TO IA-IOA-I FLIGHT MANUAL USAF SERIES A-lOA AIRCRAFT

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These instructions provide you with a general knowledge of the aircraft and it s characteristics and specific normal and emergency operating proc e dures. Your experience is recognized; therefore, basic flight principles are avoided. Instructions in this manual are for a crew inex perienced in the operation of this aircraft. This manual provides the best possible operating instruc - tions under most circumstances. Multiple emergen - cies, adverse weather, terrain. etc., may require mod- ification of the procedures. PERMISSIBLE OPERATIONS The flight manual takes a "positive approach" and normally states onl y what you can do . Unusual oper at ions or configurations are prohibited unless specifically covered herein. Clearance must be ob- tained before any questionable operation, which is not specifically permitted in this manual, is attempted. ii T.O ' ... ·I OA · I HOW TO BE ASSURED OF HAVING LATEST DATA Refer to T . O. 0- 1- 1-5 for in forma tion concerning cur- rent fli ght manuals. safely s upplem ents, operational supp leme nts. a nd checkli sts. Also, chec k the flight ma nu a l cover page, the title bl oc k of each safe ty and operational sLlp plement , and all sta tus pages con- tained in the flight manLlal or attached to formal safety and op e rational s uppl ements . Clear up all di sc repan c ie s before flight. ARRANGEMENT The manual is divided into seve n independent sec- tio ns to simplify reading it straig ht through o r lIsing it as a refere nce manual. SAFETY SUPPLEMENTS Inform at ion involving safety will be promptly for- warded to you in a safety s upple ment. Supplements coveri ng loss of life will get to y ou within 48 hour s by teletype. and sllf1p lemcnt s covering serious damage to equipment within 10 days by mail. The cover page of the flight manual and the title block of each safety supplement should be checked to determine the effect they may have on existing supplements . OPERATIONAL SUPPLEMENTS In fo rmati o n invohin g c hange s to operating pro - cedures will be forwarded to yo u by operational ~ up plements. The procedure for handling operational supplements is the same as for sa fet y supplemcnts. MAJCOM T.O.1A·10A·1 FLIGHT MANUAL REVIEW CONFERENCE (FMRC) The Interim/Formal Safety/Operational supple- ments listed on the Title Page were incorporated as written or the intent was incorporated as deter- mined by the MAJCOMs during the FMRC. CHECKLISTS The flight manual contains itemized procedure s with necessary amplifications. The checklist contains itemized procedures without the amplification . Primary line item s in the flight manual and checklist arc identical. If a formal safety or operational s up- plement affects yo ur checklist, the affected checklist page will be attached to the s upplement. HOW TO GET PERSONAL COPIES Each flight crew member is entitled to personal copies of the flight manual, safety supplements, operational s upplements. and checklists. The required quantitie s s hould be ordered before you need them to insure their prompt receipt. Check with your publica- tion distribution officer - it is his job to fulfill your T . O. reque sts. Basically, you mu st order the required quantities on the appropriate Numerical Index and Requirement Table (NIRT). T .O. 00-5-1 and 00-5-2 give detailed information for properly ordering the se publication s. Make sure a system is established at yo ur base to deliver these publications to the flight crew immediately upon receipt . FLIGHT MANUAL BINDERS Looseleaf binders and sec tionalized tabs are available for usc with your manual. They are obtained through TO 1A-10A-1 local purchase procedures and are listed