of a corresponding flame detector in eachengine compartment is increased to a level suffi-cient to signal the amplifier that a fire is present.

The following should illuminate as the selector isrotated through each of the three positions: theFAULT WARNING flasher, the L ENG FIRE andR ENG FIRE warning annunciators and, if theoptional engine fire extinguishing System isinstalled, the red lenses placarded L ENG FIREEXT - PUSH and R ENG FIRE EXT -PUSH on

the fire-extinguisher activation switches. The sys-tem may be tested anytime, either on the groundor in flight. The TEST SWITCH should beplaced in all three positions, in order to verifythat the circuitry for all six fire detectors is func-tional. Illumination failure of all the firedetection system annunciators when the TESTSWITCH is in any one of the three flame-detec-tor-test positions indicates a malfunction in oneor both of the two detector circuits (one in eachengine) being tested by that particular position ofthe TEST SWITCH.

FIRE EXTINGUISHING SYSTEM

The optional engine fire extinguishing system(Figure 8-2) incorporates an explosive cartridgeinside the extinguisher of each engine. Eachengine has its own self-contained extinguishingsystem, which can be used only once betweenrechargings. This system cannot be crossfed.When the activation valve is opened, the pressur-ized extinguishing agent is discharged through aplumbing network which terminates in strategi-cally located spray nozzles.

The fire extinguisher control switches used toactivate the system are located below theglareshield annunciator panel on the C90A. Thecontrol switches are on either side of the annun-ciator panel on the C90B. Their power is derivedfrom the hot battery bus. The detection system isoperable whenever the generator buses are active.But the extinguishing system can be dischargedat any time, since it is operated from the hot bat-tery bus. Therefore, even though the airplanemay be parked with the engines off, the fireextinguishing system may be discharged.

Each push-to-actuate switch incorporates threeindicator lenses. The red lens, placarded L (or) RENG FIRE EXT - PUSH, warns of the presenceof fire in the engine. The amber lens, placardedD, indicates that the system has been dischargedand the supply cylinder is empty. The green lens,placarded OK, is provided only for the preflighttest function.

To discharge the cartridge, raise the break-awaywired clear plastic cover and press the face of the


8-3

Figure 8-1 Fire Detection System

8-4


Figure 8-2 Fire Extinguishing System


8-5

lens. This is a one-shot system and will be com-pletely expended upon activation. The amber Dlight will illuminate and remain illuminated,regardless of battery switch position, until thepyrotechnic cartridge has been replaced.

Fire Extinguisher Test System

The fire extinguisher system test functions, incor-porated in the rotary TEST SWITCH - FIREDET & FIRE EXT, test the circuitry of the fireextinguisher system. During preflight, the pilotshould rotate the TEST SWITCH to each of thetwo positions (RIGHT EXT and LEFT EXT) andverify the illumination of the amber D light andthe green OK light on each fire-extinguisher acti-vation switch below the glareshield. Illuminationduring this check indicates that the bottle chargedetector circuitry and squib firing circuitry areoperational and that the squib is in place.

A gage, (Figure 8-3) calibrated in psi, is providedon each supply cylinder for determining the levelof charge. The gages should be checked duringpreflight. The cylinder and gages are located inthe main wheel wells.

Figure 8-3 Fire Extinguisher CylinderPressure Gage


9-i

CHAPTER 9

CONTENTS

Page

PNEUMATICS

INTRODUCTION ..................................................................................................................

9-1

DESCRIPTION.......................................................................................................................

9-1

ENGINE BLEED AIR PNEUMATIC SYSTEM...................................................................

9-3

Pneumatic Air Source......................................................................................................

9-3

Vacuum Air Source .........................................................................................................

9-3

Cabin Door Seal ..............................................................................................................

9-4

SURFACE DEICE SYSTEM .................................................................................................

9-4


9-iii

Figure Title Page

ILLUSTRATIONS

9-1

Pneumatic System Diagram.....................................................................................

9-2

9-2

Pneumatic Pressure Gage ........................................................................................

9-3

9-3

Gyro Suction Gage ..................................................................................................

9-4

9-4

Surface Deice Boot Installation ...............................................................................

9-4

9-5

Surface Deice System Diagram ...............................................................................

9-5

9-6

Surface Deice Controls ............................................................................................

9-6


9-1

CHAPTER 9PNEUMATICS

INTRODUCTION

The pneumatic and vacuum systems are necessary for the operation of surface deicers, instru-ment air, production of vacuum, rudder boost, flight hourmeter, cabin door seal, pressurizationcontroller, and pressurization outflow and safety valves. Pilots need to know how the bleed air isdistributed and controlled for these various uses. This section identifies these systems andcovers the pneumatic manifold and controls in detail.

DESCRIPTION

The Pneumatic and Vacuum Systems section ofthe training manual presents a description anddiscussion of pneumatic and vacuum systems.

The sources for pneumatic air, vacuum, andacceptable gage readings are discussed.

VALVE

L R

COBLEED AIR

515

20

AIR

9-2


PRESSURESWITCH

PNEUMATIC PRESSUREGAUGE (IN COCKPIT)

RIGHTSQUATSWITCH

OPEN INFLIGHT(N.C.)

TO DEICEBOOTS

EJECTOR

LEFT SQUATSWITCH

LANDING GEARHYDRAULIC FILL CAN

CLOSED ONGROUND (N.O.)

L SERVO

R SERVO

VALVE

L N.C.

R N.C.CHECK VALVE CHECK VALVE

AIRSTAIRDOOR SEAL

LINE

EMERGENCYEXIT SEAL

LINE

50 PSID

P SWITCH

VACUUMREGULATOR

GYRO SUCTION(IN COCKPIT)

GYROINSTRUMENTS

PRESSURIZATIONCONTROLLER,

OUTFLOW & SAFETYVALVES

HIGH PRESSURE BLEED AIR

REGULATED BLEED AIR

VACUUM

4 PSIPRESSURE

REGULATOR

18 PSIPRESSURE

REGULATOR

RUDDER BOOSTSYSTEM

13 PSIREGULATOR

GYROSUCTION

INCHES OF MERCURY

3 65430K

15K

PNEUMATICPRESSURE

PSI

0

10

20

DEICEDISTRIBUTER

VALVEFLIGHTHOURS 1/00

0 0 0 0 0

Figure 9-1 Pneumatic System Diagram


9-3

ENGINE BLEED AIR PNEUMATIC SYSTEM

The pneumatic system in Beechcraft King Airsprovides support for several operations on theairplane. These operations include surfacedeice, rudder boost, escape hatch seal, and thedoor seal. Pneumatic pressure is used to create avacuum source for the air-driven gyros, pressur-ization control, and deflation of the deice boots.

High-pressure bleed air from each engine com-pressor section, regulated at 18 psi, suppliespressure for the surface deice system, rudderboost, escape hatch and door seals, and vacuumsource (Figure 9-1). Vacuum for the flightinstruments is derived from a bleed-air ejector.One engine can supply sufficient bleed air forall these systems.

During single-engine operation, a check valvein the bleed air line from each engine preventsflow back through the line on the side of theinoperative engine. A suction gage calibrated ininches of mercury, on the copilot’s subpanel,indicates instrument vacuum (GYRO SUC-TION). To the right of the suction gage is aPNEUMATIC PRESSURE gage, calibrated inpounds per square inch, which indicates the airpressure available.

PNEUMATIC AIR SOURCE

Bleed air at a maximum rate of 90 to 120 psipressure is obtained from both engines, and flowsthrough pneumatic lines to a common manifoldin the fuselage. Check valves prevent reverseflow during single engine operation.

Downstream from the manifold, the bleed airpasses through an 18 psi regulator which incor-porates a relief valve set to operate at 21 psi incase of regulator failure. This regulated bleedair is used to supply pneumatic pressure toinflate the surface deicers, escape hatch anddoor seals, and to provide flow and pressure forthe vacuum ejector.

Bleed air is extracted from the P

3

tap of the engineat a temperature of approximately 450° F. It iscooled to approximately 70° above ambient tem-perature at the manifold in the fuselage due to heattransfer in the pneumatic plumbing.

Ordinarily, the pressure regulator valve, which isunder the right seat deck immediately forward ofthe main spar, will provide 18 +1 psi with theengine running at 70 to 80% N

1

. The PNEU-MATIC PRESSURE gage on the copilot’s rightsubpanel is provided to allow monitoring of thesystem pressure (Figure 9-2).

VACUUM AIR SOURCE

Vacuum is obtained from the vacuum ejector. Theejector is capable of supplying from 15 inches Hgvacuum at sea level, to 6 inches Hg vacuum at31,000 feet. The ejector supplies vacuum for thepressurization control system at a regulated 4.3 to5.9 inches Hg through a regulator valve.

The vacuum regulator is in the nose compartmenton the left side of the pressure bulkhead. Thevalve is protected by a foam filter.

With one engine running at 70 to 80% N

1

, thevacuum gage on the copilot’s right subpanel nor-mally should read approximately 5.9 +0/-0.2inches Hg.

PNEUMATICPRESSURE

0 20

10

PSI

Figure 9-2 Pneumatic Pressure Gage

9-4


The vacuum line for the instruments is routedthrough a suction relief valve that is designed toadmit into the system the amount of air requiredto maintain sufficient vacuum for proper opera-tion of the instruments. A GYRO SUCTIONgage (Figure 9-3), which is calibrated in inchesHg and is on the copilot’s right subpanel, indi-cates instrument vacuum.

CABIN DOOR SEAL

The entrance door to the cabin and the escapehatch uses air from the pneumatic system toinflate the seals after the airplane lifts off theground. Pneumatic air is tapped off the manifolddownstream of the 18 psi pressure regulator. Theregulated air then passes through a 4 psi regulatorand to the normally-open valve that is controlledby the left landing gear safety switch. When theairplane lifts off, the landing gear switch opensthe valve to the door and hatch seals, and theseals inflate.

SURFACE DEICE SYSTEM

The leading edges of the wings and horizontalstabilizer are protected against an accumulationof ice buildup (Figure 9-4). Inflatable bootsattached to these surfaces are inflated when nec-essary by pneumatic pressure to break away theice accumulation, and are deflated by vacuum.The vacuum is always supplied while the bootsare not in use and are held tightly against thewing. Vacuum pressure is overcome by pneu-matic pressure when the boots are inflated.

GYROSUCTION

INCHES OF MERCURY

3 64 535K

15K

Figure 9-3 Gyro Suction Gage

Figure 9-4 Surface Deice Boot Installation


9-5

Each wing has a leading-edge boot. The horizon-tal section of the tail has boots on the left andright segments of the horizontal stabilizer and onthe vertical stabilizer.

The su r face de i ce sy s t em r emoves i c eaccumulations from the leading edges of thewings and stabil izers. Ice is removed byalternately inflating and deflating the deice boots(Figure 9-5). Pressure-regulated bleed air from

Figure 9-5 Surface Deice System Diagram

9-6


the engines supplies pressure to inflate the boots.A venturi ejector, operated by bleed air, creates avacuum to deflate the boots and hold them downwhile not in use. To assure operation of thesystem in the event of failure of one engine, acheck valve is incorporated in the bleed-air linefrom each engine to prevent loss of pressurethrough the compressor of the inoperativeengine. Inflation and deflation phases arecontrolled by a distributor valve.

A three-position switch in the ICE PROTEC-TION group on the pilot’s subpanel, placardedSURFACE DEICE – SINGLE – OFF MANUAL,controls the deicing operation (Figure 9-6). Theswitch is spring-loaded to return to the OFF posi-tion from SINGLE or MANUAL. When theSINGLE position is selected, the distributorvalve opens to inflate the boots. After an inflationperiod of approximately seven seconds, an elec-

tronic timer switches the distributor to deflate theboots. When these boots have inflated anddeflated, the cycle is complete. On LJ1138 andafter, the wings and tail inflate separately. Thewings inflate for six seconds then the tail inflatesfor four seconds.

When the switch is held in the MANUAL posi-tion, all the boots will inflate simultaneously andremain inflated until the switch is released. Theswitch will return to the OFF position whenreleased. After the cycle, the boots will remain inthe vacuum hold-down condition until againactuated by the switch.

Electrical power to the boot system is requiredfor the control valve to inflate the boots in eithersingle-cycle or manual operation. With a loss ofthis power, the vacuum will hold them tightlyagainst the leading edge.

A single circuit breaker on the copilot’s side panel,receiving power from the center bus, supplies theelectrical operation of both boot systems. Shouldthe timer fail in the inflated position, the surfacedeice circuit breaker may be used as a manual con-trol. Pull the circuit breaker out to deflate theboots, and push in to inflate them. Treat the circuitbreaker as a manual control.

For most effective deicing operation, allow atleast 1/2 inch of ice to form before attempting iceremoval. Very thin ice may crack and cling to theboots instead of shedding. Subsequent cyclingsof the boots will then have a tendency to build upa shell of ice outside the contour of the leadingedge, thus making ice removal efforts ineffective.

STALLWARN

SURFACEDEICE

SINGLE PITOT

OFFMANUAL LEFT

OFF

10

DEICE WIPER

WEATHERSURF WSHLD

5

Figure 9-6 Surface Deice Controls


10-i

CHAPTER 10

CONTENTS

Page

ICE AND RAIN PROTECTION

INTRODUCTION ................................................................................................................

10-1

GENERAL ............................................................................................................................

10-1

ICE PROTECTION SYSTEMS ...........................................................................................

10-4

Description and Operation.............................................................................................

10-4

Pitot Heat .......................................................................................................................

10-5

Stall Warning Vane .......................................................................................................

10-5

Fuel System Anti-ice .....................................................................................................

10-6

Windshield Wipers ........................................................................................................

10-7

Windshield Anti-ice ......................................................................................................

10-7

Engine Air Inlet Lip Heat ............................................................................................

10-10

Engine Inertial Separators ...........................................................................................

10-11

Ice Vane Controls ........................................................................................................

10-12

Engine Autoignition System .......................................................................................

10-13

Propeller Electric Deice System..................................................................................

10-13

Wing Ice Lights ...........................................................................................................

10-14

Precautions During Icing Conditions ..........................................................................

10-15


10-iii

Figure Title Page

ILLUSTRATIONS

10-1

Ice and Rain Protection Required Equipment .......................................................

10-2

10-2

Ice and Rain Protection Controls ...........................................................................

10-3

10-3

Pitot Mast and Heat Controls.................................................................................

10-4

10-4

Stall Warning Vane and Heat Control ...................................................................

10-5

10-5

Fuel System Anti-ice .............................................................................................

10-6

10-6

Windshield Wipers ................................................................................................

10-7

10-7

Windshield Installation..........................................................................................

10-7

10-8

Windshield Anti-ice Diagram................................................................................

10-8

10-9

Windshield Anti-ice Switches ...............................................................................

10-8

10-10

Windshield Anti-ice Diagram—Normal Heat .......................................................

10-9

10-11

Windshield Anti-ice Diagram—High Heat .........................................................

10-10

10-12

Engine Air Inlet Lip Heat ....................................................................................

10-10

10-13

Inertial Separator in Retract Position...................................................................

10-11

10-14

Inertial Separator in Extend Position...................................................................

10-11

10-15

Ice Vane Controls ................................................................................................

10-12

10-16

Caution and Advisory Annunciators ...................................................................

10-12

10-17

Engine Autoignition Switches .............................................................................

10-13

10-18

Propeller Electric Deice System ..........................................................................

10-14

10-19

Wing Anti-ice Lights ...........................................................................................

10-14


10-1

CHAPTER 10ICE AND RAIN PROTECTION

INTRODUCTION

Flight in known icing conditions requires knowledge of conditions conducive to icing, and of allanti-ice and deice systems available to prevent excessive ice from forming on the airplane. Thissection identifies these systems with their controls and best usage.

GENERAL

This chapter presents a description and discus-sion of the airplane ice and rain protectionsystems. All of the anti-ice and deice systems inthis airplane are described, showing location,controls, and how they are used.

The purpose of this chapter is to acquaint thepilot with all the systems available for flight inicing or heavy rain conditions, along with theircontrols. Procedures in case of malfunction inany system are included. This also includes infor-mation concerning preflight deicing anddefrosting.

The Beechcraft King Air C90A and C90B areFAA-approved for flight in known icing condi-tions when the required equipment is installedand operational (Figure 10-1). The RequiredEquipment for Various Conditions of Flight List,contained in the “Limitations” section of the


, lists the necessaryequipment.

The ice and rain protection controls are groupedon the pilot’s and copilot’s subpanels, except thewindshield wiper control, which is overhead(Figure 10-2).

10-2


Figure 10-1 Ice and Rain Protection Required Equipment


10-3

Figure 10-2 Ice and Rain Protection Controls

10-4


ICE PROTECTION SYSTEMS

DESCRIPTION AND OPERATION

There are seven pilot-controlled anti-ice/deicesystems:

●

Windshield anti-ice

●

Surface deice (leading-edge boots)

●

Inertial separators (ice vanes)

●

Pitot heat

●

Propeller deice

●

Stall warning heat

●

Fuel vent heat

The airplane is equipped with a variety of ice andrain protection systems that can be utilized dur-ing opera t ion under inc lement wea therconditions. Electrical heating elements embed-ded in the windshie ld provide adequateprotection against the formation of ice, while air

from the cabin heating systems prevents fogging,to ensure visibility during operation under icingconditions. Heavy-duty windshield wipers forboth the pilot and copilot provide further visibil-ity during rainy flight and ground conditions.

Pneumatic deicer boots on the wings and on thevertical and horizontal stabilizers remove the for-mation of ice during flight. Regulated bleed-airpressure and vacuum are cycled to the pneumaticboots for the inflation-deflation cycle. The selec-tor switch that controls the system permitsautomatic single-cycle operation or manualoperation.

Ice protection for the engine is provided by aninertial separation system utilizing an electricalactuator. Should the main electrical actuatormotor fail, a standby actuator motor is provided.The leading-edge lip of the engine air inlet iscontinuously anti-iced by engine exhaust air. Thepropellers are protected against icing by electro-thermal boots on each blade that automaticallycycle to prevent the formation of ice.

A heating element in both pitot masts preventsthe pitot openings from becoming clogged withice. The heating elements are connected to theairplane electrical system through two 5-amperecircuit-breaker switches.

Figure 10-3 Pitot Mast and Heat Controls


10-5

PITOT HEAT

Two pitot masts located on the nose of the air-craft contain heating elements to protect againstice accumulation (Figure 10-3). The pitot mastsare electrically heated to ensure proper airspeedis indicated during icing conditions. Pitot heat iscontrolled by two circuit-breaker switcheslocated on the pilot’s right subpanel. The twoswitches placarded “PITOT,” one for the leftmast and one for the right, are located next to thestall warning anti-ice switch. They are two-posi-tion switches, with down being OFF and upbeing ON.

The pitot heat system should not be operated onthe ground, except for testing or for short inter-vals to remove snow or ice from the mast. Pitotheat should be turned on for takeoff and can beleft on in flight during icing conditions, or when-ever icing conditions are expected. If duringflight at altitude there is a gradual reduction inairspeed indication, there may be pitot icing. Ifturning on the pitot heat restores airspeed, leavethe pitot heat on because icing conditions exist.With many pilots, it is standard practice to keepthe pitot heat on during all flights at higher alti-tudes to prevent pitot icing.

STALL WARNING VANE

The stall warning vane and plate (Figure 10-4) isprovided with heat to ensure against freeze-upduring icing conditions. The stall warning plate isactivated by a two-position switch located just tothe right of the surface deicer cycle switch on thepilot’s right subpanel. The down position is OFF,and the up position is ON. The vane is heatedthrough the battery switch, so it is heated whenthe battery switch is ON.

A safety switch on the left landing gear limits thecurrent flow to approximately 12 volts to preventthe vane from overheating while the airplane ison the ground. In flight, after the left strutextends, the full 28-volt current is applied to thestall warning vane. The heating elements protectthe lift transducer vane and face plate from ice. Abuildup of ice on the wing may change or disruptthe airflow and prevent the system from accu-rately indicating an imminent stall. Rememberthat the stall speed increases whenever ice accu-mulates on any airplane.

Figure 10-4 Stall Warning Vane and Heat Control

10-6


FUEL SYSTEM ANTI-ICE

There are several anti-ice systems to protect fuelflow through the fuel lines to the engine (Figure10-5). Without heat, moisture in the fuel couldfreeze and diminish or cut off the fuel flow to theengines in freezing temperatures.

Ice formation in the fuel vent system is preventedby electrically heated vents in each wing. The

fuel vent heat is operated by left and rightswitches located in the ICE PROTECTIONgroup on the pilot’s right subpanel. Theseswitches should be turned on whenever ice isanticipated or encountered.

A portion of the fuel control unit ice protection isprovided by an oil-to-fuel heat exchanger,mounted on the engine’s accessory section. Anengine oil line within the heat exchanger is

Figure 10-5 Fuel System Anti-ice


10-7

located around the fuel line. Heat transfer occursthrough conduction. This heat melts ice particleswhich may have formed in the fuel. This opera-tion is automatic whenever the engines arerunning. Refer to the

POH

“Limitations” sectionfor temperature limitations concerning the oil-to-fuel heat exchanger.

The pneumatic line, from the engine to the FCUand the pneumatic line from the FCU to the fueltopping governor, is protected by an electricallyheated jacket. This heat is automatically appliedwhen the condition levers move out of the fuelcutoff range. No other action is required.

WINDSHIELD WIPERS

Separate windshield wipers are mounted on thepilot’s and copilot’s windshield. The dual wipersare driven by a mechanism operated by a singleelectric motor, all located forward of the instru-ment panel.

The windshield wiper control is located on theoverhead light control panel (Figure 10-6). It pro-vides the wiper mechanism with SLOW, FAST,and PARK positions. The wipers may be usedeither on the ground or in flight, as required. Thewipers must not be operated on a dry windshield.The windshield wiper circuit breaker is on thecopilot’s right-side circuit-breaker panel in theWEATHER group.

WINDSHIELD ANTI-ICE

The pilot’s and copilot’s windshields each haveindependent controls and heating circuits. Thecontrol switch allows the pilot to select a high ora low intensity heat level.

The windshields are composed of three physicallayers (Figure 10-7). The inner layer is a thickpanel of glass that acts as the structural member.The middle layer is a polyvinyl sheet which car-ries the fine wire heating grids. The outer layer isa protective layer of glass bonded to the first twolayers. The outside of the windshield is treatedwith a static discharge film called a “NESAcoating.”

Figure 10-6 Windshield Wipers

Figure 10-7 Windshield Installation

10-8


The windshields are protected against icing byelectrical heating elements (Figure 10-8). Theheating elements are connected at terminalblocks in the corner of the glass to the wiringleading to the control switches mounted in thepilot’s right subpanel.

A transparent material (usually stannic oxide)which has high electrical resistance is incorpo-rated in the laminations of each windshield, pilot’sand copilot’s. Each windshield is also fitted withelectrical connections for the resistive materialand for temperature-sensing elements. The resis-tive material is arranged so as to provide primaryheated surfaces and secondary heated surfaces.

PILOT and COPILOT WSHLD ANTI-ICEswitches in the ICE PROTECTION group on thepilot’s inboard subpanel are used to control wind-shield heat (Figure 10-9). They have positionslabeled “NORMAL,” “OFF,” and “HI.” When thePILOT and COPILOT switches are in the NOR-MAL (up) position, the secondary areas of thewindshields are heated. When the switches are inthe HI (down) position, the primary areas areheated. The primary areas are smaller areas andare heated faster to the same temperatures as theNORMAL position.

Each switch must be lifted over a detent before itcan be moved into the HI position. This lever-lock feature prevents inadvertent selection of theHI position when moving the switches fromNORMAL to the OFF (center) position.

Windshield temperature is controlled automati-cally by the use of a temperature-sensing elementembedded in each windshield, and a temperaturecontroller in each windshield circuit. The temper-ature controllers operate between 90 and 110º Fto maintain the desired mean temperature of thewindshield heating surfaces.

When the low level of heating is selected, anautomatic temperature controller senses thewindshield and attempts to maintain it at approx-imately 90 to 110º F. It does so by energizing the“low” heat relay as necessary. In this mode, theentire windshield is heated (Figure 10-10).

