TM 55-1510-220-10 WARNING DATA TECHNICAL MANUAL … · TM 55-1510-220-10 C 6 URGENT CHANGE...

TM 55-1510-220-10

TECHNICAL MANUAL

OPERATOR'S MANUALFOR

ARMY RC-12G AIRCRAFT

This copy is a reprint which includes currentpages from Changes 1 through 6.

HEADQUARTERS, DEPARTMENT

OF THE ARMY10 December 1985

WARNING DATA

TABLE OF CONTENTS

INTRODUCTION

DESCRIPTION ANDOPERATION

AVIONICS

MISSION EQUIPMENT

OPERATING LIMITS ANDRESTRICTIONS

WEIGHT/BALANCE ANDLOADING

PERFORMANCE DATA

NORMAL PROCEDURES

EMERGENCY PROCEDURES

REFERENCES

ABBREVIATIONS ANDTERMS

ALPHABETICAL INDEX

URGENTTM 55-1510-220-10

C7

CHANGE

NO. 7

HEADQUARTERSDEPARTMENT OF THE ARMY

WASHINGTON, D.C., 20 April 1998

OPERATOR’S MANUALFOR


DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited

TM 55-1510-220-10, 10 December 1985, is changed as follows:

1. Remove and insert pages as indicated below. New or changed text material is indicated by verticalbar in the margin. An illustration change is indicated by a miniature pointing hand.

Remove pages

i and ii5-9 and 5-10

8-31 and 8-32

INDEX-3 and INDEX-4

Insert pages

i and ii5-9 and 5-105-10A/(5-10B blank)8-31 and 8-328-32A/(8-32B blank)INDEX-3 and INDEX-4

2. Retain this sheet in front of manual for reference purposes.

By Order of the Secretary of the Army:

Official:

J O E L B . H U D S O NAdministrative Assistant to the

Secretary of the Army04633

DENNIS J. REIMERGeneral, United States Army

Chief of Staff

DISTRIBUTION:To be distributed in accordance with Initial Distribution Number (IDN) 311135, requirements for TM 55-

1510-220-10.

TM 55-1510-220-10C 6

URGENT

CHANGE HEADQUARTERSDEPARTMENT OF THE ARMY

NO. 6 WASHINGTON, D.C., 7 August 1992

Operator's ManualFor



1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in themargin. An illustration change is indicated by a miniature pointing hand.

Remove pages Insert pages

5-9 and 5-10 5-9 and 5-10- - - - 5-10A/5-10B8-17 through 8-22 8-17 through 8-228-31 and 8-32 8-31 and 8-32



GORDON R. SULLIVANGeneral, United States Army

Official: Chief of Staff

MILTON H. HAMILTONAdministrative Assistant to the


DISTRIBUTION:To be distributed in accordance with DA Form 12-31-E, block no. 1135, -10 & CL maintenance requirements for

TM 55-1510-220-10.

DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.

URGENT

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TM 55-1510-220-10C5


NO. 5 WASHINGTON, D.C., 30 June 1992


Army RC-12G Aircraft




2-78A/2-78B 2-78A/2-78B3-47 and 3-48 3-47 and 3-483-61 and 3-62 3-61 and 3-623-75 and 3-76 3-75 and 3-76- - - - 3-76A through 3-76E/3-76F4-1 and 4-2 4-1 and 4-2Index 3 and Index 4 Index 3 and Index 4





MILTON H. HAMILTONAdministrative Assistant to the



TM 55-1510-220-10.

}

TM 55-1510-220-10C 4

URGENT


NO. 4 WASHINGTON, D.C., 20 August 1991






6-1 and 6-2 6-1 and 6-26-5 and 6-6 6-5 and 6-6


By Order or the Secretary of the Army:



PATRICIA P. HICKERSONColonel, United States Army

The Adjutant General


TM 55-1510-220-10.

URGENT

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TM 55-1510-220-10C3

URGENTNOTICE: THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE

INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCEDCHANGE ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENTPAGE.


No. 3 WASHINGTON, D.C.,.9 November 1990






9-11 and 9-12 9-11 and 9-12- - - - 9-12.1/9-12.2



CARL E. VUONOGeneral, United States Army


THOMAS F. SIKORABrigadier General, United States Army


DISTRIBUTION:To be distributed in accordance with DA Form 12-31, -10 & CL Maintenance requirements for RC-12D Airplane,

Reconnaissance.

URGENT

}

TM 55-1510-220-10C2


NO. 2 WASHINGTON, D.C., 1 November 1990

Operator's Manualfor





a and b a and bi and ii i and ii1-1 and 1-2 1-1 and 1-22-1 through 2-10 2-1 through 2-102-15 and 2-16 2-15 and 2-162-19 through 2-30 2-19 through 2-302-39 and 2-40 2-39 and 2-402-43 through 2-48 2-43 through 2-482-55 through 2-62 2-55 through 2-622-65 through 2-68 2-65 through 2-682-71 through 2-78 2-71 through 2-78- - - - 2-78A/2-78B2-79 through 2-88 2-79 through 2-883-1 through 3-4 3-1 through 3-43-9 through 3-18 3-9 through 3-183-21 through 3-26 3-21 through 3-263-29 through 3-38 3-29 through 3-383-41 through 3-44 3-41 through 3-443-49 through 3-78 3-49 through 3-784-1 through 4-6 4-1 through 4-64-9 and 4-10 4-9 and 4-105-1 and 5-2 5-1 and 5-25-7 through 5-12 5-7 through 5-126-1 through 6-14 6-1 through 6-127-1 through 7-4 7-1 through 7-47-7 through 7-9/7-10 7-7 through 7-9/7-10- - - - 7-24A/7-24B- - - - 7-54A through 7-54R7-55 and 7-56 7-56- - - - 7-90A through 7-90P- - - - 7-95 and 7-968-1 and 8-2 8-1 and 8-29-15 and 9-16 9-15 and 9-16Index 1 through Index 6 Index 1 through Index 6


}

TM 55-1510-220-10C 2

By Order of the Secretary of the Army:CARL E. VUONO

General, United States ArmyOfficial: Chief of Staff

THOMAS F. SIKORABrigadier General, United States Army



TM 55-1510-220-10.

TM 55-1510-220-10C 1

URGENT


NO. 1 WASHINGTON, D.C., 9 October 1986

Operator's Manual

for


TM 55-151U-2z0-10, 10 December 1985, is changed as follows:



6-1 and 6-2 6-1 and 6-2


By Order of the Secretary of the Army:JOHN A. WICKHAM, JR.

General, United States ArmyOfficial: Chief of Staff

R. L. DILWORTHBrigadier General, United States Army


DISTRIBUTION:To be distributed in accordance with DA Form 12-31, -10 and CL Maintenance requirements for RC-12D

Airplane, Reconnaissance.

URGENT

TM 55-1510-220-10WARNING PAGE

Personnel performing operations, procedures and practices which are included or implied in this technicalmanual shall observe the following warnings. Disregard of these warnings and precautionary information can causeserious injury or loss of life.

NOISE LEVELSSound pressure levels in this aircraft during some operating conditions exceed the Surgeon General's hearing

conservation criteria, as defined in TM MED 501. Hearing protection devices, such as the aviator helmet or ear plugsshall be worn by all personnel in and around the aircraft during its operation.

OPERATING PROCEDURESOperating procedures or practices defined in this Technical Manual must be followed correctly. Failure to do so

may result in personal injury or loss of life.Exposure to exhaust gases shall be avoided since exhaust gases are an irritant to eyes, skin and respiratory

system.

HIGH VOLTAGEHigh voltage is a possible hazard around AC inverters, ignition exciter units,; and strobe beacons.

USE OF FIRE EXTINGUISHERS IN CONFINED AREASMonobromotrifluoromethane (CF3Br) is very volatile, but is not easily detected by its odor. Although non toxic, it

must be considered to be about the same as other freons and carbon dioxide, causing danger to personnel primarily byreduction of oxygen available for proper breathing. During operation of the fire extinguisher, ventilate personnel areaswith fresh air. The liquid shall not be allowed to come into contact with the skin, as it may cause frostbite or lowtemperature burns because of its very low boiling point.

VERTIGOThe strobe/beacon lights should be turned off during flight through clouds to prevent sensations of vertigo, as a

result of reflections of the light on the clouds.

CARBON MONOXIDEWhen smoke, suspected carbon monoxide fumes, or symptoms of lack of oxygen (hypoxia) exist, all personnel

shall immediately don oxygen masks, and activate the oxygen system.

FUEL AND OIL HANDLINGTurbine fuels and lubricating oils contain additives which are poisonous and readily absorbed through the skin.

Do not allow them to remain on skin.

SERVICING AIRCRAFTWhen conditions permit, the aircraft shall be positioned so that the wind will carry fuel vapors away from all

possible sources of ignition. The fueling unit shall maintain a distance of 20 feet between unit and filler point. Aminimum of 10 feet shall be maintained between fueling unit and aircraft.

Prior to refueling, the hose nozzle static ground wire shall be attached to the grounding lugs that are locatedadjacent to filler openings.

Change 2 a

TM 55-1510-220-10SERVICING BATTERY

Improper service of the nickel-cadmium battery is dangerous and may result in both bodily injury and equipmentdamage. The battery shall be serviced in accordance with applicable manuals by qualified personnel only.

Corrosive Battery Electrolyte (Potassium Hydroxide). Wear rubber gloves, apron, and face shield when handlingbatteries. If potassium hydroxide is spilled on clothing, or other material wash immediately with clean water. If spilled onpersonnel, immediately start flushing the affected area with clean water. Continue washing until medical assistancearrives.

JET BLASTOccasionally, during starting, excess fuel accumulation in the combustion chamber causes flames to be blown

from the exhausts. This area shall be clear of personnel and flammable materials.