in the Federal Supply Schedule (FSC Group 75, Office Sup- plies, Part I). Check with your supply per so nnel for assistance in procuring these item s. DEFINITION OF WORDS "SHALL," "WILL," "SHOULD" AND "MA Y" The words "shall" and "will" indicate a mandator y requirement. The word "should" indicates a non- mandatory desire or preferred method of accomplish- ment. The word "may" indicates an acceptable or suggested mean s of accomplishment. WARNINGS, CAUTIONS AND NOTES The following definitions apply to "Warnings", "Cautions" and "Notes" found throughout the manual. WARNING' Operating procedures. technique s, etc ., which will result in personal injur y or loss of life if not carefully followed. Operating procedures. techniques, etc .. which will result in damage to equipment if not carefully followed. NOTE An operating procedure. technique. etc ., which is considered esse ntial to emphasize . YOUR RESPONSIBILITY - TO LET US KNOW Every effort is made to keep the flight manual cur- rent. Review conferences with operating personnel and a constant review of accident and flight test reports assure inclusion of the latest data in the manual. \Ve cannot correct an error unle ss we know of it s existence. In this regard. it is essential that you do your part. Comments. corrections. and questions regarding this manual or any phase of the flight manual program are welcomed. These should be for- warded through your command channels on AF Form 847 to: SM-ALC / MMSRB. \1cClelian AFB. California 95652 . Change 8 iii TO lA-lOA-l TCTO IDENTIFICATION The following TCTOs affecting A-lO Airplanes are covered in this manual. This is not a complete listing and only includes TCTOs listed by number directly affecting this manual. Refer to the Numerical Index and Requirement Table (TO 0-1-1-5) for complete listing of TCTOs for these airplanes. TO NUMBER lA-10-649 1A-10-670 1A-10-764 lA-1O-831 1A -lO-869 1A-10-883 1A-lO-90l 1A-lO-932 1A-1O-975 1A-10-986 lA-10-l059 1A-10-l084 1A-10-l087 1A-lO-1154 1A-10-1158 1A-lO-1170 lA-10-1176 lA-lO-l177 lA-lO-1186 lA-lO-1195 lA-lO-1216 lA-10-1227 lA-10-1228 lA-10-1243 1A-IO-1250 lA-10-1267 iv Change 8 DISPOSITION SECTION I, II I I, II, III, V I I I I I, III, VI I, II, III I I I I, III, V, VII I, II, V I I, II I, II, V I I I I I I, II SUBJECT Install AAU-34/ A Altimeter Revise Cabin Air Distribution System Chaff/Flare Dispensers Modification of Emergency Canopy Actuator Un- lock Assembly APU Ground Shutoff Addition of Time Delay in Jam Indication Sys- tem Modification of Wing Flap Control System Incorporation of HARS/SAS Attitude Validity Assembly into A-10 Aircraft Incorporation of An Inertial Navigation System in A-lO Aircraft Continuous Ignition During Stall Warning and Gunfiring Installation of GFU-l6A Gun Gas Diverter Sys- tem, A-lOA Aircraft and GA-10A Installation of Canopy Jettison Handle Decal, A-lO Aircraft New Air Force Serial Numbers for Option IX A-10 Aircraft Installation of TF34-GE-100A Engines and Re- placement of Aircraft Cockpit Engine Temper- ature Indicators A-lOA Aircraft Installation and Operational Checkout of the Turbine Engine Monitoring System (TEMS) Line Replaceable Units (LRU), A-lOA Aircraft Reconfiguration of Cockpit Right Console Installation of Airborne Video Tape Recorder (A VTR) and Cockpit Television Sensor (CTVS) Retrofit Incorporation of Self Defense Air-To-Air AIM-9 Missile Capability, A-10 Aircraft Modification of Air Refueling Line Installation of Electroluminescent Formation Lights Modification of Canopy Jettison Handle Alternate Closure of 65PSI Valve, A-10 Aircraft Install a "SEAT NOT ARMED" indication sys- tem on ACES II Ejection Seats, A-lOA and OA-10 Aircraft Installation of Air Force Seawater Activated Re- lease System (SEAWARS), A-10 ACES II