When the high level of heating is selected, thesame temperature controller senses the wind-shield temperature and attempts to maintain it at90 to 110º F. In this mode, however, the control-ler will energize the high heat relay switch,which applies the electrical heat to a more con-centrated but more essential viewing area of thewindshield. In high, approximately two-thirds ofthe windshield is heated at the outboard portion(Figure 10-11).

Figure 10-8 Windshield Anti-ice Diagram

Figure 10-9 Windshield Anti-ice Switches


10-9

The power circuit of each system is protected by50-ampere current limiters located in the powerdistribution panel. Windshield heater control cir-cuits are protected with 5-ampere circuit breakerslocated on a panel mounted on the forward pres-sure bulkhead (forward of the pilot’s leftsubpanel).

Windshield heat may be used at any time and inany combination. Use of windshield heat, how-ever, may cause erratic operation of the magnetic

compass because of the electrical field created bythe heating elements.

CAUTION

In the event of windshield icing duringsustained icing conditions, it may benecessary to reduce the airspeed inorder to keep the windshield ice-free.

Figure 10-10 Windshield Anti-ice Diagram—Normal Heat

10-10


ENGINE AIR INLET LIP HEAT

The lip around each air inlet is heated by hotexhaust gases to prevent the formation of ice dur-ing inclement weather (Figure 10-12).

A scoop in the left engine exhaust stack deflectsthe hot exhaust gases downward into the hollowlip tube that encircles the engine air inlet. Thegases are expelled through a line into the rightexhaust stack, where they move out with theengine exhaust gases.

Heat will flow through the inlet whenever theengine is running.

Figure 10-11 Windshield Anti-ice Diagram—High Heat

Figure 10-12 Engine Air Inlet Lip Heat


10-11

ENGINE INERTIAL SEPARATORS

An inertial vane system of separators is installedon each engine to prevent ice, or other foreignobjects such as dust or gravel, from entering theengine inlet plenum or ice accumulating on theengine inlet screen. A movable vane and a bypassdoor are closed (retracted) for normal flying con-ditions (Figure 10-13).

At temperatures above +5º C, the ice vane anddoor should be in the retract position, as ice for-mation is unlikely.

When in icing conditions with the ice vane in theextend position (Figure 10-14), the ice vane ispositioned to create a venturi effect and intro-duces a sudden turn into the engine. At the sametime the bypass door in the lower cowling at theaft end of the air duct is open.

Figure 10-14 Inertial Separator in Extend Position

Figure 10-13 Inertial Separator in Retract Position

10-12


As the ice particles or water droplets enter the airinlet, the airstream with these particles is acceler-ated by the venturi effect. Due to their greatermass, and therefore greater momentum, the fro-zen moisture particles accelerate past the screenarea and are discharged overboard through thebypass door. The airstream, however, makes thesudden turn free of ice particles and enters theengine through the inlet screen.

ICE VANE CONTROLS

The ice vane and bypass doors are extended orretracted simultaneously through a linkage sys-tem connected to electric actuators. The actuatorsare energized through switches in the ICE PRO-TECTION group located on the pilot’s leftsubpanel (Figure 10-15). The ICE VANEswitches extend the separators in the on positionand retract them in the OFF position, which isused for all normal flight operations.

The ice vanes should be extended whenever thereis visible moisture at +5º C. When the ice vanesare extended, the two green advisory annuncia-tors will illuminate, and because the airflow intothe engine will be restricted, there will be a dropin torque and a slight increase in ITT. When theice vanes and bypass doors are retracted, theannunciators will extinguish, torque will berestored, and ITT will decrease.

The anti-ice vanes are controlled by switcheslocated on the left subpanel. The LEFT andRIGHT ENGINE ANTI-ICE switches have posi-tions labeled “ON” and “OFF,” while the

ACTUATORS switch has positions labeled“STANDBY” and “MAIN.”

The actuators have dual motors to provide aredundant system. The ACTUATORS switchallows the selection of either the MAIN orSTANDBY actuator motor. The main andstandby actuators have different circuitry butshare the same torque tube drive system.

The vanes have only two positions; there are nointermediate positions. The system is monitoredby L and R ENG ANTI-ICE (green) and L and RENG ICE FAIL (yellow) annunciators (Figure10-16). Illumination of the L and R ENG ANTI-ICE annunciators indicate that the system isactuated.

Illumination of the L or R ENG ICE FAILannunciator indicates that the system did notoperate to the desired position. Immediate illumi-nation of the L or R ENG ICE FAIL annunciatorindicates loss of electrical power, whereasdelayed illumination indicates an inoperativeactuator.

The yellow ENG ICE FAIL annunciator circuitcompares the ANTI-ICE switch position to themicroswitches checking ice vane open or closed.After a 35-second delay, the annunciator willi l l umina t e i f t he sw i t ch pos i t i on andmicroswitches do not agree. In addition, if thepower source for the actuator system selected(MAIN or STANDBY) is removed, the ICEVANE FAIL light will illuminate immediately. Ineither event, the STANDBY actuator should beselected.

Figure 10-15 Ice Vane Controls

Figure 10-16 Caution and Advisory Annunciators


10-13

ENGINE AUTOIGNITION SYSTEM

The engine autoignition system provides auto-matic ignition to attempt a restart should aflame-out occur. Once armed, the systemensures ignition during takeoff, landing, turbu-lence, and penetration of icing or precipitationconditions. Should ice or rain cause an engineflameout, autoignition will automatically reig-nite the engine.

The switches used to arm the autoignition sys-tem are located on the pilot’s left subpanel,above the ice vane switches and just to the leftof the control column (Figure 10-17). The sys-tem is activated by moving the switches into theup or ARM position. Each switch must be liftedover a lock-out barrier before it can be movedinto, or out of, the ARM position. This lever-lock feature prevents inadvertent movement tothe OFF position.

If, for any reason, engine torque falls below fourhundred foot-pounds, electrical power is pro-vided to energize the engine igniters. As thishappens, the green IGNITION ON annunciatoron the panel will illuminate, indicating that theignition system is energized. During groundoperation, the system should be turned off to pro-long the life of the igniter units.

PROPELLER ELECTRIC DEICE SYSTEM

The propeller electric deicer system includes: anelectrically heated boot for each propeller blade,slip rings, brush assemblies, timer, on-off switch,and an ammeter (Figure 10-18).

When the switch is turned on, the ammeter reg-isters the amount of current (14 to 18 ampereson the C90A or 18 to 24 amperes on the C90B)passing through the system. If the current risesbeyond the limitations, a circuit-breaker switchor current limiter will shut off power to thedeicer timer. The current flows from the timerthrough the brush assemblies to the slip rings,where it is distributed to the individual propellerdeicer boots.

Heat produced by the heating elements in thedeicer boots reduces the adhesion of the ice.Adhesion thus reduced, the ice is removed by thecentrifugal effect of the propeller and the blast ofthe airstream.

NOTE

The heating sequences for the deicerboots noted in the following sectionare the sequences which are in evi-dence during normal operation.

Power to the deicer boots is cycled in 90-secondphases. The first 90-second phase heats all thedeicer boots on the RH propeller. The secondphase heats all the deicer boots on the LH propel-ler. The deicer timer completes one full cycleevery three minutes. As the deicer timer movesfrom one phase to the next, a slight momentarydeflection of the propeller ammeter needle maybe noted. Propeller deice must not be operatedwhen the propellers are static. Figure 10-17 Engine Autoignition Switches

10-14


WING ICE LIGHTS

Wing ice lights are provided to light the wingleading edges to determine ice buildup in icingconditions. The wing lights are located on theoutboard side of each nacelle. The circuit-breakerswitch is located on the pilot’s right subpanel inthe LIGHTS group above the ICE protectiongroup (Figure 10-19).

The wing ice lights should be used as required innight flight to check for wing ice accumulation.The wing ice lights operate at a high temperatureand therefore should not be used for prolongedperiods while the airplane is on the ground. Allice lights installed must be operational for flightsinto known or forecast icing conditions at night

Figure 10-18 Propeller Electric Deice System

Figure 10-19 Wing Anti-ice Lights


10-15

PRECAUTIONS DURING ICING CONDITIONS

There are some precautions which prevail duringwinter or icing conditions. An airplane needs spe-cial care and inspection before operation in cold orpotential icing weather. In addition to the normalexterior inspection, special attention should be paidto areas where frost and ice may accumulate.

Pilots should be familiar with the potential harma harmless-looking, thin layer of frost can cause.It is not the thickness of the frost that matters; itis the texture. A slightly irregular surface cansubstantially decrease proper airflow over thewings and stabilizers. Never underestimate thedamaging effects of frost. All frost should beremoved from the leading edges of the wings,stabilons, stabilizers, and propellers before theairplane is moved.

Control surfaces, hinges, the windshield, pitotmasts, fuel tank caps, and vents should also befree of frost. Deicing fluid should be usedwhen needed.

Fuel drains should be tested for free flow. Waterin the fuel system has a tendency to condensemore readily during winter months, and if leftunchecked, large amounts of moisture may accu-mulate in the fuel tanks. Moisture does notalways settle at the bottom of the tank. Occasion-ally a thin layer of fuel gets trapped under a largemass of water, which may deceive the tester.Make sure a good-sized sample of fuel is taken.

It is also important to add only the correctamount of anti-icing additive to the fuel. A higherconcentration of anti-icer does not ensure lowerfuel freezing temperatures and may hinder theairplane’s performance. Consult the “NormalProcedures” section of the

Pilot’s OperatingHandbook

to determine the correct blend.

The brakes and tire-to-ground contact should bechecked for lockup. No anti-ice solution contain-ing oil-based lubricant should be used on thebrakes. If tires are frozen to the ground, use undi-luted defrosting fluid or a ground heater to meltice around the tires, then move the airplane as

soon as the tires are free. Heat applied to tiresshould not exceed 160° F or 71° C.

Tiedowns for propellers should be installed toensure against damage to internal engine compo-nents not lubricated when the engine is notoperating. Spinning propellers can also be asource of danger to crew, passengers, and groundsupport personnel. Propeller blades held in theirtiedown position channel moisture down theblades, past the propeller hub, and off the lowerblade more effectively than in other positions orwhen left spinning. During particularly icyground conditions, the propeller hubs should alsobe inspected for ice and snow accumulation.

Pitot masts should always be covered while theairplane is resting. Once the covers are removed,make sure both masts and drains are free of ice orwater. Faulty readings could be obtained if theyare clogged.

During extended periods of taxiing or groundholding, the autoignition system should be turnedoff until right before takeoff. This will help toprolong the service life of the igniter units.

Snow, slush, or standing water on the runwaydegrade airplane performance whether landing ortaking off. During takeoff, more runway isneeded to achieve necessary takeoff speed, whilelanding roll is longer because of reduced brakingeffectiveness.

Only the surface deicers are true deicers. The restare really anti-icers and should be used to preventthe formation of ice, not melt ice already present.Accumulated ice on even the best-equipped air-plane will degrade its performance and ruin atleast the time and fuel calculations used for flightplanning. A minimum speed of 140 KIAS is nec-essary to prevent ice formation on the undersideof the wing, which cannot be adequately deiced.

Due to distortion of the wing airfoil, stalling air-speeds should be expected to increase as iceaccumulates on the airplane. For the same rea-son, stall warning devices are not accurate andshould not be relied upon. Maintain a comfort-able margin of airspeed above the normal stallairspeed when ice is on the airplane. In order to

10-16


prevent ice accumulation on unprotected surfacesof the wing, maintain a minimum of 140 knotsduring operations in sustained icing conditions.In the event of windshield icing, it may be neces-sary to reduce airspeed.

While in flight, the engine ice vanes must beextended and the appropriate annunciator lightsmonitored:

●

Before visible moisture is encountered atOAT +5º C and below

●

At night when freedom from visiblemoisture is not assured and the OAT is+5º C or below

During flight in icing conditions, fuel vent heat,pitot heat, prop deice, windshield heat, and stallwarning heat should all be ON.


11-i

CHAPTER 11

CONTENTS

Page

AIR CONDITIONING

INTRODUCTION ................................................................................................................

11-1

DESCRIPTION.....................................................................................................................

11-1

ENVIRONMENTAL SYSTEM ...........................................................................................

11-3

UNPRESSURIZED VENTILATION...................................................................................

11-5

BLEED-AIR HEATING SYSTEM ......................................................................................

11-6

ELECTRIC HEAT................................................................................................................

11-9

COOLING SYSTEM..........................................................................................................

11-10

ENVIRONMENTAL CONTROLS....................................................................................

11-11

Automatic Mode Control.............................................................................................

11-12

Manual Mode Control .................................................................................................

11-12

Bleed-Air Control........................................................................................................

11-13

Vent Blower Control ...................................................................................................

11-13


11-iii

Figure Title Page

ILLUSTRATIONS

11-1

Environmental System Schematic .........................................................................

11-2

11-2

ENVIRONMENTAL Group Switches and Knobs................................................

11-3

11-3

Air Control Knobs—Pilot Air ...............................................................................

11-4

11-4

Air Control Knobs—Defrost Air ...........................................................................

11-4

11-5

Air Control Knobs—Cabin Air .............................................................................

11-4

11-6

Air Control Knobs—Copilot Air ...........................................................................

11-4

11-7

Ram-Air Scoop ......................................................................................................

11-4

11-8

Glareshield “Eyeball” Outlets................................................................................

11-5

11-9

Cabin Floor Outlets................................................................................................

11-5

11-10

Fresh Air Source (Unpressurized Mode) ...............................................................

11-5

11-11

Cabin “Eyeball” Outlets ........................................................................................

11-5

11-12

Cockpit “Eyeball” Outlets .....................................................................................

11-6

11-13

Ambient and Bleed Air Flow Forward of Firewalls ..............................................

11-6

11-14

Air Conditioning System Control Diagram...........................................................

11-7

11-15

Mixing Plenum ......................................................................................................

11-8

11-16

Electric Heater .......................................................................................................

11-9

11-17

Grid Heating Elements ..........................................................................................

11-9

11-18

ELEC HEAT Switch............................................................................................

11-10

11-19

Cooling System Components in Nose .................................................................

11-10

11-20

Receiver-Dryer Sight Gage..................................................................................

11-11

11-21

CABIN TEMP MODE Selector Switch ..............................................................

11-12

11-22

CABIN TEMP Level Control ..............................................................................

11-12

11-23

MANUAL TEMP Switch ....................................................................................

11-13

11-iv


11-24

BLEED AIR VALVE Switches ..........................................................................

11-13

11-25

VENT BLOWER Switch ....................................................................................

11-13


11-1

CHAPTER 11AIR CONDITIONING

INTRODUCTION

Passenger comfort and safety is of prime importance. The task is to teach participants to operatethe environmental systems effectively and within the system’s limitations.

DESCRIPTION

The Environmental System section of the train-ing manual presents a description and discussionof the air conditioning, bleed-air heating, and

fresh air systems. Each system includes generaldescription, principle of operation, controls, andemergency procedures.

11-2


Figure 11-1 Environmental System Schematic


11-3

ENVIRONMENTAL SYSTEM

“Environmental System” refers to the deviceswhich control the pressure vessel’s environment.Along with insuring the circulation of air, thissystem controls temperature by utilizing heatingand cooling devices as needed.

The environmental system consists of bleed-airpressurization, heating and cooling systems andtheir associated controls. The Beechcraft KingAir series environmental system (Figure 11-1)uses turbine engine bleed air for cabin pressur-ization and cabin heating. The air conditioningsystem, driven by the electrical system, providescool air to the airplane cabin.

The ENVIRONMENTAL control section on thecopilot’s left subpanel (Figure 11-2) provides forautomatic or manual control of the system. Thissection contains all the major controls of theenvironmental function:

●

Bleed-air valve switches

●

Vent blower control switch

●

Manual temperature switch for control ofthe bypass valves in the air-to-air heatexchangers

●

Cabin-temperature-level control

●

Cabin temperature mode selector switchfor selecting automatic heating or cool-ing, manual heating or cooling

●

Electric heat control switch

Figure 11-2 Environmental Group Switches and Knobs

11-4


Four additional manual controls (Figures 11-3through 11-6) on the main instrument subpanelsmay be utilized for partial regulation of cockpitcomfort when the cockpit partition curtain isclosed and the cabin comfort level is satisfactory.They are: pilot’s air, defroster air, cabin air, andcopilot’s air control knobs. The fully out positionof all these controls will provide the maximumheating to the cockpit, and the fully in positionwill provide minimum heating to the cockpit.

The pressurization, heating, and air conditioningsystems operate in conjunction with each other oras separate systems to maintain the desired cabinpressure altitude and cabin air temperature.Occupied compartments are pressurized, heated,or cooled through a common ducting arrange-ment. Ventilation can be obtained on demandduring nonpressurized flight through a ram-airscoop on the left side of the nose (Figure 11-7).

Figure 11-3 Air Control Knobs—Pilot Air

Figure 11-4 Air Control Knobs—Defrost

Figure 11-5 Air Control Knobs—Cabin Air

Figure 11-6 Air Control Knobs—Copilot Air

Figure 11-7 Ram-Air Scoop


11-5

UNPRESSURIZED VENTILATION

Fresh-air ventilation is provided from twosources. One source, which is available duringboth the pressurized and the unpressurized mode,is the bleed-air heating system. This air mixeswith recirculated cabin air and enters the cockpitthrough glareshield “eyeball” outlets (Figure11-8) and the cabin through the floor registers(Figure 11-9). The volume of air from the floorregisters is regulated by using the cabin air con-trol knob located on the copilot’s subpanel.

The second source of fresh air, which is availableduring the unpressurized mode only, is ambientair obtained from a ram-air scoop (Figure 11-10)on the nose (left side) of the airplane. Duringpressurized operation, an electromagnet, in addi-tion to cabin pressure, forces the ram-air flapperdoor closed. During the unpressurized mode, ram

air enters the evaporator plenum through the ram-air door when the electromagnet releases. Recir-culated cabin air forced into the evaporatorplenum by a blower, mixes with ram air fromoutside, is ducted around the electric heater andmixing plenum and into the ceiling-outlet duct.Air ducted to each individual cabin (Figure11-11) or cockpit (Figure 11-12) ceiling eyeballoutlet can be directionally controlled by movingthe eyeball in the socket. Volume is regulated bytwisting the outlet to open or close the outlet.

Figure 11-8 Glareshield “Eyeball” Outlets

Figure 11-9 Cabin Floor Outlets

Figure 11-10 Fresh Air Source (Unpressurized Mode)

Figure 11-11 Cabin “Eyeball” Outlets

11-6


BLEED-AIR HEATING SYSTEM

Air pressure for cabin pressurization, heating thecabin and cockpit, and for operating the instru-ments, rudder boost, and surface deice isobtained by bleeding air from the compressorstage (P

3

) of each engine. When air is com-pressed, its temperature increases. Therefore, the

bleed air extracted from the compressor sectionof each engine for pressurization purposes is hot.This heat is utilized to warm the cabin.

Engine bleed air is ducted from the engine to theflow control unit mounted on the firewall. Thebleed air from either engine will continue to pro-vide adequate air for pressurization and heating,and for the deicer system and instruments, shouldone engine fail. The bleed air and ambient airfrom the cowling intake are mixed together bythe flow control units, and are routed aft throughthe firewall along the inboard side of eachnacelle, and inboard to the center section forwardof the main spar.

When the left landing gear safety switch is in theon-the-ground position, the ambient air valve(Figure 11-13) in each flow control unit is closed.Consequently, only bleed air is delivered to theenvironmental bleed-air duct when the airplane ison the ground. The exclusion of ambient airallows faster cabin warmup during cold weatheroperation. In flight, the ambient air valve is openwhen temperature is above -30°F, and ambientair is mixed with the engine bleed air in the flow

Figure 11-12 Cockpit “Eyeball” Outlets

Figure 11-13 Ambient and Bleed Air Flow Forward of Firewalls


11-7

control unit. During warm weather ground opera-tion, the engine bleed air into the cabin can beshut off by placing the bleed-air valve switcheson the copilot’s subpanel to the CLOSED posi-tion. Closing the bleed-air valves prevents warmbleed air from entering the cabin area, maximiz-ing the air conditioner operation.

The heat in the air may either be retained forcabin heating or dissipated for cooling purposesas the air passes through the center section to thefuselage. If the environmental bleed-air mixtureis too warm for cabin comfort, the cabin temper-ature control bypass valve (Figure 11-14) routessome or all of it through the air-to-air heat

exchanger in the wing center section. The posi-tion of the damper in the cabin temperaturecontrol bypass valve is determined by position-ing of the controls in the ENVIRONMENTALgroup on the copilot’s subpanel. An air intake onthe leading edge of the inboard wing brings ramair into the heat exchanger to cool the bleed air.

Depending upon the position of the cabin tem-perature control bypass valves, a greater or lesservolume of the bleed-air mixture will be routedthrough or around the heat exchanger. The tem-perature of the air flowing through the heatexchanger is lowered as heat is transferred tocooling fins, which are in turn cooled by ram air-

Figure 11-14 Air Conditioning System Control Diagram

11-8


flow through the fins of the heat exchanger. Afterleaving the heat exchanger, the ram air is ductedoverboard through louvers on the underside ofthe wing.

The bleed air leaving both (left and right) cabintemperature control bypass valves is thenducted into a single muffler under the rightfloorboard forward of the main spar, whichinsures quiet operation of the environmentalbleed-air system. The air mixture is then ductedfrom the muffler into the mixing plenum underthe copilot’s floorboard.

A partition divides the mixing plenum into twosections. One section supplies the floor-outletduct, and the other supplies the ceiling outletduct. Both sections receive recirculated cabin airfrom the vent blower. The air passes through theforward evaporator, so it will be cooled if the airconditioner is operating. Even in the event thevent blower becomes inoperative, some air willstill be circulated, due to the duct design in thedischarge side of the mixing plenum.

The environmental bleed-air duct is routed intothe floor-duct section of the mixing plenum, thencurves back to discharge the environmental bleedair toward the aft end of the floor duct section ofthe mixing plenum. Forward of the discharge endof the environmental bleed-air duct (Figure11-15), warm air is tapped off and ducted upthrough the top of the mixing plenum and isdelivered to the pilot/copilot heat duct, which isbelow the instrument panel. An outlet at each endof this duct is provided to deliver warm air to thepilot and copilot. A mechanically controlleddamper in each outlet permits the volume of air-flow to be regulated. The pilot’s damper iscontrolled by the PILOT AIR (see Figure 11-3)knob, on the pilot’s left subpanel, just outboardof the control column. The copilot’s damper iscontrolled by the COPILOT AIR (see Figure11-6) knob, on the copilot’s right subpanel, justoutboard of the control column. The DEFROSTAIR control knob (see Figure 11-4) is on thepilot’s right subpanel, just inboard of the controlcolumn. This knob controls a valve at the for-

ward side of the pilot/copilot heat duct whichadmits air to two ducts that deliver the warm airto the defroster, just below the windshields in thetop of the glareshield. An air plenum built into theglareshield feeds air to “eyeball” outlets on the leftand right sides. Defrost air is the air source for thepilot and copilot glareshield “eyeball” outlets;thus, the use of the DEFROST AIR control knobalso controls air to these eyeball outlets.

The remainder of the air in the environmentalbleed-air duct is discharged into the floor-out-let duct section of the mixing plenum andmixed with recirculated cabin air. This air mix-ture passes through the cabin air control valve.This valve is controlled by the CABIN AIRcontrol knob (see Figure 11-5) on the copilot’ssubpanel, just below and inboard of the controlcolumn. When this knob is pulled out to thestop, only a minimum amount of air will bepermitted to pass through the valve, therebyincreasing the amount of air available to thepilot and copilot outlets, and to the defroster.When this knob is pushed fully in, the valve isopen and the air in the duct will be directed tothe floor-outlet registers in the cabin.