RADIOACTIVE MATERIALInstruments contained in this aircraft may contain radioactive material (TB 55-1500-314-25). These items

present no radiation hazard to personnel unless seal has been broken due to aging or has accidentally been broken. Ifseal is suspected to have been broken, notify Radioactive Protective Officer.

RF BURNSDo not stand near the antennas when they are transmitting.

OPERATION OF AIRCRAFT ON GROUNDA1 all times during a towing operation, be sure there is a man in the cockpit to operate the brakes.

Personnel should take every precaution against slipping or falling. Make sure guard rails are installed whenusing maintenance stands.

Engines shall be started and operated only by authorized personnel. Reference AR 95-1.

Insure that landing gear control handle is in the DN position.b

TECHNICAL MANUAL

NO.55-1520-220-10

TM 55-1510-220-10

HEADQUARTERSDEPARTMENT OF THE ARMY

WASHINGTON, D.C., 10 December 1985

Operator’s Manual

ARMY MODEL RC-12G

REPORTING OR ERRORS AND RECOMMENDING IMPROVEMENTS

You can help improve this manual. If you find any mistakes or if you know of any way to improve the procedures,please let use know. Mail your letter, DA Form 2028 (Recommended Changes to Publications and Blank Forms), orDA Form 2028-2 located in the back of this manual directly to: Commander, U.S. Army Aviation and MissileCommand, ATTN: AMSAM-MMC-LS-LP, Redstone Arsenal, AL 35898-5230. A reply will be furnished directlyto you. You may also send in your comments electronically to our E-mail address at , orby fax at (205) 842-6546 or DSN 788-6546. Instructions for sending an Electronic DA Form 2028 may be found atthe back of this manual immediately preceding the hard copy DA Forms 2028.

CHAPTER 1.CHAPTER 2.

Section

CHAPTER 3.Section

CHAPTER 4.Section

CHAPTER 5.Section

CHAPTER 6.Section

I.II.

III.IV.

V.VI.

VII.VIII.

IX.X.

XI.XII.

I.II.

III.IV.

I.II.

I.II.

III.IV.V.

VI.VII.

VIII.IX.

I.II.

III.IV.

V.

TABLE OF CONTENTS

INTRODUCTIONAIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATIONAircraftEmergency equipmentEngines and related systemsFuel systemsFlight controlsPropellersUtility systemsHeating, ventilation, cooling and environmentalcontrol systemElectrical power supply and distribution systemLightingFlight instrumentsServicing, parking and mooringAVIONICSGeneralCommunicationsNavigationTransponder and radarMISSION EQUIPMENTMission avionicsAircraft survivability equipmentOPERATING LIMITS AND RESTRICTIONSGeneralSystem limitsPower limitsLoading limitsAirspeed limits, maximum and minimumManeuvering limitsEnvironmental restrictionsOther limitationsRequired equipment for various conditions of flightWEIGHT/BALANCE AND LOADINGGeneralWeight and balanceFuel/oilCenter of gravityCargo loading

Page

1-12-12-12-202-212-282-352-392-42

2-542-592-682-722-803-13-13-23-213-774-14-14-15-15-15-15-65-85-85-105-105-115-126-16-16-16-126-126-12

C h a n g e 7 i

TM 55-1510-220-10

Page

7-18-18-18-28-258-268-298-339-19-1A-1B-1

INDEX-1

CHAPTER 7.CHAPTER 8.

Section

CHAPTER 9.Section

APPENDIX A.APPENDIX B.INDEX

PERFORMANCE DATANORMAL PROCEDURES

I. Mission planningII. Operating procedures and maneuvers

III. Instrument flightIV. Flight characteristicsV. Adverse environment conditions

VI. Crew dutiesEMERGENCY PROCEDURES

I. Aircraft systemsREFERENCESABBREVIATIONS AND TERMS

ii

TM 55-1510-220-10CHAPTER 1

INTRODUCTION

1-1. GENERAL.

These instructions are for use by theoperator(s). They apply to the RC-12G aircraft.

1-2. WARNINGS, CAUTIONS, AND NOTES.

Warning cautions, and notes are used toemphasize important and critical instructions and areused for the following conditions:

WARNINGAn operating procedure, practice,etc., which if not correctlyfollowed, could result in personalinjury or loss of life.

CAUTIONAn operating procedure, practice,etc., which, if not strictlyobserved, could result in damageto or destruction of equipment.

NOTEAn operating procedure,condition, etc., which is essentialto highlight.

1-3. DESCRIPTION.

This manual contains the best operatinginstructions and procedures for the RC-12G aircraftunder most circumstances. The observance oflimitations, performance, and weight/ balance dataprovided is mandatory. The observance of proceduresis mandatory except when modification is requiredbecause of multiple emergencies, adverse weather,terrain, etc. Your flying experience is recognized, andtherefore, basic flight principles are not included. THISMANUAL SHALL BE CARRIED IN THE AIRCRAFT ATALL TIMES.

1-4. APPENDIX A, REFERENCES.

Appendix A is a listing of official publicationscited within the manual applicable to and available forflight crews.

1-5. APPENDIX B, ABBREVIATIONS AND TERMS.

Appendix B is a listing of abbreviations andterms used throughout the manual.

1-6. INDEX.

The index lists, in alphabetical order, everytitled paragraph, figure, and table contained in thismanual. Chapter 7, Performance Data, has anadditional index within the chapter.

1-7. ARMY AVIATION SAFETY PROGRAM.

Reports necessary to comply with the safetyprogram are prescribed in AR 385-40.

1-8. DESTRUCTION OF ARMY MATERIEL TOPREVENT ENEMY USE.

For information concerning destruction of Armymateriel to prevent enemy use, refer to TM 750-244-1-5.

1-9. FORMS AND RECORDS.

Army aviators flight record and aircraftmaintenance records which are to be used by crewmembers are prescribed in DA PAM 738-751 andWeight I and Balance TM 55-1500-342-23.

1-10. EXPLANATION OF CHANGE SYMBOLS.

Changes, except as noted below, to the text andtables, including new material on added pages, areindicated by a vertical line in the outer margin extendingclose to the entire area of the material affected;exception: pages with emergency markings, whichconsist of black diagonal lines around three edges, mayhave the vertical line or change symbol placed along theinner margins. Symbols show current changes only. Aminiature pointing hand symbol is used to denote achange to an illustration. However, a vertical line in theouter margin, rather than miniature pointing hands, isutilized when there have been extensive changes madeto an illustration. Change symbols are not utilized toindicate changes in the following:

Change 2 1-1

TM 55-1510-220-10

a. Introductory material.

b. Indexes and tabular data where the changecannot be identified.

c. Blank space resulting from the deletion of text,an illustration or a table.

d. Correction of minor inaccuracies, such asspelling, punctuation, relocation of material, etc., unlesscorrection changes the meaning of instructiveinformation and procedures.

1-1 1. AIRCRAFT DESIGNATION SYSTEM.

The designation system prescribed by AR 7050is used in aircraft designations as follows:

EXAMPLE RC-12GR - Modified mission symbol

(Reconnaissance)

C - Basic mission and type symbol (cargo)12 - Design numberG - Series symbol

1-12. USE OF WORDS SHALL, WILL, SHOULD, ANDMAY.

Within this technical manual the word "shall" isused to indicate a mandatory requirement. The word"should" is used to indicate a nonmandatory butpreferred method of accomplishment. The word "may"is used to indicate an acceptable method ofaccomplishment. The word "will" is used to express adeclaration of purpose and may also be used wheresimple futurity is required.

1-13. PLACARD ITEMS.

All placard items (switches, controls, etc.) areshown throughout this manual in capital letters.

1-2

TM 55-1510-220-10CHAPTER 2

AIRCRAFT AND SYSTEMS DESCRIPTION AND OPERATIONSection I. AIRCRAFT

2-1. INTRODUCTION.

The purpose of this chapter is to describe theaircraft and its systems and controls which contribute tothe physical act of operating the aircraft. It does notcontain descriptions of avionics or mission equipment,covered elsewhere in this manual. This chaptercontains descriptive information and does not describeprocedures for operation of the aircraft. Theseprocedures are contained within appropriate chapters inthe manual. This chapter also contains the emergencyequipment installed. This chapter is not designed toprovide instructions on the complete mechanical andelectrical workings of the various systems; therefore,each is described only in enough detail to makecomprehension of that system sufficiently complete toallow for its safe and efficient operation.

2-2. GENERAL.

The RC-12G is a pressurized, low wing, allmetal aircraft, powered by two PT6A-41 turbopropengines (fig. 2-1 and 2-2), and has all weathercapability. Distinguishable features of the aircraft arethe slender, streamlined engine nacelles, missionantennas, wing tip pods, a T-tail and a ventral fin belowthe empennage. The basic mission of the aircraft isradio reconnaissance. Cabin entrance is made througha stair-type door on the left side of the fuselage.

2-3. DIMENSIONS.

Overall aircraft dimensions are shown in figure2-3.

2-4. GROUND TURNING RADIUS.

Minimum ground turning radius of the aircraft isshown in figure 2-4.

2-5. MAXIMUM WEIGHTS.

Maximum take of gross weight is 15,000pounds. Maximum landing weight is 15,000 pounds.Maximum ramp weight is 15,090 pounds. Maximumzero fuel weight is 11,500 pounds.

2-6. EXHAUST DANGER AREA.

Danger area to be avoided by personnel whileaircraft engines are being operated on the ground aredepicted in figure 2-5. Distance to be maintained withengines operating at idle are also shown. Temperatureand velocity of exhaust gasses at varying locations aftof the exhaust stacks are shown for maximum power.The danger area extends to 40 feet aft of the exhauststack outlets. Propeller danger areas are also shown.

2-7. LANDING GEAR SYSTEM.