Parach utes P /N J 114-509-517 A-10 Stall Warning for Approach and Landing Install HAVE QUICK II on A-10 Aircraft EFFECTIVITY PAGE OJ Aircraft sernos 76-0519 through 76-0520. [JJ Aircraft not modified by T.O. IA-IO-986. T.O. 1A-10A-1 8J Aircraft serno 81-0939 (81-0001 not modified by T.O. IA-IO-1087) and subsequent and those modified by T.O. IA-IO-986. [I] Aircraft modified by T.O . IA-IO-1084. [I] Aircraft prior to serno 75-00280 not modified by T.O. IA-IO-764. ITQ] Aircraft serno 75-00280 and subsequent. [ill Aircraft sernos 75-00280 through 77-0226 not modified by T.O. IA-IO-764. §] Aircraft prior to serno 76-0512. ~ Aircraft serno 76-0512 and subsequent. ~ Aircraft serno 76-0519 and subsequent and those modified by T .O. IA-IO-649 . IE] Aircraft prior to serno 76-0521 not modified by T .O. IA-IO-649. ~ Aircraft serno 76-0521 and subsequent and those modified by T.O. IA-IO-649. ~ Aircraft serno 76-0535 and subsequent and those modified by T.O . IA-IO-869. ~ Aircraft prior to serno 79-0167 not modified by T.O. IA-IO-932. ~ Aircraft prior to serno 77-0177 not modified by T.O. IA-IO-83\. [IT] Aircraft prior to serno 77-0259. !TIl Aircraft serno 77-0177 and subsequent and those modified by T.O. IA-IO-83\. ~ Aircraft prior to serno 75-00280. llil Aircraft serno 77-0227 and subsequent and those modified by T.O. IA-IO-764. ~ Aircraft sernos 77-0259 through 79-0166. ~ Aircraft serno 77-0259 and subsequent. ~ Aircraft prior to serno 78-0582. !ill Aircraft sernos 78-0582 through 79-0166 . ~ Aircraft serno 78-0582 and subsequent. ~ Aircraft prior to serno 78-0622. Change 7 v I TO lA-lOA-l 1551 1561 1571 ~ [I[J [gJ [§J [EJ 167 1 []§J ~ i75 1 rnIJ [g] ~ []!J [![] em 196 1 1971 1981 ~~ 1102 ! 11031 ~--' 1105 1 1106 1 EFFECTIVITY PAGE (CONT) Aircraft serno 78-0622 and subsequent. Aircraft prior to serno 78-0626. Aircraft serno 78-0626 and subsequent. Aircraft serno 78-0634 and subsequent and those modified by TO 1A-10-869. Aircraft prior to serno 79-0167 not modified by TO 1A-10-975 or TO 1A-10-975D. Aircraft serno 79-0167 and subsequent and those modified by TO 1A-10-975 or TO 1A-10-975D. Aircraft serno 79-0167 and subsequent and those modified by TO 1A-1O-883. Aircraft serno 79-0167 and subsequent and those modified by TO 1A-10-932. Aircraft serno 80-0178 and subsequent and those modified by TO 1A-10-976. Aircraft prior to serno 80-0180. Aircraft serno 80-0180 and subsequent. Aircraft serno 79-0173 and subsequent. Aircraft serno 80-0255 and subsequent and those modified by TO 1A-10-901. Aircraft modified by TO 1A-10-1059. Aircraft prior to serno 78-0582 not modified by TO 1A-1O-1108. Aircraft serno 78-0582 and subsequent and those modified by TO 1A-1O-1108. Aircraft modified by TO 1A-10-1154. Aircraft modified by TO 1A-10-1158. Aircraft modified by TO 1A-10-1170. Aircraft modified by TO 1A-10-1186. Aircraft modified by TO 1A-1O-1176. Aircraft not modified by TO 1A-10-1176. Aircraft modified by TO 1A-10-1250. Aircraft modified by TO 1A-1O-1177. Aircraft modified by TO 1A-10-l243. Aircraft modified by TO 1A-10-1195. vi Change 8 EFFECTIVITY PAGE (CONT) 1107 1 Aircraft modified by TO lA-lO-1228_ 11081 Aircraft modified by TO lA-10-1216_ 110 91 Aircraft modified by TO 1A-IO-1267. C!.!QJ Aircraft modified by AOA indic ator Recalibration. ~, f ." TO lA-lOA-l Change 8 viA/(viB Blank) T . O.1A - 10A- 1 A 10A AIR SUP PORT ATTA CK AIRCRAFT ~ I. ,l: I I l - IOA-I-)O Figure 1-\ vii/(viii blank) T.O. 1A-10A-1 SECTION I DESCRIPTION AND OPERATION TABLE OF CONTENTS The Aircraft . . . ..... . . .............. 1- 1 Engines . . . . . . . . . . . . . . . . . . . . . . . . .. 1-3 Fire Extinguishing System . . . . . . . . . . . . . 1-11 Auxiliary Power Unit. . . . . . . . . . . . . . . .. 1-12 Aircraft Fuel System . . . . . . . . . . . . . . . .. 