Figure 11-15 Mixing Plenum


11-9

ELECTRIC HEAT

Additional heating is available from an electricalheater (Figure 11-16) containing eight heatingelements rated at 1,000 watts each. The eightelectrical heating elements (Figure 11-17) aredivided into two sets with four elements in eachset. One set provides heat for NORMAL HEAToperation and both sets combine for GROUNDMAX HEAT operation. The maximum output isavailable during ground operation and only fourelements are available during flight. The airplaneelectrical system is protected against an overloadby a lockout circuit that prevents use of the elec-trical heater during operation of the propellerdeicers or windshield heat

The ELEC HEAT switch (Figure 11-18), in theENVIRONMENTAL group in the copilot’s sub-panel, has three positions: GND MAX – NORM– OFF. This switch is solenoid-held in GNDMAX position on the ground and drops to

NORM position when the landing gear safetyswitch is opened at lift-off. It provides maximumelectric heat for initial warmup of the cabin. Ifuse of all electrical heating elements is notdesired for initial warmup, as in the GND MAXposition, the switch may be placed in the NORMposition, using only four elements. In the NORMposition the four heating elements automatically

Figure 11-16 Electric Heater

Figure 11-17 Grid Heating Elements

11-10


supplement bleed-air heating, in conjunctionwith the cabin thermostat. The OFF positionturns off all electric heat, leaving only bleed airto supply cabin heat.

COOLING SYSTEM

Cabin cooling is provided by a refrigerant-gasvapor-cycle refrigeration system consisting of:

●

Belt-driven compressor, installed in the nose

●

Condenser coil

●

Condenser blower

●

Evaporator

●

Receiver-dryer

●

Expansion valve

●

Cabin heat control valve

It is routed (Figure 11-19) to the condenser coil,receiver-dryer, expansion valve, cabin heat con-trol valve, and evaporator, which are all in thenose of the airplane. The rated output of the stan-dard instal la t ion in the fuselage nose is16,000 BTU.

The evaporator utilizes a solenoid-operated, hot-gas-cabin heat control valve to prevent icing. A33° F thermal switch on the evaporator controlsthe valve solenoid.

Figure 11-18 Elec Heat Switch

Figure 11-19 Cooling System Components in Nose


11-11

The vent blower blows recirculated cabin airthrough the evaporator, into the mixing plenum,and into both the floor-outlet and ceiling outletducts. If the cooling mode is operating, refriger-ant will be circulating through the evaporator andthe air leaving it will be cool. All the air enteringthe ceiling-outlet duct will be cool. This air isdischarged through “eyeball” outlet nozzles inthe cockpit and cabin. Each nozzle is movable, sothat the airstream can be directed as desired.When the nozzle is twisted, a damper opens orcloses to regulate airflow volume.

Cool air will enter the floor-outlet duct, but inorder to provide cabin pressurization, warm envi-ronmental bleed air will also enter the floor-outlet duct anytime either BLEED AIR valve isOPEN. Therefore, pressurized air dischargedfrom the floor registers will always be warmerthan that discharged at the ceiling outlets, nomatter what temperature mode is in use.

A condenser blower in the nose section drawsambient air through the condenser when the airconditioner is operating. The receiver-dryer andsight gage (Figure 11-20) are in the upper portionof the nose wheel well.

ENVIRONMENTAL CONTROLS

The ENVIRONMENTAL control section on thecopilot’s subpanel (see Figure 11-2) provides forautomatic or manual control of the system. Thissection contains all the major controls of theenvironmental function:

●

Bleed-air valve switches

●

Vent blower control switch

●

Manual temperature switch for control ofthe bypass valves in the air-to-air heatexchangers

●

Cabin-temperature-level control

●

Cabin temperature mode selector switch,for selecting automatic heating or cool-ing, manual heating or cooling, or off

●

Electric heat control switch

Figure 11-20 Receiver-Dryer Sight Gage

11-12


Four additional manual controls on the maininstrument subpanels may be utilized for partialregulation of cockpit comfort when the cockpitpartition curtain is closed and the cabin comfortlevel is satisfactory. They are: pilot’s air,defroster air, cabin air, and copilot’s air controlknobs. The fully out position of all these controlswill provide the maximum heating to the cockpit,and the fully in position will provide maximumheating to the cabin.

For warm flights, such as short, low-altitudeflights in summer, all the cabin floor registers andceiling outlets should be fully open for maximumcooling. For cold flights, such as high-altitudeflights, night flights, and flights in cold weather,the ceiling outlets should all be closed and thefloor outlets fully open for maximum heating inthe cabin.

AUTOMATIC MODE CONTROL

When the CABIN TEMP MODE selector switch(Figure 11-21) on the copilot’s subpanel is in theAUTO position, the heating and air conditioningsystems operate automatically. The systems areconnected to a control box by means of a bal-anced br idge c i rcui t . I f a warmer cabintemperature has been selected, the automatictemperature control modulates the cabin heatcontrol valves one at a time to allow heated air tobypass the air-to-air heat exchangers in the wingcenter sections. This warm bleed air is thenbrought into the cabin where it is mixed withrecirculated cabin air in the floor ducting underthe copilot floor area. The automatic temperaturecontrol system will then modulate the bypassvalves to maintain the proper temperature of theincoming bleed air.

When the au tomat i c con t ro l d r ives theenvironmental system from a heating mode to acooling mode, the bypass valves move toward thecool position (bleed air passes through the air-to-air heat exchanger). When the left valve reachesthe full cold position, the air-conditioning systemwill begin cooling. When the left bypass valve ismoved approximately 30° toward the heatposition the air-conditioning system will turn off

preventing unnecessary recycling of the air-conditioning system.

The CABIN TEMP – INCR (Figure 11-22) con-trol provides regulation of the temperature levelin the automatic mode. A temperature-sensingunit in the cabin, in conjunction with the controlsetting, initiates a heat or cool command to thetemperature controller, requesting the desiredpressure-vessel environment.

MANUAL MODE CONTROL

When the CABIN TEMP MODE selector is inthe MAN HEAT or MAN COOL position, regu-lation of the cabin temperature is accomplished

Figure 11-21 Cabin Temp Mode Selector Switch

Figure 11-22 Cabin Temp Level Control


11-13

manually by momentarily holding the MANUALTEMP switch (Figure 11-23) to either the INCRor DECR position as desired. When released, thisswitch will return to the center (no change) posi-tion. Moving this switch to the INCR or DECRposition results in modulation of the bypassvalves in the bleed-air lines. Allow approxi-mately 30 seconds per valve (one minute totaltime) for the valves to move to the full heat orfull cold position. Only one valve moves at atime. Movement of these valves varies theamount of bleed air routed through the air-to-airheat exchanger. Consequently, the temperature ofthe incoming bleed air will vary. This bleed airmixes with recirculated cabin air (which will beair conditioned if the refrigeration system isoperating) in the mixing plenum, and is thenducted to the floor registers. As a result, the cabintemperature will vary according to the position ofthe bypass valves, whether or not the air condi-tioner is operating.

When the CABIN TEMP MODE selector is inthe MAN COOL position, the air-conditioningsystem will operate, provided the bypass valvesare positioned full cool, until turned off, or theevaporator reaches 33° F when the thermal sen-sor turns air conditioning off.

BLEED-AIR CONTROL

Bleed air entering the cabin is controlled by thetwo switches (Figure 11-24) placarded BLEEDAIR VALVES – OPEN – CLOSED. When the

switch is in the OPEN position, the environmen-tal flow control units are open. When the switchis in the CLOSED position, the environmentalflow control unit is closed. For maximum coolingon the ground, turn the bleed-air valve switchesto the CLOSED position.

VENT BLOWER CONTROL

The forward vent blower is controlled by aswitch in the ENVIRONMENTAL group (Figure11-25) placarded VENT BLOWER – HIGH –LO – AUTO. When this switch is in the AUTOposition, the vent blower will operate at lowspeed if the CABIN TEMP MODE selectorswitch is in any position other than OFF (i.e.,MANual COOL, MANual HEAT, or AUTO-matic), with one exception. The vent blower willoperate in high if GND MAX HEAT is selected.

Figure 11-23 Manual Temp Switch

Figure 11-24 Bleed Air Valve Switches

Figure 11-25 Vent Blower Switch

11-14


When the VENT BLOWER switch is in theAUTO position and the CABIN TEMP MODEselector switch is in the OFF position, the blowerwill not operate. Anytime the VENT BLOWERswitch is in the LO position, the vent blower willoperate at low speed, even if the CABIN TEMPMODE selector switch is OFF. Anytime theVENT BLOWER switch is in the HIGH position,the vent blower will operate at high speed,regardless of the position of the CABIN TEMPMODE selector switch (i.e., MAN COOL, MANHEAT, OFF, or AUTO).


12-i

CHAPTER 12

CONTENTS

Page

PRESSURIZATION

INTRODUCTION ................................................................................................................

12-1

DESCRIPTION.....................................................................................................................

12-1

PRESSURIZATION SYSTEM ............................................................................................

12-3

AIR DELIVERY SYSTEM ..................................................................................................

12-4

CABIN PRESSURE CONTROL..........................................................................................

12-8

PREFLIGHT CHECK.........................................................................................................

12-10

IN FLIGHT .........................................................................................................................

12-10

DESCENT...........................................................................................................................

12-10

FLOW CONTROL UNIT...................................................................................................

12-11


12-iii

Figure Title Page

ILLUSTRATIONS

12-1

Pressurization and Air Conditioning Distribution System ....................................

12-2

12-2

Cabin Altitude for Various Airplane Altitudes Graph...........................................

12-3

12-3

Bleed Air Valves Switches ....................................................................................

12-4

12-4

Cabin Air Outflow Valve.......................................................................................

12-5

12-5

Cabin Air Safety Valve..........................................................................................

12-5

12-6

Pressurization Controls Schematic ........................................................................

12-6

12-7

Bleed Air Control (Pressurization and Pneumatics) ..............................................

12-7

12-8

Pressurization Controller .......................................................................................

12-8

12-9

Cabin Altimeter .....................................................................................................

12-8

12-10

Cabin Climb Indicator ...........................................................................................

12-9

12-11

Cabin Pressure Switch ...........................................................................................

12-9

12-12

Environmental System Circuit Breakers ...............................................................

12-9

12-13

Pressurization Controller Setting for Landing.....................................................

12-10

12-14

Flow Control Unit................................................................................................

12-11


12-1

CHAPTER 12PRESSURIZATION

INTRODUCTION

Pressurization is desirable in an airplane because it allows the altitude of the cabin to be lowerthan the altitude of the airplane, thus decreasing or eliminating the need for supplementaryoxygen. In this section, the pilot learns how the system operates, is controlled, and how tohandle malfunctions of the system.

DESCRIPTION

The Pressurization System section of the trainingmanual presents a description of the pressuriza-tion system. The function of various majorcomponents, their physical location, and opera-

tion of the pressurization system controls arediscussed. Where necessary, references are madeto the environmental system as i t affectspressurization.

��

��

12-2


Figure 12-1 Pressurization and Air Conditioning Distribution System


12-3

PRESSURIZATION SYSTEM

The pressurization system (Figure 12-1) isdesigned to provide a cabin environment withsufficient oxygen for normal breathing, regard-less of the airplane altitude, up to its designceiling. As the airplane altitude increases, theoutside ambient air pressure decreases until, atapproximately 12,500 feet, it cannot support nor-mal respiration. The pressurization systemmaintains a proportionally lower inside cabinaltitude. The pressure differential between the

inside cabin pressure and the outside ambient airpressure is measured in pounds per square inch.

As the cabin altitude chart shows (Figure 12-2),whenever cabin altitude and airplane altitude arethe same, no pressure differential exists. When-ever cabin pressure is the greater of the two,pressure differential is a positive number. If cabinpressure is less than that of the outside ambientair, pressure differential is a negative number.Maximum differential is defined as a measure ofthe highest positive differential pressure the air-plane structure can safely withstand for anextended period of time.

Figure 12-2 Cabin Altitude for Various Airplane Altitudes Graph

12-4


The King Air C9OA and C9OB, equipped withPT6A-21 engines maintain a 5.0 ±0.1 psi differ-ential and provides a cabin pressure altitude ofapproximately 6,000 feet at an airplane altitudeof 20,000 feet; and 12,000 feet at 30,000 feet.Although the King Air’s pressure vessel isdesigned to withstand a maximum differentialgreater than 5.0 psi, the airplane structure is notdesigned to withstand a negative differential.

The pressurization and environmental systems(Figure 12-1) operate in conjunction with eachother or as separate systems to maintain thedesired cabin pressure altitude and cabin air tem-pe ra tu r e . Occup i ed compar tmen t s a r epressurized, heated, or cooled through a commonducting arrangement.

“Pressure vessel” means that portion of the air-craft designed to withstand the pressuredifferential. In the King Air, the pressure vesselextends from a forward pressure bulkhead,between the cockpit and nose section to a rearpressure bulkhead, just aft of the cabin baggagecompartment, with exterior skins making up theouter seal. Windows are round for maximumstrength. All cables, wire bundles, and plumbingpassing through the pressure vessel boundariesare sealed to reduce leaks.

AIR DELIVERY SYSTEM

Bleed air from the compressor section of eachengine is utilized to pressurize the pressure ves-sel. A flow control unit in the nacelle of eachengine controls the flow of the bleed air andmixes ambient air with it to provide an air mix-ture suitable for the pressurization function. Themixture flows to the environmental bleed airshutoff valve, which is a normally closed sole-noid. This solenoid is controlled by a switchplacarded BLEED AIR VALVES – LEFT (or)RIGHT OPEN – CLOSED in the ENVIRON-MENTAL controls group (Figure 12-3) on thecopilot’s left subpanel. When this switch is in theCLOSED position, the solenoid is closed and nobleed air can enter the flow control unit or thecabin. When the BLEED AIR VALVE switch isin the OPEN position, the solenoid is electrically

held open and the air mixture flows through thevalve to the flow control package. Electricity isrequired to keep the flow control solenoid open.If there were a complete electrical failure, thesolenoid would fail to the closed position. Nomore bleed air would enter the pressure vesseland the cabin pressure would leak out.

The air entering the airplane flows through theenvironmental bleed air duct (Figure 12-1). Theair from the environmental bleed air duct ismixed with recirculated cabin air (which may ormay not be air conditioned) in the mixing ple-num, ducted upward into the crew heat duct, thenrouted into the floor outlet duct. This pressurizedair is then introduced into the cabin through thefloor registers. This air may be recirculatedthrough the air conditioning system. Finally theair flows out of the pressure vessel through theoutflow valve (Figure 12-4), located on the aftpressure bulkhead. A silencer on the outflow andsafety/dump valves (Figure 12-5) ensures quietoperation.

Figure 12-3 Bleed Air Valves Switches


12-5

Figure 12-4 Cabin Air Outflow Valve

Figure 12-5 Cabin Air Safety Valve

12-6


The mixture from both flow control units is deliv-e red to the p ressure vesse l a t a r a te o fapproximately 14 pounds per minute, dependingupon ambient temperature and pressure altitude.Pressure within the cabin and the rate of cabinpressure changes are regulated by pneumaticmodulation of the outflow valve (Figure 12-6),which controls the rate at which air can escapefrom the pressure vessel.

A vacuum-operated safety valve is mounted adja-cent to the outflow valve on the aft pressurebulkhead. It is intended to serve three functions:

●

Provide pressure relief in the event ofmalfunction of the normal outflow valve

●

Allow depressurization of the pressurevessel whenever the cabin pressureswitch is moved into the DUMP position

●

Keep the pressure vessel unpressurizedwhile the airplane is on the ground, withthe left landing gear safety switchcompressed

A negative-pressure relief function is also incor-porated into both the outflow and the safetyvalves. This prevents outside atmospheric pres-sure from exceeding cabin pressure by more than0.l psi during rapid descents, even if bleed-airinflow ceases.

Figure 12-6 Pressurization Controls Schematic


12-7

When the BLEED AIR VALVE switches on thecopilot’s left subpanel are OPEN (up), the airmixture from the flow control units enters thepressure vessel. While the airplane is on theground, a left landing gear safety switch-actuatedsolenoid valve (Figure 12-7) in each flow controlunit keeps the ambient air modulating valveclosed, allowing only bleed air to be delivered

into the pressure vessel. At lift-off, the safetyvalve closes and the ambient air shutoff solenoidvalve in the left flow control unit opens; approxi-mately 6 seconds later, the solenoid in the rightflow control unit opens. Consequently, byincreasing the volume of airflow into the pressurevessel in stages, excessive pressure bumps duringtakeoff are avoided.

Figure 12-7 Bleed Air Control (Pressurization and Pneumatics)

12-8


CABIN PRESSURE CONTROL

An adjustable cabin pressurization controller(Figure 12-8) is mounted in the pedestal. It com-mands modulation of the outflow valve. A dual-scale indicator dial is mounted in the center ofthe pressurization controller. The outer scale(CABIN ALT) indicates the cabin pressure alti-tude which the pressurization controller is set tomaintain. The inner scale (ACFT ALT) indicatesthe maximum ambient pressure altitude at whichthe airplane can fly without causing the cabinpressure altitude to climb above the valueselected on the outer scale (CABIN ALT) of thedial. The indicated value on each scale is readopposite the index mark at the forward (top)position of the dial. Both scales rotate togetherwhen the cabin altitude selector knob, placardedCABIN ALT is turned.

Cabin altitude is obtained by setting the control-ler to the desired cruising altitude, and observingthe cabin altitude on the scale. The maximumcabin altitude selected may be anywhere from–1,000 to +10,000 feet MSL. The rate controlselector knob is placarded RATE – MIN – MAX.The rate at which the cabin pressure altitudechanges from the current value to the selected

value is controlled by rotating the rate controlselector knob. The rate of change selected maybe from approximately 200 to approximately2,000 feet per minute. Normal setting on the rateknob will be from 9 o’clock to 12 o’clock.

The actual cabin pressure altitude (outer scale)and cabin differential (inner scale) is continu-ously indicated by the cabin altimeter (Figure12-9), which is mounted in the right side of the

Figure 12-8 Pressurization Controller

Figure 12-9 Cabin Altimeter


12-9

panel that is located above the pedestal. Immedi-ately to the left of the cabin altimeter is the cabinvertical speed (CABIN CLIMB) indicator (Fig-ure 12-10), which continuously indicates the rateat which the cabin pressure altitude is changing.

The cabin pressure switch (Figure 12-11), to theleft of the pressurization controller on the pedes-tal, is placarded CABIN PRESS – DUMP –PRESS – TEST. When this switch is in theDUMP (forward lever locked) position, thesafety valve is held open, so that the cabin willdepressurize and/or remain unpressurized. Whenit is in the PRESS (center) position, the safetyvalve is normally closed in flight, and the outflowvalve is controlled by the pressurization control-

ler, so that the cabin will pressurize. When theswitch is held in the spring-loaded TEST (aft)position, the safety valve is held closed, bypass-ing the landing gear safety switch, to facilitatetesting of the pressurization system on theground. Circuit breakers for the system (Figure12-12) are on the copilot’s side panel under theheading ENVIRONMENTAL.

Figure 12-10 Cabin Climb Indicator

Figure 12-11 Cabin Pressure Switch

Figure 12-12 Environmental System Circuit Breakers

12-10


PREFLIGHT CHECK

During runup, the pressurization system may befunctionally checked using the cabin pressuriza-tion switch. With both bleed-air valves OPEN,adjust the cabin altitude selector knob so that theCABIN ALT dial indicates an altitude 500 feetBELOW field pressure altitude. Rotate the ratecontrol selector knob to place the index betweenthe 9 and 12 o’clock positions. Move both condi-t ion levers to h igh id le . Hold the cabinpressurization switch to the TEST position andcheck the CABIN CLIMB indicator for a descentindication. Release the pressurization switch tothe PRESS position when pressurizing is con-firmed and move both condition levers to theiroriginal position.

Prior to takeoff, the CABIN ALT selector knobshould be adjusted so that the ACFT ALT scaleon the indicator dial indicates an altitude approx-imately 500 feet above the planned cruisepressure altitude prior to takeoff. The rate controlselector knob should be adjusted as desired; set-ting the index mark between the 9 and 12 o’clockpositions will provide the most comfortablecabin rate of climb. The cabin pressure switchshould be checked to ensure that it is the PRESSposition.

IN FLIGHT

As 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 alti-tude at the selected value. If the airplane climbsto an altitude higher than the value indexed onthe ACFT ALT scale of the dial on the face of thecontroller, the pressure differential will reach thepressure relief setting of the outflow valve andsafety valve. Either or both valves will then over-ride the cabin pressurization controller in order tolimit the pressure differential to the maximumpressure differential. If the cabin pressure alti-tude should reach a value of 10,000 feet (12,500for LJ 1353 and later), a pressure-sensing switchwill close. This causes the red ALTITUDEWARN annunciator light to illuminate, warning

the pilot of operation requiring the use of oxy-gen. During cruise operation, if the flight plancalls for an altitude change of 1,000 feet or more,reselect the new altitude plus 500 feet on theCABIN ALT dial.

DESCENT

During descent and in preparation for landing,set the cabin altitude selector to indicate a cabinaltitude of approximately 500 feet above thelanding field pressure altitude (Figure 12-13),and adjust the rate control selector as required toprovide a comfortable cabin-altitude rate ofdescent. Control the airplane rate of descent sothat the airplane altitude does not catch up withthe cabin pressure altitude until the cabin pres-sure altitude reaches the selected value, whichmay happen before the airplane reaches theselected altitude. Then as the airplane descendsto and reaches the cabin pressure altitude thenegative pressure relief function opens the out-

Figure 12-13 Pressurization Controller Setting for Landing


12-11

flow and safety valve poppets toward the fullyopen position, thereby equalizing the pressureinside and outside the pressure vessel. As the air-plane continues to descend below the preselectedcabin pressure altitude, the cabin will be unpres-surized and will follow the airplane rate ofdescent to touchdown.

FLOW CONTROL UNIT

A flow control unit, mounted in each nacelle onthe forward side of the firewall, controls thebleed air from the engine for use in pressuriza-tion, heating, and ventilation. The function of theflow control unit (Figure 12-14) is to vary theflow and balance of bleed air and ambient air tothe cabin pressure vessel. This is done by meansof temperature and pressure sensors and theirrelated modulating valves.

When the BLEED AIR switches on copilot’s leftsubpanel are OPEN a bleed-air shutoff electricsolenoid valve on each flow control unit opens toallow the bleed air into the unit. The flow controlunit will then adjust the flow of bleed air mixedwith ambient air into the pressure vessel. Ambi-ent air is allowed to enter the flow control unitthrough a normally-open modulating valve, andserves to add air mass and some cooling to thebleed air flow.

The ambient air valve, associated with the tem-perature sensing device, is also controlled by theleft landing gear safety switch. When the aircraftis on the ground, the valve is directed to shut offthe ambient air source from the flow controlvalve. The exclusion of ambient air allows fastercabin warm-up during cold weather operation.

Figure 12-14 Flow Control Unit

12-12


After takeoff, the landing gear safety switch sig-nals the ambient air modulating valves to open.They do so sequentially to prevent the simulta-neous opening of the modulating valves and asudden pressure surge into the cabin.

The pneumostat (pneumatic thermostat) providestemperature input to the flow control unit, whichmodulates the amount of ambient air entering theflow unit for blending. Warmer outside air opensthe modulating valve and allows more ambientair in for blending. Cold air closes the valve untilit closes completely at a preset temperature. Atthis point, bleed air will be providing all air forpressurization. A check valve prevents air fromleaking out the ambient air input.

An aneroid near the bleed air ejector flow controlactuator influences the amount of bleed air enter-ing the flow control unit. The aneroid providesaltitude sensing information to the flow controlunit, and combined with the pneumostat, pro-vides accurate bleed-air input into the pressurevessel.

The quantity of bleed-air flow into the pressurevessel is influenced directly by ambient tempera-ture and ambient pressure.


13-1

See Chapter 14, “Landing Gear and Brakes,” for information onthe hydraulic power systems.


14-i

CHAPTER 14

CONTENTS

Page

LANDING GEAR AND BRAKES

INTRODUCTION ................................................................................................................

14-1

GENERAL ............................................................................................................................

14-1

LANDING GEAR SYSTEM................................................................................................

14-2

Landing Gear Assemblies .............................................................................................

14-2

Wheel Well Door Mechanisms .....................................................................................