The landing gear is a retractable, tricycle type,electrically operated by a single DC motor. This motordrives the main landing gear actuators through a gearbox and torque tube arrangement, and also drives achain mechanism which controls the position of the nosegear. Positive down-locks are installed to hold the dragbrace in the extended and locked position. The down-locks are actuated by overtravel of the linear jackscrewsand are held in position by a spring-loaded overcentermechanism. The jackscrew in each actuator holds allthree gears in the UP position, when the gear isretracted. A friction clutch between the gearbox and thetorque shafts protects the motor from electrical overloadin the event of a mechanical malfunction. A 150-ampere current limiter, located on the DC distributionbus under the center floorboard, protects againstelectrical overload. Gear doors are opened and closedthrough a mechanical linkage connected to the landinggear. The nose wheel steering mechanism isautomatically centered and the rudder pedals relieved ofthe steering load when the landing gear is retracted.Air-oil type shock struts, filled with compressed air andhydraulic fluid, are incorporated with the landing gear.Gear retraction or extension time is approximately sixseconds.

a. Landing Gear Control Switch. Landing gearsystem operation is controlled by a manually actuated,wheel-shaped switch placarded LDG GEAR CONTR UPDN, located on the left subpanel (fig. 2-6). The controlswitch and associated relay circuits are protected by a5ampere circuit breaker, placarded LANDING

2-1

TM 55-1510-220-10

Figure 2-1. General Exterior Arrangement (Sheet 1 of 6)

GEAR RELAY on the overhead circuit breaker panel(fig. 2-27).

b. Landing Gear Down Position-Indicator Lights.Landing gear down position is indicated by three greenlights on the left subpanel, placarded GEAR DOWN.These lights may be checked by operating theANNUNCIATOR TEST switch. The circuit is protectedby a 5-ampere circuit breaker, placarded LANDINGGEAR IND, on the overhead circuit breaker panel (fig.2-27).

c. Landing Gear Position Warning Lights. Two redbulbs, wired in parallel and activated by microswitchesindependent of the GEAR

DOWN position indicator annunciators, are positionedinside the clear plastic grip on the landing gear controlswitch. These annunciators illuminate whenever thelanding gear switch is in either the UP or DN positionand the gear is in transit. Both bulbs will also illuminateshould either or both power levers be retarded belowapproximately 79 to 81% N1 when the landing gear isnot down and locked. To turn the switch annunciatorsOFF during single engine operation, the power lever forthe inoperative engine must be advanced to a positionwhich is higher than the setting of the warning hornmicroswitch. Extending the landing gear will also turnthe annunciators off, provided the gear is down andlocked. Both red annunciators indicate the samewarning conditions, but two are provided for a fail-safeindication in the event one bulb burns out. The

Change 2 2-2

TM 55-1510-220-10


circuit is protected by a 5-ampere circuit breaker,placarded LANDING GEAR IND, on the overhead circuitbreaker panel (fig. 2-27.)

d. Landing Gear Warning Light Test Button. A testbutton, placarded HDL LT TEST, is located on the leftsubpanel. Failure of the landing gear switch toilluminate red, when this test button is pressed, indicatestwo defective bulbs or a circuit fault. The circuit isprotected by a 5-ampere circuit breaker, placardedLANDING GEAR RELAY, on the overhead circuitbreaker panel (fig. 2-27).

e. Landing Gear Warning Horn When either powerlever is retarded below approximately 79 to 81% N1when the landing gear is not down and locked and if theflaps are extended beyond 40%, a warning horn, locatedin the overhead con-

trol panel will sound intermittently. To prevent thewarning horn from sounding during long descents, apressure differential "Q" switch is connected into thecopilot's static line. The switch prevents the warninghorn from sounding until airspeed drops below 140KIAS. An altitude sensing switch is installed in serieswith the 140 KIAS "Q" switch which prevents thewarning horn from sounding after climbing through12,500 feet MSL. The horn will be engaged when theaircraft descends through 10, 500 feet MSL. Thewarning horn circuit is protected by a 5-ampere circuitbreaker, placarded LANDING GEAR WARN, on theoverhead circuit breaker panel (fig. 2-27).

f. Landing Gear Warning Horn Test Switch. Thelanding gear warning horn may be tested by the testswitch on the right subpanel (fig. 2-

Change 2 2-3

TM 55-1510-220-10


6). The switch, placarded STALL WARN TEST OFFLDG GEAR WARN TEST, will sound the landing gearwarning horn and illuminate the landing gear positionwarning lights when moved to the momentary LDGGEAR WARN TEST position. The circuit is protectedby a 5-ampere circuit breaker, placarded LANDINGGEAR WARN, on the overhead circuit breaker panel(fig. 2-27).

g. Landing Gear Safety Switches. A safety switchon each main landing gear shock strut controls theoperation of various aircraft systems that function onlyduring flight or only during ground operation. Theseswitches are mechanically actuated whenever the mainlanding gear

shock struts are extended (normally after takeoff), orcompressed (normally after landing). The safety switchon the right main landing gear strut activates the landinggear control circuits, cabin pressurization circuits andthe flight hour meter when the strut is extended. Thisswitch also activates a down-lock hook, preventing thelanding gear from being raised while the aircraft is onthe ground. The hook, which unlocks automatically aftertakeoff, can be manually overridden by pressing downon the red button, placarded DN LOCK REL locatedadjacent to the landing gear switch. If the override isused and / the landing gear control switch is raised,power will be supplied to the warning horn circuit andthe horn will sound. The safety switch on the left mainlanding gear strut activates the left and

2-4

TM 55-1510-220-10


right engine ambient air shut-off valves when the strut isextended.

h. Landing Gear Alternate Engage Handle. Duringmanual landing gear extension, the landing gear motormust be disengaged from the landing gear drivemechanism. This is accomplished by a manually-operated clutch disengage lever (fig. 2-7 and 2-8)located adjacent to the landing gear alternate extensionhandle (fig. 27 and 2-8). To disengage the clutch, pullthe

alternate engage handle up and turn clockwise. Toengage the clutch, turn the alternate engage handlecounterclockwise and release.

CAUTIONContinued pumping of handleafter GEAR DOWN positionindicator lights (3) are illuminatedcould damage the drivemechanism, and preventsubsequent gear retraction.

Change 2 2-5

TM 55-1510-220-10

1. APR-39 ANTENNA 13. OUTBOARD WING FUEL SUMP DRAIN2. UHF/VHF COMM DIPOLE ANTENNA 14. VHF INTERCEPT DIPOLE ANTENNA3. GRAVITY LINE DRAIN 15. ENGINE OIL VENT4. DC EXTERNAL POWER RECEPTACLE 16. UHF DF MONOPOLE ANTENNA5. STANDBY FUEL BOOST PUMP DRAIN 17. ANTENNA DEICE SYSTEM BOOT6. BLEED AIR HEAT EXCHANGER AIR INTAKE 18. HOLD DOWN EJECTOR TUBE7. BATTERY RAM AIR VENT 19. STROBE BEACON8. APR-39 BLADE ANTENNA VHF DF LOOP ANTENNA9. HYDRAULIC RESERVOIR DRAIN 20. OXYGEN REGULATOR VENT

10. MARKER BEACON ANTENNA 21. TAIL NAVIGATION LIGHT11. AC EXTERNAL POWER RECEPTACLE 22. TAILCONE ACCESS DOOR12. STALL WARNING VANE 23. AIR DATA PROBE


i. Landing Gear Alternate Extension Handle.Manual landing gear extension is provided through amanually powered system as a backup to the electricallyoperated system. Before manually extending the gear,make certain that the landing gear switch is in the downposition with the LANDING GEAR RELAY circuit

breaker pulled. Pulling up on the alternate engagehandle, located on the floor, and turning it clockwise willlock it in that position. When the alternate engagehandle is pulled, the motor is electrically disconnectedfrom the system and the alternate drive system is lockedto the gearbox and motor. When the alternate drive islocked in,

Change 2 2-6

TM 55-1510-220-10


the chain is driven by a continuous action ratchet whichis activated by pumping the landing gear alternateextension handle adjacent to the alternate engagehandle.

(1) After a manual landing gear extension hasbeen made, do not stow the handle, move any landinggear controls, or reset any switches or circuit breakers.The gear cannot be retracted manually.

Change 2 2-7

TM 55-1510-220-10

Figure 2-2. General Interior Arrangement (Sheet 1 of 2)

j. Tires. The aircraft is equipped with dual 22 x6.75 x 10, 8 ply rated, tubeless, rim inflation tires on themain gear. The nose gear is equipped with a single 22 x6.75 x 10, 8 ply rated, tubeless, rim inflated tire.

k. Steerable Nose Wheel. The aircraft can bemaneuvered on the ground by the steerable nose wheelsystem. Direct linkage from the rudder pedals (fig. 2-7)to the nose wheel steering linkage allows the nose wheelto be turned 12° to the left of center or 14° to the right.When rudder pedal steering is augmented by the mainwheel braking action, the nose wheel can be deflectedup to 480 either side of center. Shock loads whichwould normally be transmitted to the rudder pedals areabsorbed by a spring mechanism in the steering linkage.Retraction of the landing gear automatically centers thenose wheel and disengages the steering linkage fromthe rudder pedals.

I. Wheel Brake System. The main wheels areequipped with multiple-disc hydraulic brakes actuated bymaster cylinders attached to the rudder pedals at thepilot's and copilot's position. Braking is permitted fromeither set of rudder pedals. Brake fluid is supplied to thesystem from the reservoir in the nose compartment.The toe brake sections of the rudder pedals areconnected to the master cylinders which actuate thesystem for the corresponding wheels. No emergencybrake system is provided.

WARNINGRepeated and excessiveapplication of brakes, withoutallowing sufficient time forcooling, between applications,will cause a loss of brakingefficiency, possible failure ofbrake or wheel structure, possibleblowout of tires, and in extremecases may cause the wheel brakeassembly to be destroyed by fire.

2-8. PARKING BRAKE.