1 -1 3 Electrical Power System . . . . . . . . . . . . .. 1-20 Hydraulic Power Supply System . . . . . . . . 1-24 Landing Gear System . . . . . . . . . . . . . . .. 1-25 Nosewheel Steering System . . . . . . . . . . . 1-28 Wheel Brake System. . . . . . . . . . . . . . . . . 1-28 Primary Flight Control System . . . . . . . . .. 1-29 Manual Reversion Flight Control System .... . . .. .. . .. . . . .. .. . .. . . 1- 40 Secondary Flight Control System . . .... . 1-43 Boarding Ladder . . . . . . . . . . . . . . . . . . .. 1- 46 Canopy . . ... . .... .. . .. . . . . ... ... . 1-48 Ejection Seat . . . . . . . . . . . . . . . . . . . . .. 1-52 THE AIRCRAFT The A-1O is a single-seat close air support aircraft (figures 1-1 and 1-2) manufactured by Fairchild Republic Company, Farmingdale, New York. The aircraft is a low wing, low tail configuration with two high bypass turbo fan engines installed in nacelles mounted on pylons extending from the aft fuselage. Twin vertical stabilizers are mounted on the outboard tips of the horizontal tail. The tricycle forward retracting landing gear is equipped with an anti-skid system and a steerable nosewheel. The nose gear is installed to the right of the aircraft centerline to permit near centerline gunfire. The nose gear retracts fully into the fuselage while the main gears partially retract into streamlined pods in the wings. A titanium armor installation surrounds the cockpit. The primary flight controls are equipped with artifi- cial feel devices to simulate aerodynamic feel. The Oxygen System . . . . . . . . . . . . . . . . . . .. 1-56 Environment System . . . . . . . . . . . . . . . . . 1-58 Flight Instruments . . . . . . . . . . . . . . . . . .. 1-67 Standby Flight Instruments . . . . . . . . . . . . 1-77 Communication/Navigation . . . . . . . . . . .. 1-77 Lighting System . . . . . . . . . . . . . . . . . . .. 1-141 Armament System . . . . . . . . . . . . . . . . . . 1-1 50 Head-Up Display System . . . . . . . . . . . . . . 1-152 Gunsight Camera System . . . . . . . . . . . .. 1-168 Cockpit Television Sensor/Airborne Video Tape Recorder System . . ....... 1- 168 TV Monitor System . . . ... ... . .. .... . 1- 168 Target Identification Set Laser . .... . . .. 1-168 Chaff/Flare Dispensing System . . . . . . . .. 1-168 Electronic Countermeasure System ....... . ......... .. . .. ... 1-1 68 Servicing Diagram. . . . . . . . . . . . . . . . . .. 1-1 71 elevator and aileron controls split into redundant separate systems before leaving the armor protection. The controls are powered by two independent hydraulic systems, either of which has the capability of controlling the airplane. If both hydraulic systems fail, the airplane can be flown using a manual reversion system. The ailerons consist of an upper and lower panel that become speed brakes when opened. The windshield front panel is resistant to small arms fire and birds . The windshield side panels are resistant to spall spray caused by penetrations. The fuselage fuel cell sumps are self-sealing on the lower portion and tear resistant on the upper portion. The cells are filled with a flexible foam to prevent fuel tank explosion . Single-point ground refueling and engine feed lines are self-sealing. The escape system provides a zero/zero capability (zero velocity, and zero pitch and roll attitude) either with the canopy removed or through the canopy. The Change 1 1-1 ~ , N GENERAL ARRANGEMENT DIAGRAM AILERON I SPEED BRAKE RIG HT WI NG TA NK GAU-SA GUN LEFT MA IN TA NK (AFT) SINGlE - POINT REFUELING FIRE EX TI NG UISHE R BOTTLES~ LEFT WI NG TANK armament system includes a high fire rate 30mm seven-barrel gun with ammunition stored in a drum. A variety of stores are carried on II pylons, 4 on each wing and 