14-3

Steering..........................................................................................................................

14-3

Hydraulic Landing Gear ................................................................................................

14-4

Landing Gear Extension and Retraction .......................................................................

14-5

Hydraulic Fluid Level Indication System .....................................................................

14-8

Landing Gear Warning System ...................................................................................

14-11

Manual Landing Gear Extension.................................................................................

14-12

Hydraulic Schematics..................................................................................................

14-12

Tires.............................................................................................................................

14-17

Shock Struts.................................................................................................................

14-17

Landing Gear Operating Limits ..................................................................................

14-17

KING AIR WHEEL BRAKES ...........................................................................................

14-19

Series Brake System....................................................................................................

14-19

Parking Brake ..............................................................................................................

14-19

Brake Service ..............................................................................................................

14-22

Brake Wear Limits ......................................................................................................

14-22

Cold Weather Operation..............................................................................................

14-23


14-iii

Figure Title Page

ILLUSTRATIONS

14-1

Main Gear Assembly .............................................................................................

14-2

14-2

Nose Gear Assembly .............................................................................................

14-2

14-3

Main Gear Door Mechanism .................................................................................

14-3

14-4

Landing Gear Electrical Schematic .......................................................................

14-4

14-5

Hydraulic Landing Gear Plumbing Schematic ......................................................

14-5

14-6

Hydraulic Landing Gear Diagram .........................................................................

14-6

14-7

Hydraulic Landing Gear Power Pack ....................................................................

14-7

14-8

Landing Gear Control Switch Handle ...................................................................

14-8

14-9

Hydraulic Fluid Low Indicator ..............................................................................

14-8

14-10

Safety Switch .........................................................................................................

14-9

14-11

Gear Position Indicator ..........................................................................................

14-9

14-12

Gear Position Indicator—No Illumination ..........................................................

14-10

14-13

Landing Gear Control Switch Handle—Red In-Transit Indicators .....................

14-10

14-14

Handle Light Test ................................................................................................

14-10

14-15

Landing Gear Alternate Extension Placard .........................................................

14-12

14-16

Landing Gear Relay Circuit Breaker ...................................................................

14-12

14-17

Landing Gear Retraction Schematic ....................................................................

14-14

14-18

Landing Gear Extension Schematic.....................................................................

14-15

14-19

Hand Pump Emergency Extension Schematic ....................................................

14-16

14-20

Landing Gear Maintenance Retraction Schematic ..............................................

14-18

14-21

Brake System Schematic .....................................................................................

14-20

14-22

Parking Brake Schematic.....................................................................................

14-21

14-iv


14-23

Brake Fluid Reservoir .........................................................................................

14-22

14-24

Brake Wear Diagram...........................................................................................

14-23


14-v

Table Title Page

TABLES

14-1

Landing Gear Warning Horn Operation ..............................................................

14-11

14-2

Landing Gear Operating Limits...........................................................................

14-17


14-1

CHAPTER 14LANDING GEAR AND BRAKES

INTRODUCTION

An understanding of the landing gear system will aid the pilot in proper handling of landinggear operation and emergency procedures. This chapter, in addition to describing the system,identifies inspection points and abnormal conditions to be considered. This chapter alsoincludes brakes, since an understanding of the brake system will help the pilot operate thebrakes safely and with minimum wear. In addition to system description, operating andservicing procedures are covered.

GENERAL

This chapter presents a description and discus-sion of the landing gear system, landing gearcontrols, and limits. The indicator system andemergency landing gear extension are alsodescribed.

This chapter also presents a description and dis-cussion of the wheel brake system. Correct use ofthe brakes and parking brakes, brake systemdescription, and what to look for when inspectingbrakes are also detailed.

14-2


LANDING GEAR SYSTEM

LANDING GEAR ASSEMBLIES

Components

Each landing gear assembly (main and nose)consists of a shock strut, torque knee (scissors),drag leg, actuator, wheel, and tire. Brake assem-blies are located on the main gear assemblies; theshimmy damper is mounted on the nose gearassembly (Figures 14-1 and 14-2).

Operation

The upper end of the drag legs and two points onthe shock struts are attached to the airplane struc-

ture. When the gear is extended, the drag bracesare rigid components of the gear assemblies.

The landing gear incorporates Beech air/oilshock struts that are filled with both compressedair and hydraulic fluid. Airplane weight is borneby the air charge in the shock struts. At touch-down, the lower portion of each strut is forcedinto the upper cylinder; this moves fluid throughan orifice, further compressing the air charge andthus absorbing landing shock. Orifice action alsoreduces bounce during landing. At takeoff, thelower portion of the strut extends until an internalstop engages.

A torque knee connects the upper and lower por-t ions of the shock s t ru t . I t a l lows s t ru tcompression and extension but resists rotationalforces, thereby keeping the wheels aligned withthe longitudinal axis of the airplane. On the nosegear assembly, the torque knee also transmits

Figure 14-1 Main Gear Assembly

Figure 14-2 Nose Gear Assembly


14-3

steering motion to the nosewheel, and nosewheelshimmy motion to the shimmy damper.

The shimmy damper, mounted on the right sideof the nose gear strut, is a balanced hydraulic cyl-inder that bleeds fluid through an orifice todampen nosewheel shimmy.

WHEEL WELL DOOR MECHANISMS

The landing gear doors consist of one set of nosegear doors and two sets of main gear doors.Landing gear doors are mechanically actuated bygear movement during extension and retraction.

The nose gear doors are hinged at the sides andare spring-loaded to the open position. As thelanding gear is retracted, a roller on each side ofthe nose gear assembly engages a cam assemblyon each door, and draws the doors closed behindthe gear. The reverse action takes place, andspring-loading takes effect as the nose gear isextended.

The main gear doors are hinged at the sides andare connected to a landing-gear, door-actuatortorque tube assembly (Figure 14-3) with twopush-pull links. The torque tube assembly alsocontains an uplock roller support assemblywhich, when contacted by the uplock cam on themain gear shock cylinder, rotates the torque tubeto pull the doors closed upon gear retraction, orpush the doors open upon gear extension.

Roller movement is transmitted through linkageto close the doors. During extension, roller actionreverses cam movement to open the doors. Whenthe cam has left the roller, springs pull the link-age over-center to hold the doors open.

STEERING

Direct linkage to the rudder pedals permits nose-wheel steering when the nose gear is down. Onespring-loaded link in the system absorbs some ofthe force applied to any of the interconnectedrudder pedals until the nosewheel is rolling. Atthis time the resisting force is less, and more

Figure 14-3 Main Gear Door Mechanism

14-4


pedal motion results in more nosewheel deflec-tion. Since motion of the pedals is transmitted viacables and linkage to the rudder, rudder deflec-tion occurs when force is applied to the rudderpedals. With the nose landing gear retracted,some of the force applied to any of the rudderpedals is absorbed by the spring-loaded link inthe steering system, so that there is no motion atthe nosewheel but rudder deflection still occurs.The nosewheel is self-centering upon retraction.

When force on the rudder pedal is augmented bya main wheel braking action, the nosewheeldeflection can be considerably increased.

HYDRAULIC LANDING GEAR

The retractable tricycle landing gear is electri-cally controlled (Figure 14-4) and hydraulicallyactuated. The system utilizes folding braces,called “drag legs,” that lock in place when thegear is fully extended.

The individual landing gear actuators incorporateinternal/mechanical downlocks to hold the gearin the fully extended position. The landing gear isheld in the up position by hydraulic pressure.

Figure 14-4 Landing Gear Electrical Schematic


14-5

Hydraulic pressure to the system is supplied by ahydraulic power pack (Figure 14-5). A hydraulicreservoir located in the left center wing sectionprovides hydraulic fluid to the power pack. Thereservoir incorporates a dipstick to provide avisual check of fluid level.

An electrically actuated selector valve controlsthe flow of hydraulic fluid to the individual gearactuators. The selector valve receives electricalpower through the landing gear control switch.

Accidental retraction of the landing gear is pre-vented through safety switches located on themain landing gears.

LANDING GEAR EXTENSION AND RETRACTION

The nose and main landing gear assemblies areextended and retracted by a hydraulic power packin conjunction with hydraulic actuators (Figure14-6). The hydraulic power pack is located in thecenter of the center section, just forward of themain spar. One hydraulic actuator is located ateach landing gear.

The power pack (Figure 14-7) consists of: ahydraulic pump, a 28-VDC motor, a two-sectionfluid reservoir, filter screens, a four-way gearselector valve, an up selector solenoid, a fluid

Figure 14-5 Hydraulic Landing Gear Plumbing Schematic

14-6


level sensor, and an uplock pressure switch. Formanual extension the system has a hand-lever-operated pump. The pump handle is located onthe floor between the pilot’s seat and thepedestal.

Three hydraulic lines (one for normal extensionand one for retraction, routed from the powerpack, and one for emergency extension routedfrom the hand pump) are routed to the nose andmain gear actuators. The normal extension linesand the manual extension lines are connected tothe upper end of each hydraulic actuator. Thehydraulic lines for retraction are fitted to the

lower ends of the actuators. Hydraulic fluid underpressure (generated by the power pack pump andcontained in the accumulator) acts on the pistonfaces of the actuators (which are attached to fold-ing drag braces), resulting in the extension orretraction of the landing gear.

When the actuator pistons are repositioned tofully extend the landing gear, an internalmechanical lock in the nose gear actuator and theover-center action of the nose gear drag legassembly lock the nose gear in the down position.In this position, the internal locking mechanismin the nose gear actuator will actuate the actuator

Figure 14-6 Hydraulic Landing Gear Diagram


14-7

downlock switch to interrupt current to the pumpmotor. The motor will continue to run until allthree landing gears are down and locked. Aspring-loaded downlock assembly is fitted toeach main gear upper drag leg, providing positivedownlock action for the main gear.

In flight, with the LDG GEAR CONTROL in theDN position, as the landing gear moves to thefully down position, the downlock switches areactuated, and they cause the landing gear relay tointerrupt current to the pump motor. When thered GEAR-IN-TRANSIT lights in the LDG

GEAR CONTROL switch handle extinguish, andthe green NOSE–L–R indicators illuminate, thelanding gear is in the fully down-and-lockedposition.

A solenoid mounted on the valve body end of thepump is energized when the LDG GEAR CON-TROL is in the UP position and actuates the gearselect valve, allowing system fluid to flow to theretract side of the system. The gear select valve isspring-loaded in the down position and will moveto the up position only when energized. The nosegear actuator will unlock when 200 to 400 psi of

Figure 14-7 Hydraulic Landing Gear Power Pack

14-8


hydraulic pressure is applied to the retract port ofthe nose gear actuator. The landing gear willbegin to retract after the nose gear actuator isunlocked.

Hydraulic system pressure performs the uplockfunction, holding the landing gear in the retractedposition. When the hydraulic pressure reachesapproximately 1,850 psi, the uplock pressureswitch will cause the landing gear relay to openand interrupt the current to the pump motor. Thesame pressure switch will cause the pump toactuate should the hydraulic pressure drop toapproximately 1,600 psi.

The landing gear control circuit is protected by a2-ampere circuit breaker located on the pilot’sinboard subpanel. Power for the pump motor issupplied through the landing gear motor relayand a 60-ampere circuit breaker, both of whichare located under the cabin floor in the wing cen-ter section. The motor relay is energized bycurrent from the 2-ampere circuit breaker and thedownlock switches.

HYDRAULIC FLUID LEVEL INDICATION SYSTEM

A caution annunciator placarded “HYD FLUIDLOW” (Figure 14-9), in the annunciator panel,will illuminate (yellow) whenever the hydraulicfluid level in the landing gear power pack reser-voir is low. The annunciator is tested by pressingthe HYD FLUID SENSOR TEST button locatedon the pilot’s subpanel.

If the HYD FLD LOW annunciator comes on,normal extension may be attempted, but the pilotshould be prepared for an emergency manualextension.

Figure 14-8 Landing Gear Control Switch Handle

Figure 14-9 Hydraulic Fluid Indicator


14-9

Control

The landing gear hydraulic power pack motor iscontrolled by the landing gear switch handleplacarded “LDG GEAR CONTROL” with UPand DN positions, located on the pilot’s rightsubpanel (Figure 14-9). The switch handle mustbe pulled out of a detent before it can be movedfrom either the UP or DN position.

Safety switches (Figure 14-10) called “squat”switches, on the main gear shock strut, open thecontrol circuit when the oleo strut is compressed.The squat switches must close to actuate a sole-noid, which moves a downlock hook on the LDGGEAR CONTROL switch to the released posi-tion. This mechanism prevents the LDG GEARCONTROL switch handle from being placed inthe UP position when the airplane is on theground. The downlock hook automaticallyunlocks when the airplane leaves the ground.

The downlock hook disengages when the air-plane leaves the ground because the squatswitches close and a circuit is completed throughthe solenoid that moves the hook. In the event ofa malfunction of the downlock solenoid or thesquat switch circuit, the downlock hook can beoverridden by pressing downward on the red

DOWN LOCK REL button. The release button islocated just left of the LDG GEAR CONTROLswitch handle.

The LDG GEAR CONTROL handle shouldnever be moved out of the DN detent while theairplane is on the ground. If it is, the landing gearwarning horn will sound intermittently, and thered gear-in-transit lights in the LDG GEARCONTROL switch handle will illuminate (pro-vided the MASTER SWITCH is ON), warningthe pilot to return the handle to the DN position.

Position Indicators

Landing gear position is indicated by an assem-bly of three lights in a single unit located on thepilot’s right subpanel (Figure 14-11). The unithas a light transmitting cap that is marked as fol-lows: “NOSE–L–R.” Light bulbs in eachsegment, when illuminated, make the segmentappear green and indicate that particular gear isdown and locked. Absence of illumination mayindicate an unsafe gear indication (Figure 14-12).The green position indicator lights may bechecked by pushing on the light housing.

Figure 14-10 Safety Switch

Figure 14-11 Gear Position Indicator

14-10


Two red parallel-wired indicator lights, located inthe LDG GEAR CONTROL switch handle, illu-minate to show that the gear is in-transit (Figure14-13) or unlocked. Gear UP is indicated whenthe red lights go out. The red lights in the handlealso illuminate when the landing gear warningsystem is activated.

The red control handle lights may be checked bypressing HD LT TEST button (Figure 14-14)located adjacent to the LDG GEAR CONTROLswitch handle.

Each normally closed, up-position switch islocated in the upper portion of its respectivewheel well. When the gear is in the fullyretracted position, each strut actuates its respec-tive up-position switch to open the circuit fromthe in-transit light to ground. As soon as the gearmoves from the fully retracted position, eachstrut actuates its respective up-position switch toilluminate the in-transit light by providing a pathto ground through the down-position switch. Thein-transit light goes out when the drag brace ineach landing gear passes over-center to actuateits respective down-position switch to themomentary contacts. In this position, the switchopens the circuit to the in-transit light and com-pletes a path to ground for the down-positionlights. The down-position switch on each landinggear also functions as a warning switch for thesystem.

Figure 14-12 Gear Position Indicator—No Illumination

Figure 14-13 Landing Gear Control Switch Handle—Red In-Transit Indicators

Figure 14-14 Handle Light Test


14-11

The landing gear in-transit light will indicate oneor all of the following conditions:

●

Landing gear handle is in the UP posi-tion, and the airplane is on the groundwith weight on the landing gear.

●

With flaps up or approach and one orboth power levers retarded below approx-imately 79 ±2% N

1

, one or more landinggears are not down and locked.

●

Any landing gear is not in the fullyretracted position.

●

Flaps are beyond the APPROACH posi-tion (36% or more) with any gear notdown, regardless of power lever position.

Thus, the function of the landing gear in-transitlight is to indicate that the landing gear is intransit.

The up indicator, down indicator, and warninghorn systems are essentially independent sys-tems. A malfunction in any one system willprobably leave the other two systems unaffected.

LANDING GEAR WARNING SYSTEM

The landing gear warning system is provided towarn the pilot that the landing gear is not downand locked during specific flight regimes. Variouswarning modes result, depending upon the posi-tion of the flaps.

With the flaps in the UP or APPROACH positionand either or both power levers retarded belowabout 79% N

1

, the warning horn will sound inter-mittently. The horn can be silenced by pressingthe GEAR WARN SILENCE button adjacent tothe LDG GEAR CONTROL switch handle. Onthe C90B, the warning horn is silenced by press-ing the silence button located on the left powerlever. The landing gear warning system will berearmed if the power levers are advancedsufficiently.

With the FLAPS beyond the APPROACH posi-tion, the warning horn activates regardless of thepower lever settings and cannot be canceled.

Landing gear warning horn operation is shown inTable 14-1 below.

Table 14-1 LANDING GEAR WARNING HORN OPERATION

GEAR POSITION FLAPS POWER HORN SILENCE MODE

Up Up +77 to 81% No N/A

Up Up –77 to 81% Yes Silence button

Up Approach –77 to 81% Yes Silence button

Up Past approach Any Yes Lower gear

14-12


MANUAL LANDING GEAR EXTENSION

A hand pump handle, placarded “LANDINGGEAR ALTERNATE EXTENSION” (Figure14-15), is located on the floor between the pilot’sseat and the pedestal. The pump is located underthe floor, below the handle, and is used whenemergency extension of the gear is required.

To engage the system, pull the LANDINGGEAR RELAY circuit breaker (Figure 14-16),located below and to the left of the LDG GEARCONTROL switch handle on the pilot’s sub-pane l , and ensu re tha t t he LDG GEARCONTROL handle is in the DN position.Remove the pump handle from the securing clip,and pump the handle up and down until the greenNOSE–L–R gear-down indicator lights illumi-nate and further resistance is felt. Place thehandle in the fully down position and secure inthe retaining clip.

WARNING

If for any reason the green GEARDOWN lights do not illuminate (e.g.,in case of an electrical system failureor in the event an actuator is not locked

“down”), continue pumping until suffi

-

cient resistance is felt to ensure thatthe gear is down and locked. Do notstow pump handle. The landing gearcannot be manually retracted in flight.

WARNING

After a manual landing gear extensionhas been made, do not move any land-ing gear controls or reset any switchesor circuit breakers until the airplane ison jacks.

After a practice manual extension of the landinggear, the gear may be retracted hydraulically bypushing the LANDING GEAR RELAY circuitbreaker in and moving the LDG GEAR CON-TROL handle to the UP position.

HYDRAULIC SCHEMATICS

The hydraulic gear schematics shown are for thegear extended, gear retracted, hand pump emer-gency extension, and gear maintenance retractionmodes. Power is available to the contacts of thelanding gear remote power relay.

Figure 14-15 Landing Gear Alternate Extension Placard

Figure 14-16 Landing Gear Relay Circuit Breaker


14-13

When the relay is open, power comes down fromthe 2-amp gear control circuit breaker to thelanding gear control assembly switch and on tothe three downlock switches. Each gear is downand locked, so these three switches are open andno circuit passes through them. This is the staticcondition of the system after a normal gearextension.

Landing Gear Retraction

When the aircraft is airborne, the pilot selectsGEAR UP (Figure 14-17). Circuits are madefrom the gear selector switch to the uplock pres-sure switch. The pressure switch is closed at thistime, so the circuit is complete to the gear upmain switch and landing gear remote powerrelay. This relay now closes and provides thepower circuit to the hydraulic pump motor. Back-ing up to the pressure switch, a circuit is made tothe hydraulic selector valve up-solenoid. Powerto this solenoid will position the selector valvebody to route hydraulic fluid in the appropriatedirection to retract the gear.

After approximately six seconds the retractioncycle is complete. Once the landing gear reachesfull-up travel, each actuator physically bottomsout. The pressure on the retract line builds rap-idly until pressure reaches approximately 1,850psi. The uplock pressure switch opens at thistime, breaking the power circuit to the pumpmotor and stopping the hydraulic pump. Thispressure switch will close periodically whenpressure drops to approximately 1,600 psi, due tothe normal system pressure leak-down, and re-energize the pump to restore needed uplock pres-sure. Consequently, when the gear is retracted,pressure will be maintained between approxi-mately 1,600 and 1,850 psi to keep the gears intheir retracted position. An accumulator pre-charged to 800 psi, located in the left winginboard of the nacelle, is designed to aid in main-taining the system pressure in the gear-up mode.

Landing Gear Extension

For normal gear extension, a pilot selects GEARDOWN (Figure 14-18), and circuits are madefrom the landing gear control assembly through

any one of the three actuator downlock switches,back through the landing gear control assembly,the service valve, and finally to the landing gearremote power relay. The power relay closes andprovides a power circuit to the pump motor. Theselector valve is not being powered at this time.Thus, fluid under pump pressure is routedthrough the selector valve body in the appropri-ate direction to extend the landing gear.

The gear comes down under fluid pressure untileach main gear downlock and the nose gear actu-ator downlock switches are depressed. When allthree gears are down and locked, the control cir-cuit to the pump motor is broken, and the pumpstops. Notice that no pressure switches areinvolved. Consequently, there is no downlockpressure maintained. The mechanical downlockson each main gear drag brace, and an internalmechanical lock in the nose gear actuator, pre-vent gear retraction.

Hand Pump Emergency Extension

A hand-pump handle, placarded “LANDINGGEAR ALTERNATE EXTENSION,” is locatedon the floor between the pilot’s seat and the ped-estal. The pump is located under the floor belowthe handle and is used when emergency exten-sion of the gear is required.

To engage the system, pull the LANDINGGEAR RELAY circuit breaker, located on thepilot’s inboard subpanel, and place the LDGGEAR CONTROL switch handle in the DNposition (Figure 14-19). Remove the pump han-dle from the securing clip, and pump the handleup and down until the green NOSE–L–R geardown indicator lights illuminate. Place the pumphandle in the fully down position and secure inthe retaining clip.

After a practice manual extension of the landinggear, the gear may be retracted hydraulically bypushing the LANDING GEAR RELAY circuitbreaker in and moving the LDG GEAR CON-TROL switch handle to the UP position.

14-14


Figure 14-17 Landing Gear Retraction Schematic


14-15

Figure 14-18 Landing Gear Extension Schematic

14-16


Figure 14-19 Hand Pump Emergency Extension Schematic


14-17

If an alternate landing gear extension becomesnecessary, there is no limit to the amount ofcycles the hydraulic gear may be pumped. Dur-ing a complete or partial electrical failure, thegear down lights, in-transit lights, and gear warn-ing horn may not be operating. A positive methodof checking that the gear is down is throughresistance when pumping the extension handle.When all three gears are extended, hydraulicpressure is built up until the pressure relief valveopens, relieving the pressure built up by the han-dle. This can be felt by the pilot as increasedresistance while pumping, followed by a give asthe relief valve opens.

Landing Gear Maintenance Retraction

A service valve (Figure 14-20), located forwardof the power pack assembly, may be used in con-junction with the hand pump to raise the gear formaintenance purposes. With the aircraft on jacksand an external electrical power source attached,unlatch the hinged retainer and pull up on the redknob located on top of the service valve. Thehand pump can then be used to raise the gear tothe desired position. After the required mainte-nance has been performed, push the red knobdown, and use the hand pump to lower the gear.The valve is not accessible to the pilot.

CAUTION

If the red knob on the service valve ispushed down while the landing gear isretracted, the electrical power on, and

the landing gear control handle is inthe down position, the landing gearwill extend immediately.

A fill reservoir, located just inboard of the leftnacelle and forward of the front spar, contains acap and dipstick assembly to facilitate mainte-nance of the system fluid level. A line plumbed tothe upper portion of the fill reservoir is routedoverboard to act as a vent.

TIRES

The nose landing gear wheel is equipped with a6.50 x 10, 6-ply-rated, tubeless, rim-inflation tire.Each main landing gear wheel is equipped withan 8.50 x 10, 8-ply-rated, tubeless, rim-inflationtire. For increased service life, 10-ply-rated tiresof the same size may be installed. Check the


for correct tirepressure.

SHOCK STRUTS

Shock struts should always be properly inflated.Do not over- or under-inflate, and never tow ortaxi an aircraft when any strut is flat. Correctinflation is approximately 3 inches for the mainstrut and 3.0 to 3.5 inches for the nose strut.

LANDING GEAR OPERATING LIMITS

The landing gear operating limits are shown inTable 14-2 below.