Dual parking brake valves are installed belowthe cockpit floor. Both valves can be closedsimultaneously by pressing both brake pedals to build uppressure, then pulling out the handle placardedPARKING BRAKE, on the left subpanel. Pulling thehandle fully out sets the check valves in the system andany pressure being applied by the toe brakes ismaintained. Parking brakes are released when thebrake handle is pushed in. The parking brake may beset from either cockpit position.

CAUTIONParking brakes shall not be setduring flight.

Change 2 2-8

TM 55-1510-220-10

Figure 2-2. General Interior Arrangement (Sheet 2 of 2)

2-9. ENTRANCE AND EXIT PROVISIONS.

NOTETwo keys are provided in theloose tools and equipment bag.Both keys will fit the locks on thecabin door, emergency hatch,tailcone access door and the rightand left nose avionicscompartment doors.

a. Cabin Door.

CAUTIONStructural damage may be causedif more than one person is on thecabin door at a time. The door islimited to a maximum of 300pounds.

A swing-down door (fig. 2-9), hinged at thebottom, provides a stairway for normal and emergencyentrance and exit. Two of the steps are movable andfold flat against the door in the closed position. A stepfolds down over the door

sill when the door opens to provide a platform (step) fordoor seal protection. A plastic encased cable providessupport for the door in the open position, a handhold,and a convenience for closing the door from inside. Ahydraulic damper permits the door to lower graduallyduring opening. A rubber seal around the door seals thepressure vessel while the aircraft is in flight. The doorlocking mechanism is operated by either of the twomechanically interconnected handles, one inside andthe other outside the door. When either handle isrotated, three rotating cam-type latches on either side ofthe door capture posts mounted on the cargo door. Inthe closed position, the door becomes an integral part ofthe cargo door. A button adjacent to the door handlemust be depressed before the handle can be rotated toopen the door. A bellows behind the button is inflatedwhen the aircraft is pressurized to prevent accidentalunlatching and/or opening of the door. A small roundwindow just above the second step permits observationof the pressurization safety bellows. A placard adjacentto the window instructs the operator that the safety lockarm is in position around the bellows shaft whichindicates a properly locked door. Pushing the red buttonadjacent to the window will illuminate the

Change 2 2-9

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Figure 2-3. Principal Dimensions2-10

TM 55-1510-220-10

Figure 2-4. Ground Turning Radius2-11

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Figure 2-5. Exhaust and Propeller Danger Areas2-12

TM 55-1510-220-10

Figure 2-6. Subpanels2-13

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Figure 2-7. Cockpit2-14

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Figure 2-8. Control PedestalChange 2 2-15

TM 55-1510-220-10

Figure 2-9. Cabin and Cargo Doors2-16

TM 55-1510-220-10

inside door mechanism. A CABIN DOOR annunciatorlight on the caution/advisory panel will illuminate if thedoor is not closed and all latches fully locked. The cabindoor opening is 21.5 inches wide by 50.0 inches high.

b. Cargo Door. A swing-up door (fig. 2-9), hinged atthe top, provides cabin access for loading cargo or bulkyitems. After initial opening force is applied, gas springswill completely open the cargo door automatically. Thedoor is counterbalanced and will remain in the openposition. A door support rod is used to hold the door inthe open position, and to aid in overcoming the pressureof the gas spring assemblies when closing the door.Once closed, the gas springs apply a closing force toassist in latching the door. A rubber seal around the doorseals the pressure vessel while in flight. The door lockingmechanism is operated only from inside the aircraft, andis operated by two handles, one in the bottom forwardportion of the door and the other in the upper aft portion ofthe door. When the upper aft handle is operated perplacard instructions, two rotating cam-type latches on theforward side of the door and two on the aft side rotate,capturing posts mounted on the fuselage side of the dooropening. The bottom handle, when operated per placardinstructions, actuates four pin lug latches across thebottom of the door. A button on the upper aft handle mustbe pressed before the handle can be released to open orlatch the door. A latching lever on the bottom handlemust be lifted to release the handle before the lowerlatches can be opened. These act as additional aids inpreventing accidental opening or unlatching of the door.The cabin and cargo doors are equipped with dualsensing circuits to provide the crew remote indication ofcabin/cargo door security. An annunciator light placardedCABIN DOOR will illuminate if the cabin or cargo door isopen and the BATT switch in ON. If the battery switch isOFF, the annunciator will illuminate only if the cargo dooris not securely closed and latched. The cargo doorsensing circuit receives power from the hot battery bus.The cargo door opening is 52.0 inches wide by 52.0inches high.

CAUTIONInsure that the cabin door is closedand locked. Operating the cargodoor while the cabin door is openmay damage the door hinge andadjacent structure.

(1) Opening cargo door.

CAUTIONAvoid side loading of the gassprings to prevent damage to themechanism.

1. Handle access door (lower forwardcorner of door) Unfasten and open.

2. Handle Lift hook and move to OPENposition.

3. Handle access door Secure.

4. Handle access door (upper aft cornerof door) Unfasten and open.

5. Handle Press button and lift to OPENposition then latch in place.

6. Handle access door Secure.

7. Door support rod Attach one end tocargo door ball stud (on forward side of door).

8. Support rod detent pin Check in place.

9. Cabin door sill step Push out on andallow cargo door to swing open. Gas springs willautomatically open the door.

10. Door support rod Attach free end toball stud on forward fuselage door frame.

(2) Closing cargo door.CAUTION

Avoid side loading of the gassprings to prevent damage to themechanism.

1. Door support rod Detach fromfuselage door frame ball stud, then firmly grasp free endof rod while exerting downward force to overcome thepressure of gas spring assemblies, Then remove supportrod from door as gas spring assemblies pass over-centerposition.

2. Cargo door - Pull closed, using fingerhold cavity in fixed cabin door step.

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TM 55-1510-220-10

3. Handle access door (upper aft cornerof door) Unfasten and open.

4. Handle Press button and pull handledown until it latches in closed position.

5. Handle access door Secure.6. Handle access door (lower forward

corner of door) Unfasten and open.7. Handle Move to full forward position.8. Safety hook Check locked in position

by pulling aft on handle.9. Handle access door Secure.

c. Cabin Emergency Hatch. The cabin emergencyhatch, placarded EXIT PULL, is located on the right cabinsidewall just aft of the copilot's seat. The hatch may bereleased, from the inside with a pull-down handle. A flushmounted pull out handle allows the hatch to be releasedfrom the outside. The hatch is of the non-hinged plugtype, which removes completely from the frame when thelatches are released. The hatch can be key locked fromthe inside, to prevent opening from the outside. Theinside handle will unlatch the hatch whether or not it islocked, by overriding the locking mechanism. Thekeylock should be unlocked prior to flight to allow removalof the hatch from the outside in the event of anemergency. The key remains in the lock when the hatchis locked and can be removed only when the hatch isunlocked. The key slot is in the vertical position when thehatch is unlocked. Removal of the key from the lockbefore flight assures the pilot that the hatch can beremoved from the outside if necessary.

d. Cabin Door Caution Light. As a safety precaution,two illuminated MASTER CAUTION lights, on the glareshield and a steadily illuminated CABIN DOOR yellowcaution annunciator light on the caution/advisory panelindicate the cabin door is not closed and locked. Thiscircuit is protected by 5-ampere circuit breakers placardedANN PWR and ANN IND, located on the overhead circuitbreaker panel (fig. 2-27).

2-10. WINDOWS.a. Cockpit Window. The pilot and copilot have side

windows, a windshield and storm windows, which providevisibility from the cockpit. The storm windows may beopened on the ground or during unpressurized flight.

b. Cabin Window. The outer cabin windows are oftwo-ply construction, are the pressure type and areintegral parts of the pressure vessel. All cabin windowsare painted over except for the three windows locatedfarthest aft on the right side, and the two farthest aft onthe left side. The unpainted windows are provided withflaps which may be raised or lowered when it is desired toseal out light.2-11. SEATS.

a. Pilot and Copilot Seats. The pilot and copilotseats (fig. 2-10) are separated from the cabin by movablecurtains. The controls for vertical height adjustment andfore and aft travel are located under each seat. Theforward and aft adjustment handle is located beneath thebottom front inboard corner of each seat. Pulling up onthe handle allows the seat to move fore or aft. The heightadjustment handle is located beneath the bottom frontoutboard corner of each seat. Pulling up on the handle,allows the seat to move up and down. Both seats haveadjustable headrests and armrests which will raise andlower for access to the cockpit. Handholds on either sideof the overhead panels and a fold-away protectivepedestal step are provided for pilot and copilot entry intothe cockpit. For the storage of maps and the operator'smanual, pilot and copilot seats have an inboard-slanted,expandable pocket affixed to the lower portion of the seatback. Pocket openings are held closed by shock cordtension.

b. Pilot and Copilot Seat Belts and ShoulderHarnesses. Each pilot and copilot seat is equipped with alap-type seat belt and shoulder harness connected to aninertia reel. The shoulder harness belt is of the "Y"configuration with the single strap being contained in aninertia reel attached to the base of the seatback. The twostraps are worn with one strap over each shoulder andfastened by metal loops into the seat belt buckle. Thespring loading at the inertia reel keeps the harness snugbut will allow normal movement required during flightoperations. The inertia reel is designed with a lockingdevice that will secure the harness in the event of sudden forward

2-18

Figure 2-10. Pilot and Copilot Seats

2-19

TM 55-1510-220-10

Section II. EMERGENCY EQUIPMENT

2-12. DESCRIPTION.The equipment covered in this section includes allemergency equipment, except that which forms part of acomplete system. For example, landing gear system, etc.Chapter 9 describes the operation of emergency exits andlocation of all emergency equipment.

2-13. FIRST AID KITS.Four first aid kits are included in the survival kit.

2-14. HAND-OPERATED FIRE EXTINGUISHER.