3 on the fuselage . AIRCRAFT DIMENSIONS The overall dimensions of the aircraft under normal conditions of gross weight, tire and strut inflation are as follows: Overall length 53 ft 4 in. Wing span 57 ft 6 in. Horizontal tail span 18 ft 10 in. Height to top of fin 14 ft 8 in. Wheel base 17 ft 9 in. Wheel tread 17ft3in. Refer to Section II for minimum turning radius and ground clearance dimensions. AIRCRAFT GROSS WEIGHT The aircraft operating weight is approximately 28,000 I pounds. This weight includes pilot, gun (full of am- munition), 11 empty pylons, oil, windshield wash, and unusable fuel. Refer to Section V for gross weight limits. For specific aircraft weight, refer to Weight and Balance Data. T.O. 1-1 B-40. ENGINES The aircr.aft is powered by two General Electric TF34-GE - 100i [2] -100A engines (figure 1-3). Sea level, standard day, static thrust for an installed engine is approximately 8,900 pounds at maximum thrust. The engine incorporates a single-stage bypass fan and a 14-stage axial flow compressor. Bypass air produces over 85070 of engine thrust. Therefore, engine fan speed is the best indication of thrust. Variable inlet guide vanes automatically modulate throughout the engine operating range. An accessory gearbox drives a hydraulic pump, fuel pump and fuel control, oil pump, and an electric generator. An air bleed for aircraft systems is provided. Engine ac- celeration time from IDLE to MAX thrust will be approximately 10 seco nds at sea level. Engine thrust droop results from differential expansion of the engine turbines and casings during transients from low to high thrust operation. The duration and extent of the thrust droop is dependent upon the rate / range of throttle movement. Thrust droop is decreased if the engines have been idling for a period of time. Thrust droop is further decreased if the engines have been run up before takeoff. An example of the worst T.O. 1A-10A-1 condition would be a scramble takeoff where takeoff is accomplished shortly after engine start. Maximum droop occurs approximately 10 seconds after the throttle is advanced from IDLE to MAX. After ap- proximately 4 minutes of operation at MAX thrust, power output returns to normal. Elimination of thrust droop can be observed on the engine fan speed indicator. Thrust will increase as fan speed increases . ~ TURBINE ENGINE MONITORING SYSTEM ITEMS) The turbine engine monitoring system (TEMS) pro- vides a means for supporting the on-condition maintenance concept for the TF34-IOOA engine in- stalled on t he A-lOA aircraft. I nformation is pro- vided to the system electronic processor unit (EPU) automatically whenever the engine is operated . I f any engine operation limit is exceeded, or when the TEMS DATA switch is pressed, a frame of data is recorded. An overlimits event will be displayed on the umbilical display unit (UDU), located in the nosewheel well. This data is provided in code form for the technician or pilot to determine engine condi- tion . The codes are provided in figure 2-9. ENGINE OIL SYSTEM The engine oil system is self-contained and all the oil supp lied is used for lubrication and cooling . Usable oil capacity is 5.6 quarts, and maximum oi l consump- tion is 0.5 pints / hr . An oil pressure indicator and an independent light on the caution li ght panel monitor oil pressure of each engine. Oil grade and spe(ifica - tion tl) be us ed arc co\ered ill the senicing diagram, figure 1-71. Engine Oil Pressure Indicators An oil pressure indicator (~I, figure FO - I) i~ pro- \ided for each engine. They indicate oi l pressure in psi, and are powered by the 26 V in\lrUlllent t ransforll1er bus . Engine Oil Pressure Caution Lights An engine oil pressure caution light (independenl of the oil pressure indicators) (figure 1-65) , on the cau- tion light panel, is provided for each engine . The lights are placarded L-ENG OIL. PRESS and R- ENG OIL. PRESS and will come Oil if the pressure is less than n.5( ± 2.5) psi. ft.J~cJ 'W-!