Table 14-2 LANDING GEAR OPERATING LIMITS

AIRSPEED KIAS REMARKS

Maximum landing gear operation (V

LO

)• Extension• Retraction

182163

Do not extend or retract the landing gear above this speed.

Maximum Landing gear extended (V

LE

) 182 Do not exceed this speed with the landing gear extended.

14-18


Figure 14-20 Landing Gear Maintenance Retraction Schematic


14-19

KING AIR WHEEL BRAKES

The King Air series brakes are a non-assistedhydraulic brake system. The main landing gearwheels are equipped with multi-disc dual hydrau-lic brakes. These brakes are actuated by toepressure on the rudder pedals by either the pilotor copilot. The depression of either set of pedalscompresses the piston rod in the master cylinderattached to each pedal. The hydraulic pressureresulting from the movement of the pistons in themaster cylinders is transmitted through flexiblehoses and fixed aluminum tubing to the discbrake assemblies on the main landing gearwheels. This pressure forces the brake pistons onthe wheel to press against the multiple liningsand discs of the brake assembly.

As with any airplane, proper traction and brakingcontrol cannot be expected until the landing gearis carrying the full weight of the airplane. Useextreme care when braking to prevent skiddingand the resulting flat sections on tires caused byskidding. Braking should be smooth and even allthe way to the end of ground roll.

SERIES BRAKE SYSTEM

The dual brakes are plumbed in series (Figure14-21). Each rudder pedal is attached to its ownmaster cylinder. The pilot’s master cylinders areplumbed through the copilot’s master cylinders,thus allowing either set of pedals to perform thebraking action. The pilot’s and copilot’s rightrudder pedals control the brake in the right mainlanding gear. Similarly, the pilot’s and copilot’sleft rudder pedals control braking in the left maingear. This arrangement allows differential brak-ing for taxiing and maneuvering on the ground.

PARKING BRAKE

The parking brake utilizes the regular brakes anda set of valves (Figure 14-22). Dual parkingbrake valves are installed adjacent to the rudderpedals between the master cylinders of the copi-lot’s rudder pedals and the wheel brakes. The twolever-type valves are located just aft of the flightcompartment under the center aisle floorboard. Apush-pull cable from the valve control levers runsto the pedestal, terminating with a knob. The con-trol knob for the parking brake valves, placarded“PARKING BRAKE–PULL ON,” is below thelower left corner of the pilot’s subpanel.

To set the parking brake: depress the brake pedalsto build up pressure in the brake system, thendepress the button in the center of the parkingbrake control, and pull the control handle aft orON. This procedure closes both parking brakevalves simultaneously. The parking brake valvesshould retain the pressure previously pumpedinto the system.

The parking brake can be released from either thepilot’s or copilot’s side when the brake pedals aredepressed briefly to equalize the pressure on bothsides of the valves, and the PARKING BRAKEhandle is pushed in to allow the parking brakevalves to open.

To avoid damage to the parking brake system,tires, and landing gear, the parking brake shouldbe left off and wheel chocks or tiedowns installedif the airplane is to be left unattended, becausethe airplane may be moved by ground personnelin the pilot’s absence. Also, ambient temperaturechanges can expand or contract the brake fluid,causing excessive brake pressure or brakerelease.

14-20


Figure 14-21 Brake System Schematic


14-21

Figure 14-22 Parking Brake Schematic

14-22


BRAKE SERVICE

Brake fluid is supplied to the master cylindersfrom a reservoir accessible through the nose avi-onics compartment door, prior to LJ-1531. OnLJ-1531 and subsequent the door was replacedwith an access panel (Figure 14-23). The brakefluid reservoir is located on the upper corner ofthe left side of the nose avionics compartment.

Brake system servicing is limited primarily tomaintaining the hydraulic fluid level in the reser-voir. A dipstick is provided for measuring thefluid level. When the reservoir is low on fluid,add a sufficient quantity of MIL-H-5606 hydrau-lic fluid to fill the reservoir to the full mark on thedipstick. Check all hydraulic landing gear con-nections for signs of seepage and correct ifnecessary. Do not check while the parking brakeis deployed.

Standard brakes used on this airplane areequipped with automatic brake adjusters. Theautomatic brake adjusters reduce brake drag,thereby allowing unhampered roll. Airplaneswith the automatic adjusters tend to exhibit asofter pedal and a somewhat longer pedal stroke.

BRAKE WEAR LIMITS

Brake lining adjustment is automatic, eliminatingthe need for periodic adjustment of the brakeclearance. Check brake wear periodically toassure that dimension “A,” in the Brake WearDiagram (Figure 14-24), does not reach zero.When it reaches zero, refer to the Beechcraft ser-vicing and maintenance instructions for King Airbrakes and wheels. The parking brake must be set(pressure on the brakes) before this can be done.

Figure 14-23 Brake Fluid Reservoir


14-23

COLD WEATHER OPERATIONWhen operating in cold weather, check thebrakes and the tire-to-ground contact for freezelock-up. Anti-ice solutions may be used on thebrakes or tires if freeze-up occurs. No anti-icesolution which contains a lubricant, such as oil,should be used on the brakes. It will decrease theeffectiveness of the brake friction areas.

When possible, taxiing in deep snow or slushshould be avoided. Under these conditions thesnow and slush can be forced into the brakeassemblies. Keep flaps retracted during taxiing toavoid throwing snow or slush into the flap mech-anisms and to minimize damage to flap surfaces.

Figure 14-24 Brake Wear Diagram


15-i

CHAPTER 15

CONTENTS

Page

FLIGHT CONTROLS

INTRODUCTION ................................................................................................................

15-1

DESCRIPTION.....................................................................................................................

15-1

FLAPS SYSTEM..................................................................................................................

15-2

C90A Flap Operation ....................................................................................................

15-3

C90B Flap Operation.....................................................................................................

15-3

Landing Gear Warning System .....................................................................................

15-4

Flap Airspeed Limits .....................................................................................................

15-4

RUDDER BOOST SYSTEM ...............................................................................................

15-4


15-iii

Figure Title Page

ILLUSTRATIONS

15-1

Flap Control System ..............................................................................................

15-2

15-2

Flap Control Lever.................................................................................................

15-3

15-3

Flap Position Indicator...........................................................................................

15-3

15-4

Flap System Circuit Breaker..................................................................................

15-3

15-5

Airspeed Indicator .................................................................................................

15-4

15-6

Rudder Boost System Diagram .............................................................................

15-5

15-7

Rudder Boost Switch .............................................................................................

15-6


15-1

CHAPTER 15FLIGHT CONTROLS

INTRODUCTION

Familiarization with the flap system operation and limits is necessary to provide optimumperformance in takeoff, approach, and landing modes. This chapter identifies and describes flapaction so the pilot will understand their operation, controls, and limits.

A basic understanding of how the rudder boost system works, and its value in engine-out situa-tions, will assist the pilot in making full use of its advantages. This chapter also presentsfamiliarization with and operation of the rudder boost system.

DESCRIPTION

This chapter presents a description and discus-sion of flap system. The four-segment Fowler-type system, its controls and limits are consid-ered with reference to operation as outlined in the


.

The rudder boost system section of this chapterpresents a description and discussion of the rud-der boost system. This system is designed toreduce pilot effort in single-engine flightconfigurations.

��

��

��

��

20

20 20

105

510 10

5

5

LO

C

GS

15-2


FLAPS SYSTEM

The flaps, two panels on each wing, are driven byan electric motor through a gearbox mounted onthe forward side of the rear spar (Figure 15-1).The motor incorporates a dynamic braking sys-tem through the use of two sets of motorwindings. This system helps to prevent overtravelof the flaps. The gearbox drives four flexibledriveshafts, each of which is connected to a jack-screw actuator at each flap.

The flaps are operated by a sliding lever locatedjust below the condition levers on the pedestal(Figure 15-2). Flap travel, from 0% (UP) to100% (DOWN), is registered at 20, APPROACH,40, 60, and 80 and DOWN in percentage of travelon an electric indicator on top of the pedestal(Figure 15-3).

The flap control has a position detent providedfor quick selection of 30% (15º) flaps forAPPROACH. Full flap deflection is approxi-mately 43º. The indicator is operated by apotentiometer driven by the right hand inboard

Figure 15-1 Flap Control System


15-3

flap. Flap position limit switches are also drivenby the RH inboard flap.

The flap motor power circuit is protected by a20-ampere circuit breaker placarded FLAPMOTOR, located on the right hand circuit breakerpanel (Figure 15-4). A 5-ampere circuit breaker,placarded FLAP IND & CONTROL, for the flapcontrol circuit is also located on this panel.

C90A FLAP OPERATION

From the UP position to the APPROACH posi-t i on , t he flaps canno t be s topped a t anintermediate point. Between APPROACH andDOWN, the flaps may be stopped as desired bymoving the handle to DOWN position until theflaps have moved to the desired position, thenmoving the handle back to APPROACH. In likemanner, the flaps may be raised to any positionbetween DOWN and APPROACH by raising thehandle to UP until the desired setting is reached,then returning the handle to APPROACH. TheAPPROACH detent acts as a stop for any posi-tion greater than 35%. Moving the handle fromDOWN to APPROACH will not retract the flaps.When the flaps are at APPROACH and the han-dle is moved from APPROACH position to theUP position, the flaps retract completely and can-not be stopped in between.

C90B FLAP OPERATION

Flaps are selectable to 3 positions: up, approach(15º), and down (43º). If a go-around is initiatedwith flaps fully extended, retraction to eitherapproach or full-up positions can be accom-plished with a single switch position selection.

Figure 15-2 Flap Control Lever

Figure 15-3 Flap Position Indicator

Figure 15-4 Flap System Circuit Breaker

15-4


LANDING GEAR WARNING SYSTEM

The landing gear warning system is provided towarn the pilot that the landing gear is not downand locked during specific flight regimes. Thewarning horn will sound continuously when theflaps are lowered beyond the APPROACH (30%)position, regardless of the power lever setting,until the landing gear is extended or the flaps areretracted. Although the landing gear warning sys-tem is affected by the flap position, this subject isdiscussed more completely in the LANDINGGEAR section of this training manual.

FLAP AIRSPEED LIMITS

Airspeed indicator (Figure 15-5) markings showthe maximum speeds and operating range of theflaps V

FE

). The white triangle indicates maxi-mum flaps-to or at-approach (30%) speed. Theupper limit of the narrow white arc is the maxi-mum speed permissible with flaps extendedbeyond APPROACH (more than 30%). Approachspeed (flaps 30%) is 184 KIAS. BeyondAPPROACH position, the maximum speed is148 KIAS.

Lowering the flaps will produce these results:

●

Attitude—Nose up

●

Airspeed—Reduced

●

Stall speed—Lowered

NOTE

All illustration needles may not reflectnormal indications.

RUDDER BOOST SYSTEM

A rudder boost system (Figure 15-6) is providedto aid the pilot in maintaining directional controlin the event of an engine failure or a large varia-tion of power between the engines. Incorporatedinto the rudder cable system are two pneumaticrudder-boosting servos that actuate the cables toprovide rudder pressure to help compensate forasymmetrical thrust.

Figure 15-5 Airspeed Indicator


15-5

P SWITCH

18 PSIPNEUMATICPRESSURE

REGULATOR CHECKVALVE

CHECKVALVE

N.C.N.C.

LEFTRUDDERSERVO

LEFT P3AIR

RIGHTP3 AIR

AFT PRESSURE BULKHEAD

ELECTRICAL LINES

HIGH PRESSURE P3 AIR

REGULATED P3 AIR

13 PSIPRESSURE

REGULATOR

RUDDERBOOST

OFF

RUDDER

5BOOST

CENTER BUS

FILTER

RIGHTRUDDERSERVO

Figure 15-6 Rudder Boost System Diagram

15-6


The rudder boost system consists of pneumaticactuators in the empennage which provide therequired rudder deflection upon loss of anengine. A differential pressure switch, mountedon the pneumatic manifold, senses engine P

3

pressures. Upon sensing a loss of P

3

on oneengine, this pressure switch will energize a sole-noid to direct pneumatic manifold air to theappropriate actuator.

During operation, a differential pressure switchsenses bleed air pressure differences between theengines. If the bleed air pressure differential

exceeds about 50 psi differential pressure, a sig-nal from the differential pressure switch to one ofthe lines to the rudder boost servos causes thesolenoid valve to open, and one of the servos isactuated. The pressurized servo will then pull onone of the rudder cables. Tension springs in theconnection between the servos and the ruddercables take up the slack in the rudder cable whenone or the other of the servos is actuated.

A drop in bleed air pressure from the left enginewill actuate the appropriate servo and the rightrudder pedal will move forward. A drop in bleed

Figure 15-7 Rudder Boost Switch


15-7

air pressure from the right engine will cause theleft rudder pedal to move forward. Pedal riggingcauses the opposite pedal to move in the oppositedirection. This system is intended to help com-pensate for asymmetrical thrust only. Appropriatetrimming is to be done with the trim controls.

The system is controlled by a toggle switch (Fig-ure 15-7), placarded RUDDER BOOST - OFF,located on the pedestal below the aileron trimcontrol knob. The switch is to be in RUDDERBOOST position before flight.

The circuit is protected by the 5-ampereRUDDER BOOST circuit breaker on the rightside panel.

A preflight check of the system can be performedduring the run-up by retarding the power on oneengine to idle, and advancing power on the oppo-site engine until the power difference betweenthe engines is great enough to close the switchthat activates the rudder boost system. Movementof the appropriate rudder pedal (left engineidling, right rudder pedal moves forward) will benoted when the switch closes, indicating the sys-tem is functioning properly for low engine poweron that side. Repeat the check with oppositepower settings to check for movement of theopposite rudder pedal.


16-i

CHAPTER 16

CONTENTS

Page

AVIONICS

INTRODUCTION .................................................................................................................

16-1

DESCRIPTION......................................................................................................................

16-1

AVIONICS POWER DISTRIBUTION.................................................................................

16-3

KING SILVER CROWN II EQUIPMENT ...........................................................................

16-7

Audio Control System ...................................................................................................

16-7

Communications Transceiver System ...........................................................................

16-8

VOR/LOC/GS Receiver System ...................................................................................

16-8

DME System .................................................................................................................

16-9

RNAV System...............................................................................................................

16-9

ADF System ................................................................................................................

16-10

COLLINS PRO LINE II EQUIPMENT ..............................................................................

16-11

NAV System................................................................................................................

16-11

DME System ...............................................................................................................

16-12

COMM System............................................................................................................

16-13

ADF System ................................................................................................................

16-15

Transponder System ....................................................................................................

16-15

DB-415 AUDIO SYSTEM ..................................................................................................

16-15

Normal Operation........................................................................................................

16-17

EMERGENCY OPERATION ....................................................................................

16-17

SLAVED COMPASS SYSTEMS .......................................................................................

16-18

KCS-55A, MCS-65, MCS-103, and C-14A-43 Systems ............................................

16-18

16-ii


Collins PN-101 System...............................................................................................

16-22

PITOT-STATIC SYSTEM..................................................................................................

16-23

Introduction.................................................................................................................

16-23

Description..................................................................................................................

16-23

Pitot and Static System ...............................................................................................

16-23


16-iii

Figure Title Page

ILLUSTRATIONS

16-1

Nav/Comm Control Panel......................................................................................

16-2

16-2

Avionics Master Switch OFF ................................................................................

16-3

16-3

Avionics Master Switch ON..................................................................................

16-4

16-4

Alternate Avionics Bus Power...............................................................................

16-4

16-5

Avionics Buses ......................................................................................................

16-5

16-6

Inverter Power Supply ...........................................................................................

16-6

16-7

King Audio Control System ..................................................................................

16-7

16-8

King Communications Transceiver .......................................................................

16-8

16-9

King VOR/LOC/GS Receiver ...............................................................................

16-9

16-10

King DME System.................................................................................................

16-9

16-11

King ADF System................................................................................................

16-10

16-12

Pro Line II NAV Control .....................................................................................

16-12

16-13

Pro Line II DME Control.....................................................................................

16-13

16-14

Pro Line II Single DME Installation....................................................................

16-13

16-16

Pro Line II COM Control ....................................................................................

16-14

16-15

Pro Line II Dual DME Installation ......................................................................

16-14

16-17

Pro Line II ADF Control......................................................................................

16-15

16-18

Pro Line II Transponder Control .........................................................................

16-16

16-19

DB-415 Audio System Diagram..........................................................................

16-16

16-20

DB-415 Audio Switch Panel ...............................................................................

16-17

16-21

DB-415 Normal Operation Schematic.................................................................

16-18

16-22

DB-415 Emergency Operation Schematic...........................................................

16-19

16-23

Slaved Compass System Block Diagram ............................................................

16-20

16-iv


16-24

Compass System Connections ............................................................................

16-21

16-25

Collins PN-101 Compass System .......................................................................

16-22

16-26

Pitot and Static System Schematic ......................................................................

16-23

16-27

Pitot-Static Normal-Alternate Air Source Valve ................................................

16-24

16-28

Schematic Diagram of Pitot and Static System...................................................

16-25

16-29

Airspeed Calibration-Emergency System Graph ................................................

16-26

16-30

Altimeter Correction-Emergency System Graph ................................................

16-26

16-31

Ice Protection Switches .......................................................................................

16-27


16-1

CHAPTER 16AVIONICS

INTRODUCTION

Avionics systems, as a vital part of the airplane, are becoming more sophisticated andcomplex. These systems lighten the pilot load, particularly during IFR operations. It is there-fore important for the flight crew to understand how the various nav/comm systems function,and how to use them effectively. This section describes the standard avionics installation andhow it operates.

DESCRIPTION

King Air avionics controls (Figure 16-1), alongwith the weather radar, are mounted on an isola-tion panel in the center of the instrument panel,easily accessible to the pilot or copilot. Individ-ual audio switches, across the top of the panel,control audio to the speakers or headphones.

The King Silver Crown II line of panel-mountedavionics equipment is installed on many KingAirs. Although not all equipment types in the Sil-ver Crown II line will be discussed here, all ofthe main units typically installed in a King Airwill be addressed.

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Figure 16-1 Nav/Comm Control Panel


16-3

The Collins Pro Line II remote-mounted avionicspackage is optionally available on the C90A andstandard on the C90B, and is also described inthis section.

AVIONICS POWER DISTRIBUTION

All avionics equipment may be turned on and offby the avionics master switch (Figures 16-2 and16-3). In the event that this switch fails, powermay be restored by pulling the avionics master cir-cuit breaker, located in the upper right-hand cornerof the main circuit breaker panel (Figure 16-4).

The King Air C90A has three avionics buses(Figure 16-5) to feed DC power to the varioustypes of avionics equipment. To determine spe-cifically what equipment is being fed from aspecific bus or power source, refer to the wiringdiagram entitled “DC Power Distribution”which is supplied with each airplane. There are,

however, some general rules of thumb whichusually apply.

For example:

●

Items numbered one (e.g., comm 1, nav1, etc.) are fed by the number one avion-ics bus, which in turn is fed from theelectrical system triple-fed bus. It isimportant to note that in the event of adual generator failure, the items fed bythe number one avionics bus would con-tinue to operate for a limited period oftime, being fed directly by the battery.

●

Items numbered two (e.g., comm 2, nav2, etc.) are fed by the number 2 avionicsbus. The number two avionics bus is fedby the left generator bus.

●

Additional avionics items which are notfed by the previous buses are fed by thenumber 3 avionics bus. The number 3av ion i c s bus i s f ed by t he r i gh tgenerator bus.

Figure 16-2 Avionics Master Switch OFF

16-4


Figure 16-3 Avionics Master Switch ON

Figure 16-4 Alternate Avionics Bus Power


16-5

During a normal engine starting sequence, whena generator is brought on line, both generator busties close. Therefore, assuming the avionics mas-ter switch is turned ON, all avionics systems willreceive power from their respective buses undernormal circumstances. Also, when runningequipment checks on the ground with the exter-nal power switch ON and an APU connected, allthree avionics buses will be powered. In theseinstances, the bus ties are automatically closed.

However, assume the need to make a quickground check of comm 2, prior to starting

engines, and without an APU connected. In thissituation, manually close the bus ties with theappropriate switch located on the pilot’s outboardsubpanel.

As a general rule of thumb, an APU should beconsidered essential for running avionics equip-ment on the ground. For electronic flightinstrument system (EFIS) equipped airplanes, theavionics equipment and one of the invertersrequire approximately fifty amperes of currentfrom the battery. This amount of current drain

Figure 16-5 Avionics Buses

16-6


would deplete the battery in a short periodof time.

Also, especially with EFIS equipment installed,it is desirable to have the avionics nose compart-ment doors removed to allow sufficient crossventilation and cooling of the equipment. Partic-ularly during practice sessions with the avionicsequipment which exceed fifteen minutes induration.

AC power is available from either of two 400 Hzinverters. Under normal circumstances, the num-ber one inverter is fed from the left generator busand the number two inverter is fed from the rightgenerator bus (Figure 16-6). However, in theevent that the operating inverter loses power fromits appropriate bus, the inverter automaticallyswitches over to the center bus as its powersource.

Figure 16-6 Inverter Power Supply


16-7

KING SILVER CROWN II EQUIPMENT

The King Silver Crown II line of panel-mountedavionics equipment is installed on many KingAirs. Although not all equipment types in the Sil-ver Crown II line will be discussed here, all ofthe main units typically installed in a King Airwill be addressed. For additional information oneach system, please consult the appropriatepilot’s guide.

AUDIO CONTROL SYSTEM

The KMA 24 Audio Control System (Figure16-7) consists of a rotary microphone selectorswitch, speaker and phone switches for eachreceiver installed in the aircraft, and an integral

marker beacon receiver with marker beaconlights.

The microphone selector switch connects themicrophone to each transmitter installed on theaircraft. On versions of the KMA 24 which donot have the capability of handling audio from aNumber 2 ADF, the proper comm receiver audioswitch may be automatically selected by simplypushing either the speaker or phone AUTOswitch to the ON position and placing the micro-phone selector switch to the desired transmitter.

Versions of the KMA 24 which have the capabil-ity of handling audio from a second ADF do nothave the AUTO buttons and therefore the appro-priate comm receiver must be manually selectedeach time the microphone selector switch ischanged to a different transmitter.

Figure 16-7 King Audio Control System

16-8


Each receiver may be connected to either thespeaker and/or the phones by pushing the appro-priate alternate action pushbutton switch to the inor ON position.

The marker beacon receiver has a HIGH (buttonpushed in) and a LOW (button out) sensitivityposition. The marker beacon lights may also betested by pushing and holding the test button in.A built in photo cell automatically adjusts thelighting intensity depending on ambient lightingconditions.

COMMUNICATIONS TRANSCEIVER SYSTEM

The KY 196 (Figure 16-8) is capable of transmit-ting and receiving a frequency range of 118.0through 135.975 MHz in either 25 or 50 kHzsteps.

The large frequency knob changes the frequencyto the left of the decimal point while the smallerknob changes the frequency to the right of thedecimal point. The smaller knob makes 50 kHzchanges when pushed in and 25 kHz changeswhen pulled out.

The ON/OFF/VOLUME control switch turns theunit on when rotated clockwise past the initialdetent. Further clockwise rotation increases thevolume level. Pulling out this control “opens up”

the receiver squelch circuit, enabling the pilot tohear weaker stations. This might be an appropri-ate action when attempting to receive a weaktransmitter from a distance, such as listening toan ATIS at a distant point.

The left frequency display indicates the fre-quency to which the transceiver is actively tuned.The right display indicates the “standby” fre-quency. In order to transfer or swap the twofrequencies, the pilot pushes the transfer buttonmomentarily. (The frequency selector knob onlychanges the “standby” frequency.)

Transmitter operation is annunciated by the illu-mination of the letter “T” located between theactive and standby frequencies.

VOR/LOC/GS RECEIVER SYSTEM

Operation of the KN 53 (Figure 16-9) is virtuallyidentical to that of the KY 196 comm transceiverwith the following exceptions:

●

Pulling out on the volume control knobactivates the Morse code identificationcircuit, thus allowing the “ident” to beheard through the audio system.