WARNINGRepeated or prolonged exposure tohigh concentrations ofmonobromotrifluoromethane(CF3Br) or decomposition productsshould be avoided. The liquid shallnot be allowed to come into contactwith the skin, as it may cause frostbite or low temperature burnsbecause of its very low boilingpoint.

One hand-operated fire extinguisher is mounted below thepilot's seat and a second extinguisher is located on theleft cabin sidewall, aft of the cabin door. They are of themonobromotrifluromethane (CF3Br) type. Theextinguisher is charged to a pressure of 150 to 170 PSIand emits a forceful stream. Use an extinguisher withcare within the limited area of the cabin to avoid severesplashing.

NOTEEngine fire extinguisher systems

are described in Section III.

2-15. SURVIVAL KITS.There are two different survival kits authorized forinstallation in the aircraft. Depending on the anticipatedmission, either an overwater or overland kit may beinstalled.

2-16. SURVIVAL RADIOS.Provisions are installed for installation of two AN/PRC-90radio sets: one in the pocket behind the pilots seat, andone adjacent to the cabin entrance door. Each radio isequipped with a placard giving specific operatinginstructions.

Change 2 2-20

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Section III. ENGINES AND RELATED SYSTEMS

2-17. ENGINES.The aircraft is powered by two .PT6A-41 turboprop

engines (fig. 2-11).'The engine has a three stage axial,single stage centrifugal compressor, driven by a singlestage reaction turbine. The power turbine, a two stagereaction turbine, counter-rotating with the compressorturbine, drives the output shaft. Both the compressorturbine and the power turbine are located in theapproximate center of the engine with their shaftsextending in opposite directions. Being a reverse flowengine, the ram air supply enters the lower portion of thenacelle and is drawn in through the aft protective screens.The air is then routed into the compressor. After it iscompressed, it is forced into the annular combustionchamber, and mixed with fuel that is sprayed in through14 nozzles mounted around the gas generator case. Acapacitance discharge ignition unit and two spark igniterplugs are used to start combustion. After combustion, theexhaust passes through the compressor turbine and twostages of power turbine then is routed through twoexhaust ports near the front of the engine. A pneumaticfuel control system schedules fuel flow to maintain thepower set by the gas generator power lever. Theaccessory drive at the aft end of the engine providespower to drive the fuel pumps, fuel control, the oil pumps,the refrigerant compressor (right engine), the starter-generator, and the turbine tachometer transmitter. Thereduction gearbox forward of the power turbine providesgearing for the propeller and drives the propellertachometer transmitter, the propeller overspeed governor,and the propeller governor.

2-18. ENGINE COMPARTMENT COOLING.

The forward engine compartment including theaccessory section is cooled by air entering around theexhaust stack cutouts, the gap between the propellerspinner and forward cowling, and exhausting throughducts in the upper and lower aft cowling.

2-19. AIR INDUCTION SYSTEMS - GENERAL.Each engine and oil cooler receives ram air ducted

from an air scoop located within the lower section of theforward nacelle. Special components of the engineinduction system protect the power plant from icing andforeign object damage.

2-20. FOREIGN OBJECT DAMAGE CONTROL.

The engine has an integral air inlet screen designedto obstruct objects large enough to damage thecompressor.

2-21. ENGINE ICE PROTECTION SYSTEMS.

a. Inertial Separator.

CAUTIONAfter the ice vanes have beenmanually extended, they may bemechanically actuated only. Noelectrical extension or retractionshall be attempted as damage tothe actuator may result. Linkage inthe nacelle area must be reset priorto operation of the electric system.

An inertial separation system is built into eachengine air inlet to prevent moisture particles from enteringthe engine inlet plenum under icing conditions. Amovable vane and a bypass door are lowered into theairstream when operating in visible moisture at 5°C orcolder, by energizing electrical actuators with theswitches, placarded ICE VANE RETRACT EXTEND,located on the overhead control panel. A mechanicalbackup system is provided, and is actuated by pulling theT-handles just below the pilot's subpanel placarded ICEVANE No.1 ENG No.2 ENG. Decrease airspeed to 160knots or less to reduce forces for manual extension.Normal airspeed may then be resumed.

(1) The vane deflects the ram airstream slightlydownward to introduce a sudden turn in the airstream tothe engine, causing the moisture particles to continue onundeflected, because of their greater momentum, and tobe discharged overboard.

(2) While in the icing flight mode, the extendedposition of the vane and bypass door is indicated by greenannunciator lights, No.1 VANE EXT and No.2 VANE EXT.

(3) In the non-ice protection mode, the vaneand bypass door are retracted out of the airstream byplacing the ice vane switches in the RETRACT position.The green annunciator lights will extinguish. To assureadequate oil cooling, retraction should be accomplished at15°C and

2-21

TM 55-1510-220-10

Figure 2-11. PT6A-41 Engine (Sheet 1 of 2)

above. The vanes should be either extended or retracted;there are no intermediate positions.

(4) If for any reason the vane does notattain the selected position within 15 seconds, a yellowNo.1 VANE FAIL or No.2 VANE FAIL light illuminates onthe caution/advisory panel. In this event, the manualbackup system should be used. When the vane issuccessfully positioned with the manual system, theyellow annunciator lights will extinguish. During manualsystem use, the electric motor switch position must matchthe manual handle position for a correct annunciatorreadout.

NOTEIce vane extension is confirmed bya drop in torque, and a rise in TGT.

b. Engine Air Inlet Deice System.

(1) Description. Hot engine exhaust gas isutilized for heating the air inlet lips to prevent theformation of ice. Hot exhaust gas is picked up insideeach engine exhaust stack and carried by plumbing to theinlet lip. The gas flows through the inside of the lip to thebottom where it is allowed to escape.

(2) Engine air inlet deice system switches.These switches are not used.

(3) Fuel heater. An oil-to-fuel heat exchanger,located on the engine accessory case,

Change 2 2-22

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Figure 2-11. P76A-41 Engine (Sheet 2 of 2)

operates continuously and automatically to heat the fuelsufficiently to prevent ice from collecting in the fuelcontrol unit. Each fuel control unit is protected againstice. Fuel control heat is automatically turned on for allengine operations.

2-22. ENGINE FUEL CONTROL SYSTEM.

a. Description. The basic engine fuel systemconsists of an engine driven fuel pump, a fuel control unit,a fuel flow divider, a dual fuel manifold and fourteen fuelnozzles.

b. Fuel Control Unit. One fuel control unit ismounted on the accessory case of each engine. This unitis a hydro-mechanical metering device which determines

the proper fuel schedule for the engine to produce theamount of power requested by the relative position of itspower lever. The control of developed engine power isaccomplished by adjusting the engine compressor turbine(N1) speed. N1 speed is controlled by varying theamount of fuel injected into the combustion chamberthrough the fuel nozzles. Engine shutdown isaccomplished by moving the appropriate condition leverto the full aft FUEL CUTOFF position, which shuts off thefuel supply.

Change 2 2-23

TM 55-1510-220-10

2-23. POWER LEVERS.

CAUTIONMoving the power levers intoreverse range without the enginesrunning may result in damage tothe reverse linkage mechanism.

Two power levers are located on the controlpedestal (fig. 2-8). These levers regulate power in thereverse, idle, and forward range, and operate so thatforward movement increases engine power. Powercontrol is accomplished through adjustment of the N1speed governor in the fuel control unit. Power isincreased when N1 RPM is increased. The power leversalso control propeller reverse pitch. Distinct movement(pulling up and then aft on the power lever) by the pilot isrequired for reverse thrust. Placarding beside the levertravel slots reads POWER. Upper lever travel range isdesignated INCR (increase), supplemented by an arrowpointing forward. Lower travel range is marked IDLE,LIFT and REVERSE. A placard below the lever slotsreads: CAUTION REVERSE ONLY WITH ENGINESRUNNING.

2-24. CONDITION LEVERS.

Two condition levers are located on the controlpedestal (fig. 2-8). Each lever starts and stops the fuelsupply, and controls the idle speed for its engine. Thelevers have three placarded positions: FUEL CUTOFF,LO IDLE, and HIGH IDLE. In the FUEL CUTOFFposition, the condition lever controls the cutoff function ofits engine-mounted fuel control unit. From LO IDLE toHIGH IDLE, they control the governors of the fuel controlunits to establish minimum fuel flow levels. LO IDLEposition sets the fuel flow rate to attain 52 to 55% (at sealevel) minimum N1 and HIGH IDLE position sets the rateto attain 70% minimum N1.The power lever for thecorresponding engine can select N1 from the respectiveidle setting to maximum power. An increase in low idleN1 will be experienced at high field elevation.

2-25. FRICTION LOCK KNOBS.

Four friction lock knobs (fig. 2-8) are located on thecontrol pedestal to adjust friction drag. One knob is belowthe propeller levers, one below the condition levers, andtwo under the power levers. When a knob is rotatedclockwise, friction restraint is increased opposingmovement of the affected lever as set by the pilot.Counterclockwise rotation of a knob will decrease friction

drag thus permitting free and easy lever movement. TwoFRICTION LOCK placards are located on the pedestaladjacent to the knobs.