~ f ~- ''1 ENGINE FUEL SYSTEM The engine fuel system (figure FO-4) pro\idcs fuel re- quired for combu s tion, cOl1lrols engine variable geometry actuation, and provides engine oil cooling. Fuel is supplied to the engine fuel pumps, where it is pre~ s urized and directed to Ihe fuel control. From the fuel cont ro l, metered fuel passes 1hrough 1ill' engi ne oil cooler to the dislribution vahc. Change 7 1-3 T.O. 1A-10A-1 TF-34 ENGINE FAN COMPRESSOR ACCESSORY DRIVES PRESSURE TURBINE LOW PRESSURE TURBINE T-ICA- 1 -32 Figure 1-3 ENGINE FUEL CONTROL Th e engi ne fuel control (figure FO-4) is a hydro- mechani ca l type which modulates fuel flow to maintain a constant core speed as called for by throttle position. An electrical control unit regulates fuel flow at maximum power to maintain ITT limits. In the event of an ITT control unit electrical failure, the sys tem can be disabled with the engine fuel flow sw itch . In this mode, the engine will be speed- co ntroll ed throughout the entire range of operation, requiring pilot monitoring to prevent engine overt em perat ure . 1-4 The fuel control also prevents compressor discharge pressure from exceeding structural limits of the compressor. At sea level static, this limit is normally encountered at maximum power when engine inlet temperature is 0 of or colder. The limit can also be encountered on a standard day at sea level above approximately 330 knots. In this case it will not be possible to obtain rated ITT. The fuel control automatically controls the position of the compressor inlet guide vanes and the first five stator stages to prevent compressor stall . The engine fuel control does not require electrical power, but the ITT control unit is powered by the auxiliary AC essential bus. Engine Fuel Flow Indicators A fuel flo" indicator (43, figure FO-I) is provided for each engine. They show fu el flow in pounds per hour. The indicators are powered by the right AC bus. Engine Fuel Flow Switches Two engine fuel flow switches (figure 1-4), one for each cngine, are located on the engine co ntrol panel. These switches are placarded ENG FUEL FLOW L and R and each ~\\itch has two positions, placarded NORM and OVERRIDE. With the switch in NORM, the engine fuel flow' is scheduled on the basis of throttle position and limited to the m ,j ximum power trim setting hy the ITT amplifier. In the event of an ITT amplifier failure, the temperature control sy~tem can be deactivated by placing the appropriate switch in OVERRIDE. \Vhen this is donc, the engine will be speed controlled by I he t hroltle posit ion alone. Selec- tion of OVERRIDE when ITT is b el ow the maximum power trim setting will produce no change in engine operation. The engine fuel flow switches are powered by the auxiliary AC essential bus. THROTTLES A mechan:cal throll!e (figure 1-4) controls the opera- tion of each engine . Each throttle has three positive stop positions placarded OFF, IDLE, and MAX. To move from OFF to IDLE the throttle is raised and moved forward to the first stop position. To mOve to OFF the throttle is retarded to the IDLE stop, then raised and moved aft to OFF , The DC fuel pump is energized when either throttle is positioned to IDLE or above, and there is no pressure from the left main tank boost pump. On [i] and ~, when the throttle is at IDLE stop, the following actions take I place provided engine core rpm is below 56~o, and electrical power and an air source are available. • ATS valve opens causing the ENG START CYCLE light to come on • ECS shutoff valve