●

There is no transmit annunciator on thissystem.

Figure 16-8 King Communications Transceiver


16-9

DME SYSTEM

The remote-mounted KN 63 DME with theKDI 572 panel-mounted indicator operates in astraightforward manner. The indicator is capableof displaying DME distance, ground speed, andtime to station simultaneously (Figure 16-10).

The mode selector allows the unit to be chan-neled by either nav 1 or nav 2. Selecting the HLD(hold) position allows the DME to remain chan-neled to the previously selected frequency, and isannunciated by either H1 or H2 depending on

whether nav 1 or nav 2 was previously used. Themode selector also allows the DME to be turnedoff.

RNAV SYSTEM

For operational information on the KNS 81 sys-tem, refer to the appropriate flight manualsupplement.

Figure 16-9 King VOR/LOC/GS Receiver

Figure 16-10 King DME System

16-10


ADF SYSTEM

The KR 87 ADF (Figure 16-11) has two basicmodes of operation, ANT (antenna) and ADF. Inthe ANT mode, the bearing pointer in theRMI/ADF indicator will not point to the stationbut provides improved audio reception. The ADFmode is used for navigation purposes, allowingthe bearing pointer to point to the station. TheADF mode is selected by pushing the alternateaction pushbutton in, and the ANT mode isselected by allowing the same pushbutton toremain in the “out” position. The selected modeis annunciated on the left side of the ADFdisplay.

This unit incorporates a BFO (beat frequencyoscillator) circuit which allows non-directionalbeacons to be identified which are not modulat-ing the carrier with audio. These types of stationsare sometimes used outside of the United States.The BFO circuit, when activated by pushing the

BFO pushbutton to the “in” position, generates a1020 Hz tone which will be heard each time theNDB transmitter is turned on. This allows theMorse code to be identified in a normal fashion.

As with the KY 196 comm and KN 53 nav, twofrequencies may be displayed on the KR 87 ADF.The frequency on the left is always the frequencyin use, however, the right display window isshared by several different functions. Like thecomm and the nav, the right window may displaythe standby frequency. However, pushing theFLT/ET alternate action pushbutton changes thefunction of the right display window.

When FLT is annunciated to the right of the rightdisplay window, the display is being used to dis-play flight time. Initially, the flight timer beginsoperation when the unit is turned on. Then, dur-ing takeoff, the flight timer is reset to zero andbegins counting again when the weight of the air-craft is off the landing gear “squat switch.” The

Figure 16-11 King ADF System


16-11

flight timer continues to operate until the weightof the aircraft is once again on the landing gearsquat switch, at which time the display is “fro-zen,” and will remain so until power is removedor the aircraft takes off again.

Pushing the FLT/ET button again will switch theright window to display an elapsed time function.The elapsed timer may be reset to zero at anytime by momentarily pushing the SET/RST(set/reset) button. Elapsed time will continue toaccumulate until the SET/RST button is pushedagain or power is removed.

The elapsed timer also has a “countdown” modeof operation, which may be initiated by holdingthe SET/RST button in for approximately threeseconds, or until the ET annunciator begins toflash. Now, the countdown time (in minutes andseconds) may be set into the right display byrotating the two concentric knobs which are nor-mally used to change the frequency. Set theminutes with the large knob and the seconds withthe small knob.

In order to start the countdown cycle (as whenpassing the final approach fix) push the SET/RSTbutton. Time remaining will now be continuouslydisplayed until the timer reaches zero, at whichtime it will revert to a count up mode of opera-tion and will now automatically display theelapsed time above and beyond that which wasoriginally set in. Additionally, when the count-down mode switches to the count up mode, theright display window will flash for fifteen sec-onds in order to alert the pilot to the fact that hehas gone beyond the originally preset time.

With both the flight and elapsed timer, the dis-play will initially be read in minutes andseconds (up to 59 minutes and 59 seconds(59:59). After the first hour, these timers willdisplay hours and minutes.

Pushing the FREQ (frequency transfer) buttoninitially changes the right window back to thestandby frequency display. Subsequent pushes ofthe FREQ button transfers the standby and in-usefrequencies back and forth (flip-flops).

COLLINS PRO LINE II EQUIPMENT

A new series of Collins CTL control heads isused for the nav, comm, ADF, and Transponder.The ADF-60 and TDR-90 (ADF and transponderunits respectively), however, are retained fromthe earlier Collins Pro Line System.

The Pro Line II family presently consists of aVHF comm (VHF-22), a VOR/LOC/GS/MBreceiver (VIR-32), and a DME (DME-42). Theseunits employ many state-of-the-art features,including extensive self-diagnostic capabilitiesand multiple frequency storage. Some of the fea-tures of this equipment will be described here.For additional information, see the current Col-lins Pro Line II pilot’s guide.

The comm and nav units have many features incommon; therefore, we will use the features ofthe nav (VIR-32/CTL-32) as a building block forthe comm, which will be described next. Featurescommon to both will be described under the navexplanation; differences will be pointed outunder the comm explanation.

NAV SYSTEM

The VIR-32/CTL-32 nav system is comprised ofa VOR/localizer receiver, a glide slope receiver,and a marker beacon receiver, all contained inone “black box” located in the nose avionicscompartment.

The nav receiver (Figure 16-12) may be tuned tothe correct frequency in any one of three ways:

1. The ACTive frequency may be tuned directlyby first holding down the ACT push buttonfor approximately three seconds. The lower(PREset) frequency display will be dashedout. The two concentric frequency selectknobs will now directly channel the ACTivefrequency. Features such as DME hold, pre-set channels, etc., are still operable in thissituation.

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2. The PREset (standby) frequency may be ini-tially selected and displayed in the lowerfrequency window. It may be necessary tocancel the direct tuning mode (describedabove) by again holding the ACT pushbuttonfor approximately three seconds. Once thePREset frequency is displayed in the lowerwindow, it may be transferred up to theACTive window by holding the XFR/MEMswitch to the XFR position momentarily.

3. Up to four frequencies may be placed into thefour channel slots of the memory. This isdone by repeatedly pressing the XFR/MEMswitch to the MEM position until the desiredchannel number appears in the upper(ACTive) window (e.g., CH-1). Now the fre-quency may be selected using the twoconcentric frequency select knobs and will bedisplayed in the lower (PREset) window.Once selected, the frequency may be storedby simply pressing the STOre button twice.Subsequent frequencies/channels maybestored in a similar fashion.

Regardless of the frequency selection methodused, when a new frequency is selected, the com-pare annunciator (labeled ACT) will flash once if,in fact, the VIR-32 receiver has properly tuned tothe frequency displayed in the active window. Ifthe compare annunciator continues to flash, atuning fault is indicated. The test button shouldbe pressed momentarily in order to display thefault and diagnostic code (see pilot’s guide forfurther details).

DME hold may be selected by placing the modeselector switch in the HLD position. This topicwill be further discussed under the topic of DME,to be covered later in this section.

DME SYSTEM

By using frequency scanning techniques, theDME-42 is capable of working with up to threeDME stations simultaneously (Figure 16-13). Itcan display DME distance (NM), ground speed(GS), time to station (MIN), and station identifi-

Figure 16-12 Pro Line II NAV Control


16-13

cation to any one of these stations; however, theDME stays locked onto all three stations.

In a single DME-42 installation, the threefrequencies or channels are connected to thenav 1 and nav 2 control heads (CTL-32s)(Figure 16-14).

In a dual DME-42 installation, the number oneDME-42 is only connected to the number onenav control head. Likewise, the number twoDME-42 is only connected to the number twonav control head. In this configuration, eachDME-42 is purposefully limited to displayingonly two channels (Figure 16-15).

COMM SYSTEM

In most respects, the VHF-22 comm works justlike the features previously explained on theVIR-32 navigation receiver (Figure 16-16). Theprimary differences are as follows:

●

There are six frequency memory posi-tions instead of four.

●

In place of the HLD annunciator, there isa TX (unit transmitting) annunciator.

●

SQ OFF (squelch off) replaces HLD onthe mode selector.

●

Two short tones indicate a fault. Push thetest button to display fault code.

●

Continued turning of the small knobresults in 50 kHz steps. When reversedone click, however, a 25 kHz step results.

Figure 16-13 Pro Line II DME Control

Figure 16-14 Pro Line II Single DME Installation

16-14


Figure 16-16 Pro Line II COM Control

ACTPRE

NAV 1ACTPRE

NAV 2

1

CH1 - NAV 1 ACTIVE2 - NOT USED3 - NAV 1 PRESET

CH1 - NOT USED2 - NAV 2 ACTIVE3 - NAV 3 PRESET

2 3 DME 1

1 2 3 DME 2

DME 2

Figure 16-15 Pro Line II Dual DME Installation


16-15

ADF SYSTEM

The ADF control head also works like the navcontrol head in many respects, however, themodes on the mode selector switch are appropri-ate to an ADF (Figure 16-17).

Although the ADF utilizes one of the new ProLine II control heads (CTL-62), the actual ADFunit is of an older generation and it does not dis-play test codes.

TRANSPONDER SYSTEM

Like the ADF, the transponder unit is of an earliergeneration has a new Pro Line II control head(CTL-92), and does not display test codes (Fig-ure 16-18).

The transponder control head can store one pre-selected code, such as 1200, ready for use at thepush of the PRE button.

DB-415 AUDIO SYSTEM

The majority of the King Air C90As built to datehave the DB-415 audio system installed.Although other optional audio systems may beinstalled, the standard DB-415 system is the onlyone which will be described in this section.

The avionics system has dual DB-415 audio sys-tems which are totally independent of each other(Figure 16-19). The only exception to this rule isthat there is only one emergency/normal switchon the radio panel which serves both the pilot’sand copilot’s audio systems. Therefore, if theemergency/ normal switch is in either position,both systems will be in that mode of operation.

Figure 16-17 Pro Line II ADF Control

16-16


Figure 16-18 Pro Line II Transponder Control

Figure 16-19 DB-415 Audio System Diagram


16-17

NORMAL OPERATION

Under “normal” circumstances, the nor-mal/emergency switch should remain in theNORM position (Figure 16-20). The followingoperating rules apply when in the normal modeof operation. Rules will only be listed for thepilot’s audio system; however, they apply equallyto the copilot’s audio system.

●

The volume control on the microphoneselector switch regulates the volume levelfor both the pilot’s speaker and the pilot’sheadphones.

●

The speaker switch turns the speaker ONand OFF.

●

The headphones are operational at alltimes as long as they are plugged intotheir jack.

●

The speaker and headphone audio chan-nels are independent of each other andfailure of one does not necessarily implya failure of the other.

●

To select any audio source (e.g., comm 1,ADF, etc.) turn ON the appropriate audioselector switch.

●

The switch labeled VOICE-BOTH-RANGE (Figure 16-21) works with theADF and nav receivers. When in theVOICE position, the voice portion of theaudio will be heard and not the Morsecode station identification. When in theRANGE (ident) position, only the Morsecode station identification will be heard,not the voice portion. When in the BOTHposition, the voice and range portions ofthe audio will be heard. If the pilot’saudio system has failed entirely, the pilotmay still listen to audio through the copi-lot’s speaker. If this is undesirable for anyreason, or if both the pilot’s and copilot’saudio systems have failed, place theemergency/normal switch in the EMERposition.

EMERGENCY OPERATION

When in the emergency mode of operation, thefollowing operating rules apply (Figure 16-22):

●

All audio sources (comm 1, nav 2, ADF,etc.) are connected directly to theheadphones.

●

To eliminate any specific audio source,turn down the volume control on thataudio source (e.g., nav 1). This rule doesnot apply to comm 1 and comm 2.

●

The volume control located on the micro-phone selector switch has no function inthe emergency mode.

●

If the emergency/normal switch shouldfall for any reason, the audio systemsmay still be placed into the emergencymode of operation by pulling the two cir-cuit breakers labeled PLT AUDIO andCOPLT AUDIO located directly beneaththe avionics master circuit breaker, on themain circuit breaker panel.

Figure 16-20 DB-415 Audio Switch Panel

16-18


SLAVED COMPASS SYSTEMS

The most common compass system for the KingAir C90A is the King KCS-55A; however, theSperry C-14A-43 or the Collins MCS-65 orMCS-103 systems could be installed. As far asthe pilot is concerned, all of these systems oper-ate in a similar manner. They will be treated asone common system in the following discussion.

Occasionally, a Collins PN-101 system will beinstalled on the copilot’s side. This system oper-ates in a slightly different manner and will bediscussed separately in this section.

KCS-55A, MCS-65, MCS-103, AND C-14A-43 SYSTEMS

From an operational standpoint, all three of thesesystems may be treated identically. All of thesesystems require 400 Hz electrical power from aninverter. In the unlikely event that both invertersfail, these systems would be inoperative.

Figure 16-21 DB-415 Normal Operation Schematic


16-19

Each of them has the following components (Figure 16-23).

● Flux sensor (also called a flux gate or fluxvalve)—The function of this device is tosense the earth’s magnetic field relative tothe airplane and convert that informationinto an electrical signal which representsthe airplane’s magnetic heading.

● Slaving amplifier—The magnetic head-ing signal from the flux sensor is tooweak to be used directly; therefore, it isamplified (made larger or stronger) bythe slaving amplifier. The output signal isnow strong enough to directly drive a

torque motor in the directional gyro andthus maintain the gyro rotor in alignmentwith magnetic north.

● Directional gyro—Once the gyro rotor isaligned with magnetic north, it will have anatural tendency to stay there for a shortperiod of time, due to a force called gyro-scopic rigidity in space. This force willcontinue to keep the gyro mechanism inrelatively good alignment as long as thegyro rotor continues to turn at its designspeed. When the gyro drifts out of align-ment (precesses) the condition will besensed, and the magnetic heading referenceinformation from the slaving amplifier will

Figure 16-22 DB-415 Emergency Operation Schematic


+

FLUX SENSOR

SENSES MAGNETICHEADING ANDCONVERTES IT TOAN ELECTRICSIGNAL.

SLAVING AMPLIFIER

AMPLIFIES THEMAGNETIC HEADINGSIGNAL.

DIRECTION GYRO

PROVIDES GYROSTABILIZED MAGNETICHEADING.

H.S.I./R.M.I

DISPLAYS GYROHEADING.

MAGNETICHEADING

MAGNETICHEADING

FREE SLAVE

INCREASE

DECREASE

TO AUTOPILOTGYRO HEADING/MAGNETIC HEADING

SLAVING METER

DISPLAYS DIFFERENCE BETWEENSENSED AND INDICATED MAGNETIC HEADING

COLLINS MC-65, MC-103, KING KCS 55AAND SPERRY C-14A-43 (AC POWERED)

Figure 16-23 Slaved Compass System Block Diagram


again drive the gyro rotor back into align-ment with magnetic north, using the torquemotor previously described.

● Horizontal situation indicator(HSI)—The gyro heading information(which should be the same as magneticheading) is sent to a compass card onthe HSI to display the magnetic head-i n g t o t h e p i l o t . T h i s h e a d i n ginformation is then sent from the HSIto the compass card on the oppositeradio magnetic indicator (RMI). In thisway, gyro-stabilized, magnetic head-ing information is displayed in front ofeach p i lo t f rom two independen tsources, the pilot’s and the copilot’scompass systems (Figure 16-24).

● Slaving meter—The slaving meter com-pares the sensed magnetic heading at theflux sensor (system input) to the dis-placed magnetic heading at the HSI(system output). The difference, if any, isdisplayed on the slaving meter by dis-placement of the slaving needle from thecenter position (which indicates synchro-nization or zero error). It is normal forthis needle to deviate occasionally due toprecession, however, it should alwayscome back to center. If it is displaced toone side for more than approximately oneminute the gyro may be precessingexcessively and/or the slaving systemmay not be doing its Job. In any case, the

accuracy of the compass system shouldbe checked by cross referencing theheading information from the oppositesystem and/or the magnetic compass.

● SLAVE/FREE switch—This lever-lock-ing switch is used to select either theslaved or the free mode of operation forthe compass system.

This switch should normally remain inthe slaved mode of operation. In thismode, when power is initially applied tothe system, it will automatically “slave”itself to the correct magnetic heading andremain there throughout the flight, cor-recting for precession as necessary.

The free mode of operation is generallyreserved for occasions when the slaved(automatic) mode of operation has failedand the pilot wishes to revert to a direc-tional gyro mode of operation. This modemay also be used for flight in polarregions where extreme levels of magneticvariation exist. In this mode of operation,the flux sensor and the slaving amplifierare disconnected from the rest of the sys-tem. The result is that the pilot now has adirectional gyro (which will precess andmust be corrected manually using theincrease/decrease switch) which uses theHSI to display the heading informationfrom the directional gyro.

Figure 16-24 Compass System Connections


● Increase/decrease switch—This is a tog-gle switch which is spring loaded to thecenter (OFF) position. The switch may beused when in the free mode of operationto manually change the directional gyroto the left or right, thus increasing ordecreasing the displayed heading infor-mation. When in the slaved mode ofoperation, momentarily holding thisswitch in either position causes the sys-tem to “reset” itself to the fast-slavemode of operation, thereby correctingany displayed error at a rapid rate. Thiscould be helpful if for any reason thegy ro had t umb led o r p r ece s sedexcessively.

COLLINS PN-101 SYSTEMThis compass system (Figure 16-25) is fre-quently installed on the copilot’s side. It has theadvantage of being directly powered by the28-volt DC electrical system. If both invertersfail, the system would continue to operate. How-ever, the PN-101 system does not have a manualback-up mode of operation (FREE) if the slavingsystem (flux sensor and/or slaving amplifier)fails. The PN-101 system does have a fast-slaveswitch which may be momentarily held in the UPposition to initiate the fast-slaving sequence (seefast-slaving explanation under increase/decreaseswitch above). Except for the differences men-tioned here, the basic operation of this system isvirtually identical to that of the Collins MCS-65,MCS-103 and the Sperry C-14A-43 systemsdescribed previously.

FLUXSENSOR

SLAVINGAMPLIFIER

DIRECTIONALGYRO

H.S.I./R.M.I.

FASTSLAVE

(D.C. POWERED)

Figure 16-25 Collins PN-101 Compass System


PITOT-STATIC SYSTEM

INTRODUCTIONBecause the pitot-static system is vital to the safeoperation of the airplane, the pilot must be famil-iar with the system, the instruments affected, andthe alternate system for emergency use. Thistraining unit identifies and describes how the sys-tems works, the instruments affected by it, andthe use of the alternate static air source.

DESCRIPTIONThe Pitot-Static section of the training manualpresents a description and discussion of the pitot-static system. The dual pitot-static system is vitalto airspeed indications in the airplane. The prin-ciple of operation, sources of static and pitotpressure, instruments that depend on the system,

and the alternate static air source are covered inthis section.

PITOT AND STATIC SYSTEMThe pitot and static system (Figure 16-26) pro-vides a source of impact air and static air foroperation of the flight instruments.

A heated pitot mast is located on each side of thelower portion of the nose. Tubing from the leftpitot mast is connected to the pilot’s airspeedindicator and tubing from the right pitot mast isconnected to the copilot’s airspeed indicator. Thepilot’s system is completely independent of thecopilot’s system.

The normal static system provides two separatesources of static air-one for the pilot’s flightinstruments, and one for the copilot’s. Each of

Figure 16-26 Pitot and Static System Schematic


the normal static air lines opens to the atmo-sphere through two static air ports-one on eachside of the aft fuselage-four ports total.

An alternate static airline is also provided for thepilot’s flight instruments. In the event of a failureof the pilot’s normal static air source (if forexample, ice accumulations should obstruct thestatic air ports), the alternate source can beselected by lifting the red spring clip retainer offthe PILOT’S EMERGENCY STATIC AIRSOURCE valve handle, located on the right sidepanel (Figure 16-27), and moving the handle aftto the ALTERNATE position. This will connectthe alternate static air line to the pilots flightinstruments. The alternate line obtains static airjust aft of the rear pressure bulkhead, from insidethe unpressurized area of the fuselage.

The pilot’s altimeter, vertical speed indicator, andairspeed indicator are connected to the pilot’sstatic air source. When the system is switched tothe pilot’s alternate air source, the pilot’s altime-ter and vertical airspeed indicator are affected, aswell as the pilot’s airspeed indicator. With alter-

nate air, the pilot’s airspeed indicator andaltimeter will read higher that actual, and the ver-tical speed indicator will show a momentaryclimb. The copilot’s airspeed indicator, altimeter,and vertical speed indicator are all on the copi-lot’s static air source and cannot be switched tothe alternate source (Figure 16-28).

Refer to the Airspeed Calibration-EmergencySystem, and the Altimeter Correction-EmergencySystem graphs in the Performance section of thePilot’s Operation Handbook for operation whenthe alternate static air source is in use.

A sample Airspeed Calibration-Emergency Sys-tem graph from the Performance section of thePOH is shown in (Figure 16-29). When thepilot’s system is switched to ALTERNATE, usethis graph to determine the required IndicatedAirspeed to maintain a desired Calibrated Air-speed. For example, with flaps down, to maintaina Normal system IAS of 100 knots, an emer-gency system IAS of about 98 knots is required.

Figure 16-27 Pitot-Static Normal-Alternate Air Source Valve


A sample ALTIMETER CORRECTION-EMER-GENCY SYSTEM graph is shown in (Figure16-30). In this sample, to maintain an altitude of3000 feet MSL at 120 KIAS it would be neces-sary to maintain an indicated altitude of 3045 feetMSL. The graph indicates the indicated altitudeis 45 feet below actual altitude.

When the alternate static air source is notneeded, ensure that the PILOT’S EMER-GENCY STATIC AIR SOURCE valve handle isheld in the forward (NORMAL) position by thespring clip retainer.

Three petcocks are provided to facilitate drain-ing moisture from the static air lines. They arelocated behind an access cover below the circuitbreakers on the right side panel. These are pri-marily intended for maintenance personnel andthe drain valves should be opened to release anytrapped moisture at each 100-hour inspection,and after exposure to visible moisture on theground. They must be closed after draining. Ifthe drains were to be opened in-flight, pressur-ized air form the cabin would rush into the staticports of the instruments, resulting in possibleinstrument damage.

Figure 16-28 Schematic Diagram of Pitot and Static System


Figure 16-29 Airspeed Calibration-Emergency System Graph

Figure 16-30 Altimeter Correction-Emergency System Graph


The pitot masts can be heated electrically forflight in icing conditions. The pitot heat switchedare located in the lower right hand corner of theICE PROTECTION control panel in the righthand pilot’s subpanel (Figure 16-31). It is cus-tomary, as a precautionary measure, to have the

pitot heat ON during flight at altitude when thereis visible moisture and temperatures are near plus5∞C. There is no restriction on use of pitot heatexcept not to use it excessively on the groundwhere there is no air flow around the masts.

Figure 16-31 Ice Protection Switches on Pilot’s Subpanel


17-i

CHAPTER 17

CONTENTS

Page

MISCELLANEOUS SYSTEMS

INTRODUCTION ................................................................................................................

17-1

DESCRIPTION.....................................................................................................................

17-1

OXYGEN SYSTEM .............................................................................................................

17-3

Manual Plug-in System .................................................................................................

17-3

Diluter-Demand Crew Oxygen Masks ..........................................................................

17-5

Plug-in Masks................................................................................................................

17-5

Oxygen Supply Cylinder ...............................................................................................

17-5

Oxygen System Controls...............................................................................................

17-6

Oxygen Duration ...........................................................................................................

17-6

Oxygen Duration Computation .....................................................................................

17-6

Time of Useful Consciousness ......................................................................................

17-7

PHYSIOLOGICAL TRAINING ..........................................................................................

17-7

What Is It? .....................................................................................................................

17-7

Who Needs It? ...............................................................................................................

17-7

Where Can You Get It? .................................................................................................

17-8

How Long Is the Course? ..............................................................................................

17-8

What Is Contained in the Course? .................................................................................

17-8

What Are the Prerequisites for Training?......................................................................

17-9

How Do You Apply for Training? ................................................................................

17-9

How Can You Get Further Information? ......................................................................

17-9

SERVICING THE OXYGEN SYSTEM..............................................................................

17-9

17-ii


Filling the Oxygen System ...........................................................................................