2-26. ENGINE FIRE DETECTION SYSTEM.a. Description. A flame surveillance system is

installed on each engine to detect external engine fire andprovide alarm to the pilot. Both nacelles are monitored,each having a control amplifier and three detectors.Electrical wiring connects all sensors and controlamplifiers to DC power and to the cockpit visual alarmunits. In each nacelle, one detector monitors the forwardnacelle, a second monitors the upper accessory area, anda third the lower accessory area. Fire emits an infraredradiation that will be sensed by the detector whichmonitors the area of origin. Radiation exposure activatesthe relay circuit of a control amplifier which causes signalpower to be sent to cockpit warning systems. Anactivated surveillance system will return to the standbystate after the fire is out. The system includes afunctional test switch and has circuit protection throughthe FIRE DETR circuit breaker. Warning of internalnacelle fire is provided as follows: the red MASTERWARNING lights on the glareshield illuminateaccompanied by the illumination of a red warning light inthe appropriate fire control T-handle placarded No.1 FIREPULL or No.2 FIRE PULL (fig. 2-30). Fire detectorcircuits are protected by a single 5-ampere circuit breaker,placarded FIRE DETR, located on the overhead circuitbreaker panel (fig. 2-27).

b. Fire Detection System Test Switch. One rotary switchplacarded FIRE PROTECTION TEST on the copilot'ssubpanel is provided to test the engine fire detectionsystem. Before checkout, battery power must be on andthe FIRE DETR circuit breaker must be in. Switchposition DETR 1, checks the area forward of the air intakeof each nacelle, including circuits to the cockpit alarm andindication devices. Switch position DETR 2, checks thecircuits for the upper accessory compartment of eachnacelle. Switch position DETR 3, checks the circuits forthe lower accessory compartment of each nacelle. Eachnumbered switch position will initiate the cockpitindications previously described.

c. Erroneous Fire Detection System Indications. Duringground test of the engine fire detection system, anerroneous indication of system fault may be encounteredif an engine cowling is not closed properly, or if theaircraft

2-24

TM 55-1510-220-10

is headed toward a strong external light source. In thiscircumstance, change the aircraft heading to enable avalid system check.

2-27. ENGINE FIRE EXTINGUISHER SYSTEM.a. Description. The fire extinguisher system utilizes

an explosive squib, connected to a valve which, whenopened, allows the distribution of the pressurizedextinguishing agent through a plumbing network of spraynozzles strategically located in the fire zones of theengines.

b. Fire Pull Handles. The fire control T-handles,which are used to arm the extinguisher system arecentrally located on the pilot's instrument panel (fig. 2-30), immediately below the glareshield. These controlsreceive power from the hot battery bus. The fire detectionsystem will indicate an engine fire by illuminating themaster fault warning annunciator on the pilot's andcopilot's glareshield and the respective NO. 1 FIRE PULLor No. 2 FIRE PULL annunciators in the fire control T-handles. Pulling the fire control T-handle will electricallyarm the extinguisher system and close the fuel firewallshutoff valve for that particular engine. This will causethe red annunciator in the PUSH TO EXTINGUISH switchand the respective red No. 1 FUEL PRESS or No. 2FUEL PRESS annunciator on the warning annunciatorpanel to illuminate. Pressing the lens of the PUSH TOEXTINGUISH switch (after lifting one side of its spring-loaded clear plastic guard) will fire the squib, expelling allthe agent in the cylinder at one time. The respectiveyellow caution annunciator, No. 1 EXTGH DISCH or No.2 EXTGH DISCH on the caution/advisory annunciatorpanel will illuminate and remain illuminated, until thesquib is replaced.

c. Fire Extinguisher System Test Switch. A rotarytest switch, placarded FIRE PROTECTION TEST, islocated on the copilot's subpanel. The test functions,placarded EXTGH No.1 No.2, are arranged on the leftside of the switch and provide a test of the pyrotechniccartridge circuitry. During preflight, the pilot should rotate

the test switch through the two positions and verify theillumination of the green SQUIB OK light on the PUSHTO EXTINGUISH switch and the corresponding yellowNo.1 or No.2 EXTGH DISCH light on the caution/advisoryannunciator panel.

d. Fire Extinguishing System Supply CylinderGages. A gage, calibrated in PSI, is mounted on eachsupply cylinder for determining the level of charge andshould be checked during preflight (table 2-1).

2-28. OIL SUPPLY SYSTEM.

CAUTIONMaximum allowable oilconsumption is one quart per 10hours of engine operation.

a. The engine oil tank is integral with the airinletcasting located forward of the accessory gearbox. Oil forpropeller operation, lubrication of the reduction gearboxand engine bearings is supplied by an external line fromthe high pressure pump. Two scavenge lines return oil tothe tank from the propeller reduction gearbox. A non-congealing external oil cooler keeps the engine oiltemperature within the operating limits. The capacity ofeach engine oil tank is 2.3 U.S. gallons. The total systemcapacity for each engine, which includes the oil tank, oilcooler, lines, etc., is 3.5 U.S. gallons. The oil level isindicated by a dipstick attached to the oil filler cap. Oilgrade, specification and servicing points, are described inSection IX, Servicing.

b. The oil system of each engine is coupled to aheat exchanger unit (radiator) of fin-and-tube design.These exchanger units are the only airframe mounted partof the oil system and are attached to the nacelles belowthe engine air intake. Each heat exchanger incorporatesa thermal bypass which assists in maintaining oil at theproper temperature range for engine operation.

Change 2 2-25

TM 55-1510-220-10

Figure 2-12. Overhead Control Panel

2-29. ENGINE CHIP DETECTION SYSTEM.A magnetic chip detector is installed in the bottom

of each engine nose gearbox to warn the pilot of oilcontamination and possible engine failure. The sensor isan electrically insulated gap immersed in the oil,functioning as a normally open switch. If a large metalchip or a mass of small particles bridge the detector gap,a circuit is completed, sending a signal to illuminate a redannunciator panel indicator light placarded No.1 CHIPDETR or No.2 CHIP DETR and the MASTER WARNINGlights. Chip detector circuits are protected by two 5-ampere circuit breakers, placarded No.1 CHIP DETR andNo.2 CHIP DETR on the overhead circuit breaker panel(fig. 2-27).

2-30. ENGINE IGNITION SYSTEM.a. Description. The basic ignition system consists of

a solid state ignition exciter unit, two igniter plugs, twoshielded ignition cables, pilot controlled IGNITION ANDENGINE START switches and the ENG AUTO IGNswitch. Placing an IGNITION AND ENGINE STARTswitch to ON (forward) will cause the respective igniterplugs to spark, igniting the fuel/air mixture sprayed intothe combustion chamber by the fuel nozzles. Theignition system is activated for ground and air starts, butis switched off after combustion light up.

b. Ignition and Engine Start Switches. Two three-position toggle switches, placarded IGNITION ANDENGINE START, are located on the overhead controlpanel (fig. 2-12). These switches will initiate startermotoring and ignition in the ON position, or will motor theengine in the STARTER ONLY position. The ON switchposition completes the starter circuit for engine rotation,energizes the igniter plugs for fuel combustion, andactivates the No.1 IGN ON or No.2 IGN ON light on theannunciator panel. In the center position the switch isOFF. Two 5ampere circuit breakers on the overheadcircuit breaker panel, placarded IGNITOR CONTR No.1and No.2, protect ignition circuits. Two 5-ampere circuitbreakers on the overhead circuit breaker panel, placardedSTART CONTR No.1 and No.2, protect starter controlcircuits (fig. 2-27).

2-31. AUTOIGNITION SYSTEM.If armed, the autoignition system automatically

provides combustion re-ignition of either engine shouldaccidental flameout occur. The system is not essential tonormal engine opera

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TM 55-1510-220-10

tion, but is used to reduce the possibility of power lossdue to icing or other conditions. Each engine has aseparate autoignition control switch and a green indicatorlight placarded No.1 IGN ON or No.2 IGN ON, on theannunciator panel. Autoignition is accomplished-byenergizing the two igniter plugs in each engine.

NOTEThe system should be turned OFFduring extended ground operationto prolong the life of the igniterplugs.

a. Autoignition Switches. Two switches, located onthe overhead control panel (fig. 2-12), placarded ENGAUTO IGN-ARM control the autoignition systems. TheARM position initiates a readiness mode for theautoignition system of the corresponding engine. TheOFF position disarms the system. Each switch isprotected by a corresponding START CONTR No.1 orNo.2 5ampere circuit breaker on the overhead circuitbreaker panel (fig. 2-27).

b. Autoignition Lights. If an armed autoignitionsystem changes from a ready condition to an operatingcondition (energizing the igniter plugs in the engine) acorresponding green annunciator panel light willilluminate. The annunciator panel light is placarded No.1IGN ON or No.2 IGN ON and indicates that the ignitersare energized. The autoignition system is triggered froma ready condition to an operating condition when enginetorque drops below approximately 20%. Therefore, whenan autoignition system is armed, the igniters will beenergized continuously during the time when an engine isoperating at a level below approximately 20% torque.The autoignition lights are protected by 5-ampereIGNITOR CONTR No.1 or No.2 circuit breakers, locatedon the overhead circuit breaker panel (fig. 2-27).

2-32. ENGINE STARTER-GENERATORS.

One starter-generator is mounted on each engineaccessory drive section. Each is able to function either asa starter or as a generator. In the starter function, 24volts DC is required to power rotation. In the generatorfunction, each unit is capable of 400 amperes DC output.When the starting function is selected, the starter controlcircuit receives power, through the respective 5-ampereSTART CONTR circuit breaker on the overhead circuitbreaker panel, from either the aircraft battery or anexternal power source. When the generating function isselected, the starter-generator provides electrical power.An automatic starter cutoff control is installed in each

starter-generator to provide automatic termination of thestart cycle when engine speed reaches 44% NI. Enginespeed is sensed by a magnetic sensor located in thestarter-generator. For additional description of the starter-generator system, refer to Section IX.