17-9

King Air C90A and C90B Capacity ...........................................................................

17-10

Oxygen Cylinder Retesting.........................................................................................

17-10


17-iii

Figure Title Page

ILLUSTRATIONS

17-1

Oxygen System Schematic—Typical C90A .........................................................

17-2

17-2

Plug-in Type Oxygen Mask...................................................................................

17-3

17-3

Oxygen Mask Donned ...........................................................................................

17-3

17-4

Crew Oxygen Mask ...............................................................................................

17-3

17-5

Oxygen Cylinder Installation.................................................................................

17-4

17-6

Oxygen System Control Handle ............................................................................

17-4

17-7

Oxygen Pressure Gage...........................................................................................

17-4

17-8

Oxygen Fill Valve and Gage .................................................................................

17-5

17-9

Percent of Usable Oxygen Capacity ......................................................................

17-6

17-10

FAA Altitude Chamber..........................................................................................

17-8


17-v

Table Title Page

TABLES

17-1

Oxygen Duration (Minutes)...................................................................................

17-7

17-2

Time of Useful Consciousness ..............................................................................

17-7


17-1

CHAPTER 17MISCELLANEOUS SYSTEMS

INTRODUCTION

Pilot and passenger comfort and safety are of prime importance in operating this airplane. Thetask is to teach flight crewmembers to use the oxygen system safely and effectively, whenrequired, within the requirements of applicable FARs.

DESCRIPTION

This chapter presents a description and discus-sion of the oxygen system. It includes generaldescription, principle of operation, controls, andemergency procedures. Use of the oxygen dura-tion chart involves working simulated problemsunder various flight conditions. FAR require-

ments for crew and passenger needs are part ofthe discussion, as well as the types and availabil-ity of oxygen masks. Local servicing proceduresreferenced in the


arealso included.

��

��

��

1612

8

40

NO 1 FUELTRANS

NO 1 FUELLOWBATTHOT

ANTI-ICEON

OIL

BLOWEROFF

ENG 1CHIP

NO 1 FUELFILTER

NO 1 BATTSYS

XMSNOIL

90° BOXOIL

GEN 1HOT

��

TEST

RESET

17-2


Figure 17-1 Oxygen System Schematic—Typical C90A


17-3

OXYGEN SYSTEM

Current FARs require that anytime an aircraftflies above 25,000 feet, oxygen must be immedi-ately available to the crew and passengers. TheKing Air C90A and C90B systems comply withthis requirement.

The oxygen system (Figure 17-1) provides anadequate flow for an altitude of 30,000 feet. Themasks and Oxygen Duration chart (Normal Pro-cedures section of the

POH

) are based on 3.7LPM-NTPD. The only exception is the diluter-demand crew mask when used in the 100%mode. For oxygen duration computation, eachdiluter-demand mask being used in the 100%mode is counted as two masks at 3.7 LPM-NTPD each.

MANUAL PLUG-IN SYSTEM

The manual plug-in system is of the constant-flow type (Figures 17-2 and 17-3). Each maskplug is equipped with its own regulating orifice.The pilot and copilot oxygen masks are quick-

donning oxygen masks and are connected to theoxygen supply lines at all times (Figure 17-4).When the diluter demand masks are not in use,one hangs from a bracket (on the stub partition)behind the pilot’s head and one hangs from abracket behind the copilot’s head.

Figure 17-2 Plug-in Type Oxygen Mask

Figure 17-3 Oxygen Mask Donned

Figure 17-4 Crew Oxygen Mask

17-4


Passenger masks are kept in seatback pocketsexcept in the couch installation, in which casethey are stored under the couch. The cabin outletsare located at both the forward and aft ends of thecabin. All masks are easily plugged in by pushingthe orifice in firmly and turning clockwiseapproximately one-quarter turn. Unplugging iseasily accomplished by reversing the motion.

The oxygen supply cylinder is in the aft unpres-surized area of the fuselage (Figure 17-5). Theoxygen system pressure regulator and controlvalve are attached to the cylinder, and are acti-vated by a remote push/pull knob located to therear of the cockpit overhead light control panel(Figure 17-6). When this control is pushed in, no

oxygen supply is available anywhere in the air-plane. When this control is pulled out, the oxygensystem is charged with oxygen ready for use pro-vided the oxygen supply cylinder is not empty.The oxygen supply pressure gage is located in thecopilot’s right subpanel (Figure 17-7).

Figure 17-5 Oxygen Cylinder Installation

Figure 17-6 Oxygen System Control Handle

Figure 17-7 Oxygen Pressure Gage


17-5

DILUTER-DEMAND CREW OXYGEN MASKS

The crew are provided with diluter-demand,quick-donning oxygen masks (see Figure 17-4).These masks hang on the aft cockpit partitionbehind and outboard on the pilot and copilotseats. They are held in the armed position byspring tension clips, and can be donned immedi-ately with one hand. The diluter-demand crewmasks deliver oxygen to the user only upon inha-lation. Consequently, there is no loss of oxygenwhen the masks are plugged in and the PULLON handle is pulled out, even though oxygen isimmediately available upon demand.

A small lever on each diluter-demand oxygenmask permits the selection of two modes of oper-ation: NORMAL and 100%. In the NORMALposition, air from the cockpit is mixed with theoxygen supplied through the mask. This reducesthe rate of depletion of the oxygen supply, and itis more comfortable to use than 100% aviator’sbreathing oxygen. However, in the event ofsmoke or fumes in the cockpit, the 100% positionshould be used to prevent the breathing of con-taminated air. For this reason, the selector leversshould be left in the 100% position when themasks are not in use so the masks are alwaysready for maximum emergency use.

PLUG-IN MASKS

The plug-in oxygen masks in the cabin (see 17-2)are designed to be adjustable to fit the averageperson with minimum leakage of oxygen. To donthe mask, fit the nose and mouth piece over theface and adjust the elastic headband over thehead to hold the mask firmly in place. Insert thefitting in one of the oxygen outlets in the over-head cavity, push in firmly, and turn clockwiseapproximately one-quarter turn to lock it inplace. If oxygen is available (the system is turnedon and the oxygen cylinder charged), the red flowindicator will move and the green portion willcome into view. The mixing bag will inflate withbreathing. Breath normally. System efficiency is

determined by the fit of the oxygen mask. Makecertain the masks fit properly and are in goodcondition. The hose plug must be disconnected tostop the flow of oxygen.

There are certain important considerations anytime oxygen is in use. Do not use combustibleproducts near oxygen. Common items such aschapstick, lipstick, women’s makeup, or mustachewax could spontaneously ignite in the presence ofoxygen. These items should be removed beforeusing oxygen. No smoking should be allowed inthe airplane when oxygen is in use.

OXYGEN SUPPLY CYLINDER

Oxygen for flight at high altitudes is supplied bya cylinder mounted behind the aft pressure bulk-head. The cylinder is filled by a valve accessiblethrough an access door on the right side of theaft fuselage. The high-pressure system has twopressure gages, one on the copilot’s RH sub-panel in the cockpit for in-flight use (see Figure17-7), and one adjacent to the filler valve forchecking the pressure of the system during fill-ing (Figure 17-8). The cylinder is available inthree different capacities: 22 cubic feet, 49cubic feet, or 66 cubic feet.

Figure 17-8 Oxygen Fill Valve and Gage

17-6


OXYGEN SYSTEM CONTROLS

A shutoff valve regulator in the cylinder isactuated by its a push-pull shutoff controllocated overhead between the pilot and copilotseats (see Figure 17-6). Pushing in the handledeactivates the oxygen supply, while pulling outthe handle actuates the oxygen supply. Theregulator is a constant-flow type which supplieslow-pressure oxygen through aluminumplumbing to the outlets.

OXYGEN DURATION

A preflight requirement is to check the oxygenavailable, considering the number of crew andpassengers, to assure that it is sufficient fordescent to 12,500 feet, or until loss of pressure inthe airplane can be corrected and cabin altitudepressure restored. Full oxygen system pressure is1800 ±50 psi at 70° F for the 22 cubic feet cylin-der, and 1850 ±50 psi for the larger cylinders.First, read the oxygen pressure gage and note thepressure. Determine from the OXYGEN AVAIL-ABLE WITH PARTIALLY FULL BOTTLEgraph the percent of usable capacity. To obtainthe duration in minutes of the supply, obtain theduration for a full bottle from the Oxygen Dura-tion table, considering the number of personsaboard. Multiply the full bottle duration by thepercent of full bottle available to obtain the avail-able oxygen duration in minutes.

On the C9OA or C9OB airplane, oxygen dura-tion is for a Puritan-Zep oxygen system whichmust use the red, color-coded, plug-in mask,rated at 3.7 standard liters per minute – normaltemperature pressure (SLPM – NTPD) flow.Both aircraft are approved for altitudes up to30,000 feet.

OXYGEN DURATION COMPUTATION

In this sample computation, oxygen duration iscomputed for a Puritan-Zep oxygen systemwhich utilizes the red, color-coded, plug-in maskrated at 3.7 standard liters per minute (SLPM)flow and is approved for altitudes up to 30,000feet. This table is also used for the quick-donning, diluter-demand crew oxygen masks.When selected to the 100% mode, the number ofcrew masks in use should be doubled forcomputation. To compute oxygen duration forfour passengers and two crew members usingtheir masks in 100% mode, consider eight peopleusing oxygen.

To compute the duration in minutes of availableoxygen for eight people, assume the pressuregage shows 1,500 pounds. Enter the Percent ofUsable Oxygen Capacity chart (Figure 17-9) at1,500 pounds and read across to intersect the32° F diagonal, then down to read 85% of usablecapacity. To compute the duration available, enterthe Oxygen Duration chart (Table 17-1) at the8-people-using column and read down to 55 min-utes available for a 66 cubic-foot supply bottle.Now take 85% of 55 and find the current oxygenduration available of approximately 46 minutes.

Figure 17-9 Percent of Usable Oxygen Capacity


17-7

TIME OF USEFUL CONSCIOUSNESS

In the event of decompression at altitude, the pri-mary need is for oxygen to prevent hypoxia.Hypoxia is a lack of the oxygen needed to keepthe brain and other body tissues functioningproperly. The early symptoms of hypoxia, suchas an increased sense of well-being, quickly giveway to slow reactions, impaired thinking ability,unusual fatigue and a dull headache. Therefore,the crew must act quickly to don oxygen masksand supply oxygen to the passengers before theonset of hypoxia.

The ALTITUDE WARN annunciator illuminateswhen cabin altitude exceeds 10,000 feet (12,500for LJ-1353 and later), should the red ALTI-TUDE WARN annunciator illuminate due toinadequate cabin pressure, or loss of pressuriza-tion at high altitudes, crew and passengers shoulddon oxygen masks immediately and descend to asafe altitude.

The Time of Useful Consciousness table (Table17-2) shows the average time of useful con-sciousness available at various altitudes. This isthe time from the onset of hypoxia until loss ofeffective performance. Individuals may differfrom that shown in the table. Using the Emer-gency Descent procedure in the EmergencyProcedures section of the

POH

, a very rapiddescent can minimize the exposure to hypoxia.

PHYSIOLOGICAL TRAINING

WHAT IS IT?

Physiological training is a program directedtoward understanding and surviving in the flightenvironment. It covers the problems of both highand low altitudes and recommends procedures toprevent or minimize the human factor errorswhich occur in flight.

WHO NEEDS IT?

The course is primarily of benefit to pilots. It is alsorecommended for other air crew personnel, air traf-fic controllers, aviation medical examiners andother personnel from the national aviation system.

Table 17-1 OXYGEN DURATION (MINUTES)

CYL VOLCU FT

NUMBER OF PEOPLE USING*

1 2 3 4 5 6 7 8 9 10 11 12

22 151 75 50 37 30 25 21 18 16 15 13 12

49 334 167 111 83 66 55 47 41 37 33 30 27

66 445 222 148 111 89 74 63 55 49 44 40 37

Table 17-2 TIME OF USEFUL CONSCIOUSNESS

ALTITUDE TIME

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 to 18,000 feet .......................... 30 minutes or more

17-8


WHERE CAN YOU GET IT?

A resident physiological training course at theFAA’s Aeronautical Center in Oklahoma City isdevoted entirely to problems in civil aviation(Figure 17-10). Many military installations, andthe National Aeronautics and Space Administra-tion (NASA) in Houston, Texas, conduct aresident program for non-government personnel.

HOW LONG IS THE COURSE?

The course takes one full day.

WHAT IS CONTAINED IN THE COURSE?

Many topics are covered. They include the envi-ronment to which the flyer i s exposed ,physiological functions of the body at groundlevel, and alteration of some of these functionsby changes in the environment. The higher oneflies, the more critical becomes the need for sup-plemental oxygen. This need is discussed so thatthe trainee will understand why a pilot cannot flysafely at altitudes in excess of 12,500 feet for a

prolonged period without some aid, either sup-plemental oxygen or a pressurized aircraft. Bothoxygen equipment and pressurization are dis-cussed. When humans are confronted withcertain stressful situations, there is a tendency tobreathe too rapidly. This topic (hyperventilation)and methods of control are discussed. Ear painon descent and other problems with body gasesand procedures to prevent or minimize gas prob-lems are explained. Alcohol, tobacco, and drugsare also discussed as they apply to flying. Pilotvertigo is discussed and demonstrated so that thetrainee will understand why a non-current instru-ment pilot should never attempt to fly in cloudsand other weather situations where visibility isreduced. Resident courses include an altitudechamber flight where the trainees experienceindividual symptoms of oxygen deficiency aswell as decompression. This flight will demon-strate that:

1. Proper oxygen equipment and its use willprotect an individual from oxygen deficiency.

2. An individual can experience and recognizesymptoms that will be the same as thosefound in actual flight and therefore take the

Figure 17-10 FAA Altitude Chamber


17-9

necessary action to prevent loss of judgmentand consciousness.

3. Decompression is not dangerous providedproper supervision is present, and properac t ions a re p lanned and taken whennecessary.

WHAT ARE THE PREREQUISITES FOR TRAINING?

Personnel must have a valid FAA medical certifi-cate. A fee of twenty dollars is required. Theapplicant must be eighteen years of age or older.

HOW DO YOU APPLY FOR TRAINING?

All requests for the training course must be coor-dinated with:

FAA Airman Education Section (AAC - 142)Civil Aeromedical InstituteP.O. Box 25082Oklahoma City, Oklahoma 73125

HOW CAN YOU GET FURTHER INFORMATION?

Write to the Airman Education Section at theabove address, or phone (405) 686-4837.

SERVICING THE OXYGEN SYSTEM

The oxygen system is serviced by a filler valveaccessible by removing an access plate on theright side of the aft fuselage (see Figure 17-8).The system has two pressure gages, one on theright subpanel in the crew compartment for in-flight use, and one adjacent to the filler valve forchecking system pressure during filling. A shut-off valve and regulator on the cylinder control theflow of oxygen to the crew and passenger outlets.The shutoff valve is actuated by a push-pull con-

trol located aft of the overhead light control panelin the cockpit. The regulator is a constant-flowtype which supplies low-pressure oxygenthrough system plumbing to the outlets.

The following precautions should be observedwhen purging or servicing the oxygen system:

1. Avoid any operation that could create sparks.Keep all burning cigarettes or fire away fromthe vicinity of the airplane when the outletsare in use.

2. Inspect the filler connection for cleanlinessbefore attaching it to the filler valve.

3. Make sure that your hands, tools, and cloth-ing are clean, particularly of grease or oilstains. These contaminants are extremelydangerous in the vicinity of oxygen.

4. As a further precaution against fire, open andclose all oxygen valves slowly during filling.

FILLING THE OXYGEN SYSTEM

When filling the oxygen system, only use avia-tor’s breathing oxygen (MIL-0-27210).

WARNING

DO NOT USE MEDICAL OXYGEN.It contains moisture which can causethe oxygen valve to freeze.

Fill the oxygen system slowly byadjusting the recharging rate with thepressure regulating valve on the ser-vicing cart, because the oxygen, underhigh pressure, will cause excessiveheating of the filler valve. Fill the cyl-i nde r ( 22 -cub i c - foo t cy l i nde rin s t a l l a t i on ) t o a p r e s su re o f1,800 ±50 psi at a temperature of 70°F. This pressure may be increased anadditional 3.5 psi for each degree ofincrease in temperature; similarly, foreach degree of drop in temperature,reduce the pressure for the cylinder by

17-10


3.5 psi. The oxygen system, after fill-ing, will need to cool and stabilize fora short period before an accurate read-ing on the gage can be obtained. The49- or 66-cubic-foot cylinders may becharged to a pressure of 1,850 ±50 psiat a temperature of 70° F. When thesystem is properly charged, disconnectthe filler hose from the filler valve andreplace the protective cap on the fillervalve.

KING AIR C90A AND C90B CAPACITY

Oxygen for unpressurized, high-altitude flight issupplied by a cylinder in the compartment imme-diately aft of the pressure bulkhead (see Figure17-5). A 22-, 49-, or 66-cubic-foot cylinder maybe installed.

OXYGEN CYLINDER RETESTING

Oxygen cylinders used in the airplane are of twotypes. Lightweight cylinders, stamped “3HT” onthe plate on the side, must be hydrostaticallytested every three years and the test date stampedon the cylinder. This bottle has a service life of4,380 pressurizations or 15 years, whicheveroccurs first, and then must be discarded. Regularweight cylinders, stamped “3A,” or “3AA,” mustbe hydrostatically tested every five years andstamped with the retest date. Service life on thesecylinders is not limited.


APP-i

APPENDIXSYMBOLS, ABBREVIATIONS,

CONTENTS

Page

AND TERMINOLOGY

AIRSPEED........................................................................................................................

APP-1

METEOROLOGICAL......................................................................................................

APP-2

POWER.............................................................................................................................

APP-2

CONTROL AND INSTRUMENT ...................................................................................

APP-3

GRAPH AND TABULAR................................................................................................

APP-3

WEIGHT AND BALANCE .............................................................................................

APP-4

AVIONICS........................................................................................................................

APP-5


APP-1

APPENDIXSYMBOLS, ABBREVIATIONS,

AND TERMINOLOGY

AIRSPEED

CAS—Calibrated airspeed is the indicated air-speed of an airplane corrected for position andinstrument error. Calibrated airspeed is equal totrue airspeed in standard atmosphere at sea level.

GS—Groundspeed is the speed of an airplanerelative to the ground.

IAS—Indicated airspeed is the speed of an air-plane as shown on the airspeed indicator whencorrected for instrument error. IAS values pub-lished in this training manual assume zeroinstrument error.

KCAS—Calibrated airspeed expressed in knots.

KIAS—Indicated airspeed expressed in knots.

M—Mach number is the ratio of true airspeed tothe speed of sound.

TAS—True airspeed is the airspeed of an air-plane relative to undisturbed air, which is theCAS corrected for altitude, temperature, andcompressibility.

V

YSE

—Best single-engine rate-of-climb speed.

V

A

—Maneuvering speed is the maximum speedat which application of full available aerody-namic control will not overstress the airplane.

V

F

—Design flap speed is the highest speed per-missible at which wing flaps may be actuated.

V

FE

—Maximum flap extended speed is the high-est speed permissible with wing flaps in aprescribed extended position.

V

LE

—Maximum landing gear extended speed isthe maximum speed at which an airplane can besafely flown with the landing gear extended.

V

LO

—Maximum landing gear operating speed isthe maximum speed at which the landing gearcan be safely extended or retracted.

V

MCA

—Air minimum control speed is the mini-mum flight speed at which the airplane isdirectionally controllable, as determined inaccordance with Federal Aviation Regulations.The airplane certification conditions include: oneengine becoming inoperative and windmilling, a5° bank toward the operative engine, takeoffpower on operative engine, landing gear up, flapsin takeoff position, and most rearward CG. Forsome conditions of weight and altitude, stall canbe encountered at speeds above V

MCA

, as estab-lished by the certification procedure describedabove, in which event stall speed must beregarded as the limit of effective directionalcontrol.

V

MCG

—Ground minimum control speed.

V

MO

/M

MO

—Maximum operating limit speed isthe speed limit that may not be deliberatelyexceeded in normal flight operation. V isexpressed in knots and M in Mach number.

V

R

—Decision speed/rotation speed.

V

S

—Stalling speed or the minimum steady flightspeed at which the airplane is controllable.

V

SO

—Stalling speed or the minimum steadyflight speed at which the airplane is controllablein the landing configuration.

APP-2


V

SSE

—Intentional one-engine-inoperative speedis a speed above both V

MCA

and stall speed,selected to provide a margin of lateral and direc-tional control when one engine is suddenlyrendered inoperative. Intentional failing of oneengine below this speed is not recommended.

V

X

—Best angle-of-climb speed is the airspeedwhich delivers the greatest gain of altitude in theshortest possible horizontal distance.

V

Y

—Best rate-of-climb speed is the airspeedwhich delivers the greatest gain in altitude in theshortest possible time.

METEOROLOGICAL

Altimeter setting—Barometric pressure cor-rected to sea level.

Indicated pressure altitude—The number actu-ally read from an altimeter when the barometricsubscale has been set to 29.92 inches of mercury(1013.2 millibars).

IOAT—Indicated outside air temperature is thetemperature value read from an indicator.

ISA—International standard atmosphere inwhich:

●

Air is a dry, perfect gas.

●

Temperature at sea level is 59º Fahrenheit(15º Celsius).

●

Pressure at sea level is 29.92 inches ofmercury (1013.2 millibars).

●

Temperature gradient from sea level tothe altitude at which the temperature is–69.7º F (–56.5º C), is –0.003566º F(–0.00198º C) per foot, and is zero abovethe altitude.

OAT—Outside air temperature is the free airstatic temperature, obtained either from the tem-pera tu re ind ica to r ( IOAT) ad jus ted fo rcompressibility effects or from ground meteoro-logical sources.

Pressure altitude—Altitude measured from stan-dard sea level pressure (29.92 inches Hg) by apressure (barometric) altimeter. It is the indicatedpressure altitude corrected for position andinstrument error. In this training manual, altime-ter instrument errors are assumed to be zero.Position errors may be obtained from the altime-ter correction graphs.

Station pressure—Actual atmospheric pressure atfield elevation.

Temperature compressibility effects—An error inthe indication of temperature caused by airflowover the temperature probe. The error varies,depending on altitude and airspeed.

Wind—The wind velocities recorded as variableson the charts of this training manual are to beunderstood as the headwind or tailwind compo-nents of the reported winds.

POWER

Beta range—The range of propeller blade anglecontrol from the primary low-pitch-stop bladeangle setting to the full-reverse blade anglesetting.

Cruise climb—Cruise climb is the maximumpower approved for normal climb. These powersare torque or temperature (ITT) limited.

High idle—High idle is obtained by placing thecondition lever in the HIGH IDLE position. Thislimits the power operation to a minimum of 70%of N

1

rpm.

Low idle—Low idle is obtained by placing thecondition lever in the LOW IDLE position. Thislimits the power operation to a minimum of 51%of N

1

rpm.

Maximum continuous power—Maximum con-tinuous power is the highest power rating notlimited by time. Use of this rating is at the discre-tion of the pilot.


APP-3

Maximum cruise power—Maximum cruisepower is the highest power rating for cruise andis not time-limited.

Reverse—Reverse thrust is obtained by liftingthe power levers and moving them into the Betaplus power range.

SHP—Shaft horsepower.

Minimum takeoff power—Minimum takeoffpower is the minimum power which must beavailable for takeoff without exceeding theengine limitations.

Takeoff power—Takeoff power is the maximumpower rating. Use of this rating should be limitedto normal takeoff operations and emergencysituations.

CONTROL AND INSTRUMENT

Condition lever (fuel shutoff lever)—The fuelshutoff lever actuates a valve in the fuel controlunit which shuts off the fuel at the fuel controloutlet and regulates the idle range from low tohigh idle.

ITT (interstage turbine temperature)—Eightprobes, wired in parallel, sense the temperaturebetween the compressor and power turbines, andsend the reading to the ITT indicator in degreescentigrade x 100.

N

1

tachometer (gas generator rpm)—The N

1

tachometer registers the rpm of the gas generatorin percent, with 100% representing a gas genera-tor speed of approximately 37,500 rpm.