2-33. ENGINE INSTRUMENTS.The engine instruments are vertically mounted near

the center of the instrument panel (fig. 230).

a. Turbine Gas Temperature Indicators. Two TGTgages on the instrument panel are calibrated in degreesCelsius (fig. 2-30). Each gage is connected tothermocouple probes located in the hot gases betweenthe turbine wheels. The gages register the temperaturepresent between the compressor turbine and powerturbine for the corresponding engine.

b. Engine Torquemeters. Two torquemeters on theinstrument panel indicate torque applied to the propellershafts of the respective engines (fig. 2-30). Each gageshows torque in percent of maximum using 2 percentgraduations and is actuated by an electrical signal from apressure sensing system located in the respectivepropeller reduction gear case. Torquemeters areprotected by individual. 0.5-ampere circuit breakersplacarded TORQUE METER No.1 or No.2 on theoverhead circuit breaker panel (fig. 2-27).

c. Turbine Tachometers. Two tachometers on theinstrument panel register compressor turbine RPM (N1)for the respective engine (fig. 2-30). These indicatorsregister turbine RPM as a percentage of maximum gasgenerator RPM. Each instrument is slaved to atachometer generator attached to the respective engine.

d. Oil Pressure/Oil Temperature Indicators. Twogages on the instrument panel register oil pressure in PSIand oil temperature in °C (fig. 2-30). Oil pressure is takenfrom the delivery side of the main oil pressure pump. Oiltemperature is transmitted by a thermal sensor unit whichsenses the temperature of the oil as it leaves the deliveryside of the oil pressure pump. Each gage is connected topressure transmitters installed on the respective engine.Both instruments are protected by 5-ampere circuitbreakers, placarded OIL PRESS and OIL TEMP No.1 orNo.2, on the overhead circuit breaker panel (fig. 2-27).

Change 2 2-27

TM 55-1510-220-10

e. Fuel Flow Indicators. Two gages on theinstrument panel (fig. 2-30) register the rate of flow forconsumed fuel as measured by sensing units coupled intothe fuel supply lines of the respective engines. The fuelflow indicators are calibrated in increments of hundreds of

pounds per hour. Both circuits are protected by 0.5-ampere circuit breakers placarded FUEL FLOW No.1 orNo.2, on the overhead circuit breaker panel (fig. 2-27).

Section IV. FUEL SYSTEMS2-34. FUEL SUPPLY SYSTEM.The engine fuel supply system (fig. 2-13) consists of twoidentical systems sharing a common fuel managementpanel (fig. 2-14) and fuel crossfeed plumbing (fig. 2-15).Each fuel system consists of five interconnected wingtanks, a nacelle tank, and an auxiliary inboard fuel tank.A fuel transfer pump is located within each auxiliary tank.Additionally, the system has an engine-driven boostpump, a standby fuel pump located within each nacelletank, a fuel heater (engine oil-to-fuel heat exchangerunit), a tank vent system, a tank vent heating system andinterconnecting wiring and plumbing. Refer to Section IXfor fuel grades and specifications. Fuel tank capacity isshown in table 2-2. Gravity feed fuel flow is shown infigure 2-16.

a. Engine Driven Boost Pumps.CAUTION

Engine operation using only theengine-driven primary (highpressure) fuel pump withoutstandby pump or engine-drivenboost pump fuel pressure is limitedto 10 cumulative hours. Thiscondition is indicated byillumination of either the No.1 orNo.2 FUEL PRESS warningannunciator lights and thesimultaneous illumination of bothMASTER WARNING lights. Refer toChapter 9. All time in this categoryshall be entered on DA Form 2408-13 for the attention of maintenancepersonnel.

A gear-driven boost pump, mounted on eachengine supplies fuel under pressure to the inlet of theengine-driven primary high-pressure pump for enginestarting and all normal operations. Either the engine-driven boost pump or standby pump is capable ofsupplying sufficient pressure to the engine-driven primaryhigh-pressure pump and thus maintain normal engineoperation.

b. Standby Fuel Pumps. A submerged,electrically-operated standby fuel pump, located withineach nacelle tank, serves as a backup unit for theengine-driven boost pump. The standby pumps are

switched off during normal system operations. A standbyfuel pump will be operated during crossfeed operation topump fuel from one system to the opposite engine. Thecorrect pump is automatically selected when theCROSSFEED switch is activated. Each standby fuelpump has an inertia switch included in the power supplycircuit. When subjected to a 5 to 6 G shock loading, as ina crash situation, the inertia switch will remove electricalpower from the standby fuel pumps. The standby fuelpumps are protected by two 10-ampere circuit breakersplacarded STANDBY PUMP No.1 or No.2, located theoverhead circuit breaker panel (fig. 2-27), and four 5-ampere circuit breakers (2 each in parallel) on the hotbattery bus.

c. Fuel Transfer Pumps. The auxiliary tank fueltransfer system automatically transfers the fuel from theauxiliary tank to the nacelle tank without pilot action.Motive flow to a jet pump mounted in the auxiliary tanksump is obtained from the engine fuel plumbing system,downstream from the engine driven boost pump androuted through the transfer control motive flow valve.The motive flow valve is energized to the open positionby the control system to transfer auxiliary fuel to thenacelle tank to be consumed by the engine during theinitial portion of the flight. When an engine is started,pressure at the engine driven boost pump closes apressure switch, which after a 30 second time delay toavoid depletion of fuel pressure during starting, energizesthe motive flow valve. When auxiliary fuel is depleted, alow level float switch de-energizes the motive flow valveafter a 10 second time delay provided to prevent cyclingof the motive flow valve due to sloshing fuel. In the eventof a failure of the motive flow valve or the associatedcontrol circuitry, the loss of motive flow pressure whenthere is still fuel remaining in the auxiliary fuel tank issensed by a pressure switch and float switch,respectively, which illuminates a caution annunciatorplacarded No. 1 NO FUEL XFR or No. 2 NO FUEL XFR.During engine start, the pilot should note that the NOFUEL XFR annunciators extinguish 30 to 50 seconds afterengine start. The NO FUEL XFR annunciators will notilluminate if auxiliary tanks are empty. A manual overrideis incorporated as a backup for the automatic transfersystem. This is initiated by placing the AUX TRANSFERswitch, located on the fuel

Change 2 2-28

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Table 2-2. Fuel Quantity Data

TANKS NUMBER GALLONS **POUNDS

LEFT ENGINEWing TanksNacelle TankAuxiliary Tank

511

1355779

877.5370.5513.5

RIGHT ENGINEWing TanksNacelle TankAuxiliary Tank

511

1355779

877.5370.5513.5

*TOTALS 14 542 3523.0* Unusable fuel quantity and weight (4 gallons, 26 pounds not included in totals).

** Fuel weight is based on standard day conditions at 6.5 pounds per U.S. gallon.BTO 1490

management panel to the OVERRIDE position. This willenergize the transfer control motive flow valve. Thetransfer systems are protected by. 5ampere circuitbreakers placarded AUXILIARY TRANSFER No. 1 or No.2, located on the overhead circuit breaker panel (fig. 2-27).

NOTEIn turbulence or during maneuversthe NO FUEL XFR annunciatorsmay momentarily illuminate afterthe auxiliary fuel has completedtransfer.

d. Fuel Gaging System. The total fuel quantityin the left or right main system, or left or right auxiliarysystem is measured by a capacitance type fuel gagingsystem. Two fuel gages, one for the left and one for theright fuel system, indicate the fuel quantity in pounds.Refer to Section IX for fuel capacities and weights. Amaximum of 3% error may be encountered in eachsystem. However the system is compensated for fueldensity changes due to temperature excursions. Inaddition to the fuel gages, yellow No. 1 NAC LOW or No.2 NAC LOW annunciators on the caution/advisoryannunciator panel illuminate when there is approximately20 minutes of fuel per engine remaining (on standard day,lat sea level, maximum cruise power consumption rate).The fuel gaging system is protected by individual 5-ampere circuit breakers placarded QTY IND and QTYWARN No. I or No. 2 located on the overhead circuitbreaker panel (fig. 2-27). A mechanical spiral float gage

is installed in each auxiliary fuel tank to provide anindication of fuel level when servicing the tank. The gageis installed on the auxiliary fuel tank cover, adjacent to thefiller neck (fig. 2-17). A small sight window in the upperwing skin permits observation of the gage.

e. Fuel Management Panel. The fuel managementpanel is located on the cockpit overhead between the pilotand copilot. It contains the fuel gages, standby fuel pumpswitches, the crossfeed valve switch and a fuel gagingsystem control switch and transfer control switches arealso installed.

(1) Fuel gaging system control switch. Aswitch on the fuel management panel (fig. 2-14)placarded FUEL QUANTITY, MAIN AUXILIARY, controlsthe fuel gaging system. When in the MAIN position thefuel gages read the total fuel quantity in the left and rightwing fuel system. When in the AUXILIARY position thefuel gages read the fuel quantity in the left and rightauxiliary tanks only.

(2) Standby fuel pump switches. Twoswitches, placarded STANDBY PUMP ON located on thefuel management panel (fig. 2-14) control a submergedfuel pump located in the corresponding nacelle tank.During normal aircraft operation both switches are off solong as the engine-driven boost pumps function andduring crossfeed operation. The loss of fuel

Change 2 2-29

TM 55-1510-220-10

Figure 2-13. Fuel System Schematic

pressure, due to failure of an engine driven boost pumpwill illuminate the MASTER WARNING lights on theglareshield and will illuminate the No. 1 FUEL PRESS orNo.2 FUEL PRESS on the warning annunciator panel.Turning ON the STANDBY PUMP will extinguish theFUEL PRESS lights. The MASTER WARNING lightsmust be manually cleared.

NOTEBoth standby pump switches shall

(3) Fuel transfer control switches. Twoswitches on the fuel management panel (fig. 214),placarded AUX TRANSFER OVERRIDE AUTO controloperation of the fuel transfer pumps. During normaloperation both switches are in AUTO which allows thesystem to be automatically actuated by fuel flow to theengine. If either transfer system fails to operate, the fault

condition is indicated by two illuminated MASTERCAUTION lights on the glareshield and a steadilyilluminated yellow No.1 NO FUEL XFR or No.2 NO FUELXFR light on the caution annunciator panel.