Power lever (gas generator N

1

rpm)—The powerlever serves to modulate engine power from fullreverse thrust to takeoff. The position for idlerepresents the lowest recommended level ofpower for flight operation.

Propeller control lever (N

P

rpm)—The propellercontrol is used to control the rpm setting of thepropeller governor. Movement of the lever results

in an increase or decrease in propeller rpm. Pro-peller feathering is the result of lever movementbeyond the detents at the low rpm end of thelever travel.

Propeller governor—The propeller governorsenses changes in rpm and hydraulically changespropeller blade angle to compensate for thechanges in rpm. Constant propeller rpm isthereby maintained at the selected rpm setting.

Torquemeter—The torquemeter system indi-cates the shaft output torque. Differentialpressure from the mechanism within the reduc-tion gearcase causes a bellows and servo systemto indicate torque on a meter. Instrument readoutis in foot-pounds.

GRAPH AND TABULAR

Accelerate-go—Accelerate-go is the distance toaccelerate to takeoff decision speed (V

R

), experi-ence an engine failure, continue accelerating toliftoff, then climb and accelerate in order toachieve takeoff safety speed (V

YSE

) at 50 feetabove the runway, for C90A aircraft and 35 feetfor C90B aircraft.

Accelerate-stop—Accelerate-stop is the distanceto accelerate to takeoff decision speed (V

R

) andthen bring the airplane to a stop.

AGL—Above ground level.

Best angle-of-climb—The best angle-of-climbdelivers the greatest gain of altitude in the short-est possible horizontal distance with gear andflaps up.

Best rate-of-climb—The best rate-of-climbdelivers the greatest gain of altitude in the short-est possible time with gear and flaps up.

Clearway—A clearway is an area beyond the air-port runway not less than 500 feet wide, centrallylocated about the extended centerline of the run-way, and under the control of the airportauthorities. The clearway is expressed in terms ofa clear plane, extending from the end of the run-way with an upward slope not exceeding 1.25%,

APP-4


above which no object nor any terrain protrudes.However, threshold lights may protrude abovethe plane if their height above the end of the run-way is 26 inches or less and if they are located toeach side of the runway.

Climb gradient—Climb gradient is the ratio ofthe change in height during a portion of a climbto the horizontal distance traversed in the sametime interval.

Demonstrated crosswind—Demonstrated cross-wind is the demonstrated crosswind componentfor which adequate control of the airplane duringtakeoff and landing was actually demonstratedduring certification; however, this is not consid-ered a limitation.

MEA—Minimum enroute altitude.

Net gradient of climb—Net gradient of climb isthe gradient of climb with the flaps in the takeoffposition and the landing gear retracted. “Net”indicates that the actual gradients of climb havebeen reduced by 8% to allow for turbulence andpilot technique. The net gradient of climb graphsare constructed so that the value(s) obtainedusing the airport pressure altitude and outside airtemperature will be the average gradient from 35feet above the runway up to 1,500 feet above therunway.

Route segment—Route segment is a part of aroute. Each end of that part is identified by a:

●

Geographic location, or

●

Point at which a definite radio fix can beestablished

Takeoff flight path—Takeoff flight path is theminimum gradient of climb required to clearobstacles in excess of 50 feet, measured horizon-tally from reference zero and vertically at thealtitude above the runway. Reference zero is thepoint where the airplane has reached 50 feetabove the runway, as determined from the accel-erate-go graphs.

WEIGHT AND BALANCE

Approved loading envelope—Those combina-tions of airplane weight and center of gravitywhich define the limits beyond which loading isnot approved.

Arm—Arm is the distance from the center ofgravity of an object to a line about whichmoments are to be computed.

Basic empty weight—Basic empty weight is theweight of an empty airplane, including fullengine oil and unusable fuel. This equals emptyweight plus the weight of unusable fuel, and theweight of all the engine oil required to fill thelines and tanks. Basic empty weight is the basicconfiguration from which loading data isdetermined.

Center of gravity—Center of gravity is the pointat which the weight of an object may be consid-ered concentrated for weight and balancepurposes.

CG limits—CG limits are the extreme center-of-gravity locations within which the airplane mustbe operated at a given weight.

Datum—Datum is a vertical plane perpendicularto the airplane’s longitudinal axis from whichfore and aft (usually aft) measurements are madefor weight and balance purposes.

Empty weight—Empty weight is the weight ofan empty airplane before any oil or fuel has beenadded. This includes all permanently installedequipment, fixed ballast, full hydraulic fluid, fullchemical toilet fluid, and all other operating flu-ids full, except that the engines, tanks, and linesdo not contain any engine oil or fuel.

Engine oil—Total system oil, including that por-tion of the engine oil which cannot be drainedfrom the engine.

Jack point—Jack points are points on the air-plane identified by the manufacturer as suitablefor supporting the airplane for weighing or otherpurposes.


APP-5

Landing weight—Landing weight is the weightof the airplane at landing touchdown.

Leveling points—Leveling points are thosepoints which are used during the weighing pro-cess to level the airplane.

Maximum weight—Maximum weight is thegreatest weight allowed by design, structural,performance, or other limitations.

Maximum zero fuel weight—Any weight abovethe value given must be loaded as fuel.

Moment—Moment is a measure of the rotationaltendency of a weight, about a specified line,mathematically equal to the product of theweight and the arm.

Payload—Payload is the weight of occupants,cargo, and baggage.

PPH—Pounds per hour.

Ramp weight—Ramp weight is the airplaneweight at engine start, assuming all loading iscompleted.

Station—Station is the longitudinal distancefrom some point to the zero datum or zero fuse-lage station.

Takeoff weight—Takeoff weight is the weight ofthe airplane at liftoff from the runway.

Tare—Tare is the apparent weight of any items(wheel chocks, jack stands, etc.) used on thescales but which are not a part of the airplaneweight.

Unusable fuel—Unusable fuel is the fuel remain-ing after consumption of usable fuel.

Usable fuel—Usable fuel is that portion of thetotal fuel which is available for consumption asdetermined in accordance with applicable regula-tory standards.

Useful load—Useful load is the differencebetween the airplane ramp weight and the basicempty weight.

Zero fuel weight—Zero fuel weight is the air-plane ramp weight minus the weight of fuel onboard.

AVIONICS

ADF mode—A mode of automatic directionfinder operation allowing the ADF needle topoint to the station.

NOTE

In this mode of operation, on manyreceivers the audio fidelity is severelylimited.

Air data computer—An electronic system prima-rily designed to gather information for anautopilot flight director system with outputsrelating to pitot and static data. Possible informa-tion from this system includes: pressure altitude,indicated airspeed, total air temperature, static airtemperature, and other information related toautopilot operation.

Altitude alert light—An amber light associatedwith an altitude alerter system. This light will beilluminated prior to intercepting a preselectedaltitude, or if for any reason the aircraft straysbeyond a preset limit from the selected altitudeonce the aircraft has intercepted the altitude.

Altitude preselector—An autopilot flight directorsubsystem that allows a pilot to preselect the alti-tude to which he desires to climb or descend. Thecontrolling mechanism for an altitude preselectsystem is normally combined with the samedevice which controls the altitude alerter system.

Amplifier—A basic type of electronic device thatseeks to make an electrical signal greater instrength. A public address system, for instance, isa type of amplifier. Amplifying devices are typi-cally tubes or transistors.

Analog—A type of electronic circuitry that ischaracterized by smooth, continuous operationrather than discrete steps, as would be observedwith digitally operated equipment.

APP-6


Angle-of-attack (AOA) indicator—A supplemen-tal flight instrumentation system that attempts toread out to the pilot the angle-of-attack or deckangle information. Several variations of this sys-tem are available.

Angular deviation—A means of showing dis-placement from a selected course either to orfrom a VOR station, TACAN, or NDB, showingdisplacement from the desired course in terms ofangle. This is commonly used with the VOR sys-tem having a normal course width of 10° on eachside of the course.

Annunciator—An indicator light with a message.An annunciator makes an announcement as to thespecific status of a system or subsystem.

ANT (antenna) mode—This mode of ADF oper-ation allows improved audio fidelity in order tolisten to the music or voice programs of an AMbroadcast station. However, the ADF needleoperation is defeated in this mode of operation.

Area navigation system (RNAV)—A system ofdirect point-to-point navigation having four fur-ther subclassifications:

●

Course line computer—A computer, uti-l iz ing informat ion f rom VOR andcolocated DME stations, that allows theoperator to change the location of theVOR station from its physical position towherever the operator wants.

●

OMEGA/VLF system—See relateddefinition.

●

Inertial navigation system—See relateddefinition.

●

Loran system—Operationally similar toan OMEGA/VLF system.

Asymptotic—A design characteristic of an auto-pilot or flight director system. The function ofthis characteristic is to allow the autopilot toattempt to intercept a given course or altitudewithout overshooting. This is done by continu-ously reducing the intercept angle as the aircraftapproaches the selected ground track or altitude.

Attenuation—The process of electrically reduc-ing the size of a radio or audio signal (i.e., to turndown or make smaller.)

Attitude director indicator (ADI) (flight directorindicator)—This instrument combines the basicfunctions of an attitude indicator with the steer-ing commands received from the flight directorsystem.

Attitude indicator (artificial horizon)—A gyro-scopically controlled instrument used to displaythe aircraft’s pitch and roll attitude relative to theearth’s surface. The gyro used to display thisinformation may be contained within the case ofthe displayed instrument, or it may receive itsinformation from a remotely located attitudegyro.

Audio filters—An electronic means of removinga portion of the audio which the pilot does notdesire to listen to. The pilot may choose toremove either the voice portion or the Morsecode identifier of a VOR or an ADF system.

Audio selector switches—The system ofswitches which allows one or several audio sys-tems to be “piped- in” to the speaker orheadphones of an aircraft.

Audio system—The electronic system that servesas a switchboard and amplification system for thevaried receivers that require the audio to be fun-neled to the speaker or headphones.

Autopilot/flight director modes:

●

Vertical modes—Vertical modes controlchanges in the pitch attitude of the air-craft using the elevator servo. Examplesof vertical modes are: altitude hold, alti-tude preselect, indicated airspeed hold,and vertical speed hold.

●

Lateral modes—Lateral modes controlautopilot operation by controlling theaileron and rudder servos. Examples oflateral modes are: heading hold, naviga-tion modes and submodes (e.g., en routenav tracking, approach, backcourse, etc.).


APP-7

Autopilot/flight director submode—Generallyspeaking, this concept represents two subclassifi-cations of operation within a given mode.Namely:

●

Arm—The process of activating a systemor preparing it to operate at a future time.For instance, if you push the Nav buttonto track a specific radial from a VOR sta-tion, but the CDI needle is displaced full-scale to the left or right at the momentyou push the Nav button, then the autopi-lot flight director system will initially beactivated in the nav-arm mode while theaircraft continues to intercept the selectedradial.

●

Capture—A submode allowing the auto-pilot flight director system to track aspecified lateral or vertical reference(e.g., altitude or glide slope as a verticalmode; VOR or localizer course as a lat-eral mode).

Avionics master circuit breaker—The circuitbreaker that supplies power to the avionics mas-ter switch in Beechcraft factory-installedavionics packages on Baron, Bonanza, Duke, andKing Air installations. This circuit breaker servesas a backup means of activating the avionics sys-tem should the avionics master switch fail for anyreason.

Avionics master switch—A central on/off powerswitch for the entire avionics package in an air-craft. This switch conveniently allows the pilot toturn on the entire avionics package by turning ononly one switch.

Beat frequency oscillator (BFO)—A device usedon an ADF receiver that generates a tone allow-ing the pilot to identify the Morse code beingtransmitted by some nondirectional beacons.

NOTE

Th i s type o f t r ansmi t t e r i s no temployed in the United States.

Carrier—That portion of the transmitted radioenergy which “carries” the useful information(i.e., modulation).

Compass system slaving—The process of auto-matically aligning the directional gyro in acompass system with the earth’s magnetic field todisplay the aircraft’s magnetic heading. When thecompass system is initially powered, slavingoccurs at a fast rate to quickly align the compasssystem with magnetic north. Once the fast-slav-i ng r a t e i s a ccompl i shed , t he sy s t emautomatically goes into a slow-slaving rate forcontinuous operation. It will correct for preces-sion errors of the compass system up to amaximum error of about 3° per minute.

Concentric—Two or more knobs mounted onone common system of shafts having the sameaxis. For example, most frequency selector knobsused in all avionics systems employ concentricknobs in the interest of conserving panel space.

Course deviation indicator (CDI)—An indicatorused with a VOR/localizer receiver that showsonly left/right deviation and to/from information.This instrument has a knob called an OBS knob,meaning “omni bearing selector,” which allowsthe pilot to choose the course to or from a VORstation.

Course knob—The name applied to the omnibearing selector on an HSI type of instrument.The course knob is attached to the resolver and acourse pointer on the HSI indicator.

Course width—Displacement left or right of thedesired course:

●

Angular—Degrees left or right of thedesired course. Course width using theVOR system is 10° on each side of thedesired course.

●

Linear—In the “en route mode” mostcourse line computer RNAVs have acourse width of ±5 nautical miles. In theapproach mode most course line com-puter RNAVs have a course width of±1 1/4 nautical miles on each side of thecenterline.

APP-8


Cross sidetone—Sending sidetone audio acrossthe cockpit from one side to the other; for exam-ple, this allows the pilot to hear what the copilotis saying on the transmitter. Cross sidetone maybe heard through either the phones or thespeaker.

Digital—A type of electronic circuitry technol-ogy that operates in specific steps, as opposed tothe smooth, sweeping type of operat ionemployed in analog.

Double-cue flight director system (crosspointer)—A command presentation system usingone vertical bar to indicate commanded roll-atti-tude instructions and one horizontal bar toindicate commanded pitch instructions.

Electronic flight instrument system (EFIS)—Atype of flight instrumentation system employingcathode ray tubes (television screens) to displayinformation.

Emergency/normal switch—In the event of thefailure of the audio system, this switch (whenplaced in the emergency position) allows audiofrom the aircraft receivers to be “piped” directlyto the headphones.

Fast erect—A mode of operation whereby anattitude indicator may be quickly realigned withthe earth’s horizon if for any reason the gyro hasprecessed or tumbled.

Flux valve (flux gate)—A component of a slavedcompass system that senses the earth’s magneticfield and converts this information into an electri-cal signal representing magnetic north.

Free operation—A mode of operation for aslaved compass system whereby the directionalgyro is disconnected from the slaving system.Normally this would be used when the slavingsystem fails or for operation in the polar regionswhere the earth’s magnetic field will not permitnormal slaved operation. The concept here is thatthe directional gyro is free of its master, magneticnorth.

Gain—The relative amount of amplification of aradio receiver. A gain control is commonly used

on a radar indicator to control the relative amountof amplification of the received radar echo. Thisallows the pilot to optimize the information dis-played, especially when the radar is used forterrain mapping purposes.

Go-around mode—An autopilot flight directormode intended to be used during a missedapproach. This mode will command a pitch-upattitude appropriate for a climbout with an asso-ciated wings-level command. The autopilot mayor may not remain engaged during the go-aroundmode, depending upon the type of autopilotinstalled in a specific aircraft.

Gyro erection—The process of an attitude gyrobecoming aligned with the earth’s horizon or,viewed in another way, aligned with true vertical.This happens automatically when the system firstreceives power.

Half bank—An autopilot mode of operationwhereby the bank angle is limited during turns insuch a way that the aircraft will only bankapproximately half as much as normal. This isdesigned to give the passengers the perception ofa smoother ride with no steep banks.

Heading bug—An adjustable marker used on aheading indicator to direct an autopilot and/orflight director system according to the magneticheading the pilot desires to fly. Also, this devicemay be used simply as a reminder to the pilot ofwhat heading he is to fly when not using the auto-pilot flight director system.

Heading indicator (directional gyro)—A gyro-scopically controlled instrument used to displayan aircraft’s heading relative to magnetic north.The compass card of this indicator may bereceiving the information which it displays froma remotely located gyro and an associated slavedcompass system.

Hertz—The unit of measure used to describe thenumber of cycles of alternating current persecond.

Horizontal situation indicator (HSI)—Thisinstrument, alternately called a CDI by somemanufacturers, displays heading information


APP-9

from a compass system, left/right and to/frominformation from a VOR/localizer receiver, anddeviation above and below a glide slope from aglide-slope receiver. The pilot’s workload isreduced by integrating these displays onto oneindicator.

Inertial navigation system—This system allowsdirect point-to-point navigation via a great circleroute. This system is completely self-sufficient,utilizing a group of gyros and accelerometers tosense movement along the earth’s surface.

Integrated autopilot/flight director system—Asystem utilizing both autopilot and flight direc-tion information to respond to selected modes.

Interrogation—In the secondary surveillanceradar system the ground-based radar unit is saidto “interrogate” the transponders of all aircraftflying within reception range of that radar. Oncea transponder has been interrogated, it shouldreply to the ground radar unit by sending a brieftransmission of radio energy. For general avia-tion aircraft a transponder may be interrogated inboth modes. A mode supplies azimuth and dis-tance information, and altitude information isprovided through mode C.

Keying—The process of turning on the transmit-ter by means of the push-to-talk button locatedon the microphone or the control wheel.

Latitude—The angular displacement of a geo-graphic location north or south of the equator.This is normally expressed in terms of degrees,minutes, and tenths of minutes.

Linear deviation—A means of showing lateraldisplacement from the desired navigationalcourse calibrated in miles. Linear deviationallows for parallel course boundaries whether faraway from or near a station.

Longitude—The angular displacement of a geo-graphic location east or west of the primemeridian located in Greenwich, England. This isnormally expressed in terms of degrees, minutes,and tenths of minutes.

Magnetic bearing—The direction of a nondirec-tional beacon (NDB) or VOR station relative tomagnetic north.

Meter movement—An application of an ammeterused in any instrumentation system to show devi-ation such as left/right, to/from, slaving indicator,etc.

Mode—One of several operating conditions of asystem. For instance, most airborne weatherradars have both weather mapping and terrainmapping modes of operation.

Mode A—That portion of the transponder replywhich transmits azimuth and distance informa-tion for the radar controller.

Mode C—The portion of a transponder replycontaining the pressure altitude of an aircraft asprovided by an encoding altimeter.

Modulation—The addition of useful informationto the carrier wave that is emitted from a trans-mitter; for example, talking into the microphoneor the transmission of the Morse code identifica-tion from a VOR station.

Muting—The silencing of incoming receiveraudio while one is transmitting.

Nonintegrated AP/FD system—Two separateflight control systems, each using its own com-puter. Information coming from these twosystems may or may not agree at any given time.

OMEGA/VLF system—A world-wide naviga-tion system that allows direct great circle flightfrom one point to another. This system utilizesU.S. Navy VLF communication transmitters andthe OMEGA system of navigation.

Parallax error—A problem that can cause inaccu-rate interpretations of an instrument reading. It iscaused by the user’s viewing angle not beingdirectly in line with the instrument.

Parallax error adjustment—An adjustment of somesingle-cue flight director systems which allows thecommand bars to be adjusted up or down in orderto “nestle” just above the aircraft symbol.

APP-10


Parked or stowed ADF needle—The process ofplacing the ADF needle at the 3 o’clock relativebearing location to indicate that the ADF unit isin the antenna mode and that the needle is notoperating.

Radio magnetic indicator (RMI)—The combineddisplay of magnetic heading from a compass sys-tem and relative bearing to a nondirectionalbeacon or VOR, which results in displaying theproduct, called “magnetic bearing,” to or fromthe station.

Range filter—An audio filter designed to removethe Morse code identification from a radio trans-mission. Actually, range is something of amisnomer as pilots know it today—you shouldthink of this as being an “ident” filter.

Relative bearing—The direction of a nondirec-tional beacon relative to the longitudinal axis ofthe aircraft.

Relay—An application of an electromagnet toperform switching duties. A relay may be used toswitch large quantities of current. A multiple polerelay will allow a single pole switch to switchmany circuits from a remote location.

Remote mounted avionics—Avionic equipmentwhich is not fully self-contained and mounted onthe instrument panel. Typically, the “blackboxes” for these systems are located in the for-ward avionics compartment, forward of the frontpressure bulkhead, or aft of the rear pressurebulkhead.

Resolver—The electronic device to which thecourse knob or OBS knob is attached. Thisdevice communicates the desired course, whichthe pilot selects, to the VOR receiver.

Servo system—Using an electric motor in anyone of several applications to reduce pilot work-load or allow automatic operation of somesystems; for example: autopilot servos, electricelevator trim servos, servoed altimeters, compasssystems, etc.

Sidetone—The ability to hear oneself talk whiletransmitting. The sidetone may be heard through

either the headphones or the speaker. Addition-ally, sidetone may be considered as a means ofverifying normal transmitter and receiver opera-tion. If the receiver and transmitter are workingproperly, the sidetone will sound “normal.” Ifeither the transmitter or receiver is malfunction-ing, the sidetone will sound weak or garbled.

Single-cue flight director system (V-bar sys-tem)—A command display system using a pairof bars which work in unison to display the com-manded attitude to the pilot.

Slant/range correction—A means of correctingfor the inherent error in raw slant/range datawhich will result in a true lateral distance fromthe aircraft to the DME station. Many of themore sophisticated RNAV computers provideslant/range correction.

Slant/range distance—Conventional, uncor-rected DME distance to the station.

Slaved compass system—A directional gyro sys-tem that is automatically synchronized to themagnetic heading of the aircraft. The concept ofthis system is that magnetic north is the master;therefore, the compass system is its slave.

Slaved operation—The normal mode of a slavedcompass system whereby the directional gyroautomatically remains synchronized to magneticnorth. This type of operation continually com-pensates for gyro precession and other compasssystem errors. The concept is that the compasssystem is a slave to magnetic north.

Slaving amplifier—An amplifier which takes theweak signal representing magnetic north, comingfrom the flux valve, and boosts that signal to ausable level in order to drive the directional gyroto the proper magnetic heading.

Slaving indicator—A meter used in some slavedcompass systems that displays the differencebetween sensed magnetic heading and displayedmagnetic heading. If the needle on this indicatoris centered, there is no error between sensed andindicated magnetic heading. If the needle is off tothe left or right, a small amount of error is indi-cated. Normal operation of the compass system


APP-11

causes the needle to sway to the left and rightbecause of gyro precession and other factors.

Soft ride—A mode for an autopilot whereby theresponsiveness of the autopilot to rough air isaltered in such a way that the ride is perceived tobe much smoother than it is.

Squelch—A silencing circuit employed in com-munication receivers that allows undesirablebackground noise to be omitted. Only a strongincoming signal from a transmitter will be heard.

Transponder code—A specific four-digit codethat may be selected by the pilot on his transpon-der to identify his specific aircraft.

NOTE

A common misconception is that thiscontrol changes the transponder replyfrequency. The transponder alwaysoperates on the same frequency.

Voice filter—An audio filter designed to removethe vo i ce po r t i on o f a r e ce ived r ad iotransmission.

Voice terrain advisories—Voice callouts of perti-nent altitude-above-ground information. Theinformation announced will be determined by thetype of system installed.

Waypoint—The geographic location of naviga-tional fix used in area navigation. This may beused in either a VLF/OMEGA system or aVOR/DME system utilizing a course linecomputer.

Waypoint address—The radial and distance of awaypoint from a VORTAC.

Waypoint coordinates—The latitude and longi-tude of the waypoint used with a VLF/OMEGAsystem.

Yaw damper—A system connected to the rudderservo that seeks to dampen or reduce oscillationsof the aircraft about the yaw axis. The yawdamper system significantly reduces the level ofmotion sickness experienced by passengers. This

system should be engaged soon after takeoff andunder normal operations should remain engageduntil just prior to landing.


ANN-1

ANNUNCIATORS

The Annunciators section presents acolor representation of all the annunci-ator lights in the airplane.

Please unfold to the right and leave openfor ready reference as the annunciatorsare cited in the text.


ANN-3

Figure ANN-1 Annunciators—King Air C90B (SNs LJ-1353 to LJ-1533)

Documents

King Air C90 A-B