(4) Fuel crossfeed switch. The fuel crossfeedvalve is controlled by a 3-position switch (fig. 2-14),located on the fuel management panel, placardedCROSSFEED OFF. Under normal flight conditions theswitch is left in the OFF position. During emergencysingle engine operation, it may become necessary tosupply fuel to the operative engine from the fuel systemon the opposite side. The crossfeed system is placardedfor fuel selection with a simplified diagram on theoverhead fuel control panel. Place the standby fuel pumpswitches in the off position when

2-30

TM 55-1510-220-10

Figure 2-14. Fuel Management Panel

crossfeeding. A lever lock switch, placardedCROSSFEED, is moved from the center OFF position tothe left or to the right, depending on direction of fuel flow.This opens the crossfeed valve and energizes the standbypump on the side from which crossfeed is desired. Duringcrossfeed operation with firewall fuel valve closed,auxiliary tank fuel will not crossfeed. When the crossfeedmode is energized, a green FUEL CROSS-FEED light onthe caution/advisory panel will illuminate. Crossfeedsystem operation is described in Chapter 9. Thecrossfeed valve is protected by a 5-ampere circuit breakerplacarded CROSSFEED located on the overhead circuitbreaker panel (fig. 2-27).

f. Firewall Shutoff Valves.

CAUTIONDo not use the fuel firewall shutoffvalve to shut down an engine,except in an emergency. Theengine-driven highpressure fuelpump obtains essential lubricationfrom fuel flow. When an engine isoperating, this pump may beseverely damaged (while cavitating)if the firewall valve is closed beforethe condition lever is moved to theFUEL CUTOFF position.

2-31

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Figure 2-15. Crossfeed Fuel Flow

2-32

TM 55-1510-220-10

Figure 2-16. Gravity Feed Fuel Flow

2-33

Figure 2-17. Auxiliary Fuel Tank MechanicalFuel Gage

The fuel system incorporates a fuel line shutoffvalve mounted on each engine firewall. The firewallshutoff valves close automatically when the fireextinguisher T-handles on the instrument panel are pulledout. The firewall shutoff valves receive electrical powerfrom the main buses and also from the hot battery buswhich is connected directly to the battery. The valves areprotected by circuit breakers placarded FIREWALLVALVE No.1 or No.2 on the overhead circuit breakerpanel (fig. 2-27), and FIREWALL SHUTOFF No.1 orNo.2 on the hot battery bus circuit breaker board.

g. Fuel Tank Sump Drains. A sump drain wrench isprovided in the aircraft loose tools to simplify draining asmall amount of fuel from the sump drain.

(1) There are five sump drains and one filterdrain in each wing (table 2-3).

(2) An additional drain for the extended rangefuel system line extends through the bottom of the wingcenter section adjacent to the fuselage. Anytime theextended range system is in use, a part of the preflightinspection would consist of draining a small amount offuel from this drain to check for fuel contamination.Whenever the extended range system is removed fromthe aircraft and the fuel line is capped off in the fuselage,the remaining fuel in the line shall be drained.

h. Fuel Drain Collector System. Each engine isprovided with a fuel drain collector system to return fueldumped from the engine during clearing and shutdownoperations back into its respective nacelle tank. Thesystem draws power from the No.4 feeder bus. Fueltransfer is completely automatic. Fuel from the engineflow divider drains into a collector tank mounted below theaft engine accessory section. An internal float switchactuates an electric scavenger pump which delivers thefuel to the fuel purge line just aft of the fuel purge shutoffvalve. A check valve in the line prevents the backflow offuel during engine purging. The circuit breaker for bothpumps is located in the fuel section of the overheadcircuit breaker panel; placarded SCAVENGER PUMP. Avent line, plumbed from the top of the collector tank, isrouted through an inline flame arrestor and thendownward to a drain manifold on the underside of thenacelle.

i. Fuel Vent System. Each fuel system is ventedthrough two ram vents located on the underside of thewing adjacent to the nacelle. To prevent icing of the ventsystem, one vent is recessed into the wing and thebackup vent

Table 2-3. Fuel Sump Drain Locations

NUMBER DRAIN LOCATION1 Leading Edge Tank Outboard of nacelle, underside of wing

1 Integral Tank Underside of wing, forward of aileron

1 Firewall fuel filter Underside of cowling forward of firewall

1 Sump Strainer Bottom center of nacelle forward of wheel well

1 Gravity feed line Aft of wheel well

1 Auxiliary Tank At wing root, just forward of the flap

2-34

TM 55-1510-220-10

protrudes out from the wing and contains a heatingelement. The vent line at the nacelle contains an inlineflame arrestor.

j. Engine Oil-to-Fuel Heat Exchanger. An Kengine oil-to-fuel heat exchanger, located on each engineaccessory case, operates continuously and automaticallyto heat the fuel delivered to the engine sufficiently toprevent the freezing of any water which it might contain.The temperature of the delivered fuel is thermostaticallyregulated to remain between 21°C and 32°C.

2-35. FUEL SYSTEM MANAGEMENT.a. Fuel Transfer System. When the auxiliary tanks

are filled, they will be used first. During transfer ofauxiliary fuel, which is automatically controlled, thenacelle tanks are maintained full. A check valve in thegravity feed line from the outboard wing prevents reversefuel flow. Normal gravity transfer of the main wing fuelinto the nacelle tanks will begin when auxiliary fuel isexhausted. The system will gravity feed fuel only to itsrespective nacelle tank, i.e. left or right (fig. 2-16). Fuelwill not gravity feed through the crossfeed system.

b. Operation With Failed Engine-Driven Boost Pumpor Standby Pump. Two pumps in each fuel systemprovide inlet head pressure to the engine-driven primaryhigh-pressure fuel pump. If crossfeed is used, a thirdpump, the standby fuel pump from the opposite system,will supply the required pressure. Operation under thiscondition will result in an unbalanced fuel load as fuelfrom one system will be supplied to both engines while allfuel from the system with the failed engine driven andstandby boost pumps will remain unused. A triple failure,which is highly unlikely, would result in the engine drivenprimary pump operating without inlet head pressure.Should this situation occur, the affected engine cancontinue to operate from its own fuel supply on its engine-driven primary high-pressure fuel pump.

2-36. FERRY FUEL SYSTEM.Provisions are installed for connection to long range

fuel cells.

Section V. FLIGHT CONTROLS

2-37. DESCRIPTION.The aircraft's primary flight control systems consist

of conventional rudder, elevator and aileron controlsurfaces. These surfaces are manually operated from thecockpit through mechanical linkage using a control wheelfor the ailerons and elevators, and adjustable rudder/brake pedals for the rudder. Both the pilot and copilothave flight controls. Trim control for the rudder, elevatorand ailerons is accomplished through a manually actuatedcable-drum system for each set of control surfaces. Theautopilot has provisions for controlling the position of theailerons, elevators, and rudder. Chapter 3 describes theoperation of the autopilot system.

2-38. CONTROL WHEELSElevator and aileron control surfaces are operated

by manually actuating either the pilot's or copilot's controlwheel. Switches are installed in the outboard grip of eachwheel to operate the elevator trim tabs. A microphoneswitch, a chaff dispense switch, and an autopilot/yawdamp/ electric trim disconnect switch are also installed inthe outboard grip of each wheel. A transponder identswitch is installed on top of the inboard grip of each

control wheel. These control wheels (fig. 2-18) areinstalled on each side of the instrument subpanel. Amanually wound 8-day clock is installed in the center ofthe pilot's control wheel, and a digital electric clock isinstalled in the center of the copilot's control wheel. Amap light switch, and a pitch synchronization and controlwheel steering switch are mounted adjacent to the clockin each control wheel.

2-39. RUDDER SYSTEM.a. Rudder Pedals. Aircraft rudder control and nose

wheel steering is accomplished by actuation of the rudderpedals from either pilot's or copilot's station (fig. 2-7).The rudder pedals may be individually adjusted in either aforward or aft position to provide adequate leg room forthe pilot and copilot. Adjustment is accomplished bydepressing the lever alongside the rudder pedal arm andmoving the pedal forward or aft until the locking pinengages in the selected position.

2-35

Figure 2-18. Control Wheels

2-36

TM 55-1510-220-10

b. Yaw Damp System. A yaw damp system isprovided to aid the pilot in maintaining 'directionalstability and increase ride comfort. The system may beused at any altitude and is required for flight above17,000 feet. It must be deactivated for takeoff andlanding. The yaw damp system is a part of the autopilot.Operating instructions for this system are contained inChapter 3. The system is controlled by a YAW DAMPswitch adjacent to the ELEV TRIM switch on the pedestalextension.

c. Rudder Boost System. A rudder boost system isprovided to aid the pilot in maintaining directional controlresulting from an engine failure or a large variation ofpower between the engines. Incorporated in the ruddercable system are two pneumatic rudder boosting servoswhich actuate the cables to provide rudder pressure tohelp compensate for asymmetrical thrust.

(1) During operation, a differential pressurevalve accepts bleed air pressure from each engine.When the pressure varies between the bleed air systems,the shuttle in the differential pressure valve moves towardthe low pressure side. As the pressure difference reachesa preset tolerance, a switch closes on the low pressureside which activates the rudder boost system. Thissystem is designed only to help compensate forasymmetrical thrust. Appropriate trimming is to beaccomplished by the pilot.. Moving either or both of thebleed air valve switches on the overhead control panel toPNEU & ENVIRO OFF position will disengage the rudderboost system.

NOTECondition levers must be in LOWIDLE position to perform rudderboost check.

(2) The system is controlled by a switchlocated on the extended pedestal placarded RUDDERBOOST ON OFF (fig. 2-8), and is to be turned on beforeflight. A preflight check of the system can be performedduring the run-up by retarding the power on one engine toidle and advancing power on the opposite engine until thepower difference between the engines is great enough toactivate the switch to turn on the rudder boost system.Movement of the appropriate rudder pedal (left engineidling, right rudder pedal moves forward) will be notedwhen the switch closes, indicating the system isfunctioning properly for low engine power on that side.Repeat the check with opposite power settings to check

for movement of the opposite rudder pedal. The systemis protected by a 5-

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