closes • Both en gine bleed air shutoff valves open • Ignition is supplied to the engine • Fuel is supplied when engine rotation starts. T.O. 1A-10A-1 Engine speed is normally controlled by the throttle. Under certain flight conditions, the engine fuel con- trol override s the throttle to protect the engine from ovcrtemperature, overpressure, and compressor stall. Switches and controls located on the throttle are as shown on figure 1-4. Throttle Friction Control Throllle friction for both throttles is controlled by means of the friction control located on the throttle quadrant (figure 1-4). ENGINE IGNITION SYSTEM Ignition is supplied by two ignitors in each engine. The ignition ignitors are powered hy the AC essential bus and actuated by DC powered relays. The ignitor circcits are protected by two ENG IGNITOR (LlR-1 and L/R-2) circuit breakers. Each circuit breaker pro- tects an ignition circuit in each engine; therefore. to totally disable the ignition circuit to one engine, it IS necessary to open both circuit breakers. Engine Ignition Controls The ignition system is pilot-actuated through throttle position (IDLE. core rpm below 56aJo) [I] ~, the ignition button W [IT], or by the IGN function of the engine operate switch ITIJ (figures 1-4 and 1-5). On IIJ. ignition is provided to both engines while the gun trigger is depressed to the second detent and for 30 seconds after gun trigger release. On GJ, ig- nition is provided to both engines \'ihile the gun trig- ger is depressed to the second detent or during ac- tivation of stall warning chopped tone, and continues for 1 second after gun uigger release or termination of stall warning . BLEED AIR SYSTEM Bleed air from each engine, from the APU, and from a ground receptacle are routed to a common manifold (figure 1-7). The bleed air supply system furnishes air for the following: • Engine starter system • Environment control system • Windshield rain removal and wash system • Canopy de-fog system Change 6 1-5 T.O . 1A-10A - 1 THROTTLE QUADRANT INCLUDING ENGINE CONTROL PANEL MASTER LEFT EXTERIOR THROTTLE LIGHTS SWITCH MISSILE REJE CT/U NCAGE SWITCH * TIME SHARED FUNCTION 1-6 FLARE DISPENSING BUTTON [ill OR RIGHT ENGINE IGNITION BUTTON mAND [ill " M I C" THROTTLE FRICTION CO NTR OL Figure 1-4 CHAFF DISPENSING BUTTON !ill OR LEFT ENGINE IGNITION BUTTON [I] AND [ill MISSILE SEEKER HEAD SLEW/TRACK CONTROL l- I OA- I-3 T.O. 1A - 10A -1 ENGINE IGNITION SYSTEM DIFFERENCES PROVIDED TO OLD AUTO MANUAL I ~ ---IGNITION -'1 ENGINES START ~ START 621 NEW AUTO START ~ - ------+---------- --+-- ._-------_._--_. __ . BY THROTTLE THROTTLE IN IDLE IF IGNITION NOT PROVIDED THROTTLE IN IDLE IF ENGINE CORE RPM IS BELOW 56 % I POSITION ENGINE CORE RPM IS I BELOW 56 % r -;; BU~TON-;-T;;;VIDED WHEN BUTTONS ARE PRESSED AND IGNITION NOT PROVIDED . THESE BUTTONS ARE USED ON THROTTLE I FOR 30 SEC . AFTER BUTTONS ARE RELEASED , FOR CHAFF/FLARE DISPENSING . I L I REGARDLESS OF THROTTLE POSITION, RPM OR Ii . ENGINE OPERATE SWITCH POSITION I S-Y ENGINE I IGNITION NOT PROVIDED '---+-P'-R-O-V-ID- E-D- W-H- E-N- E- N- G- INE OPERATE SWITCH IS 1I OPERATE i PLACED IN IGN AND FOR 30 SEC. AFTER SWITCH IS L SWITCH I ~~~ASED . REGARDLESS OF THROTTLE POSITION OR I Figure 1-5 • Ca nop y seal • Anti-g suit • External tank pressurization. Each of the above system s is described in detail under the respective section s. Bleed air supplied from th e engine is controlled by a s hutoff valve adja cent to each en gi ne. Both valves are opened or closed simultaneously by the bleed air switch. Bleed air from the APU and / or ext ernal source is not con- trolled by the bleed air switch. A temperature senso r is provided adjacent to the manifold for bleed air leak detection. The bleed air leak detection system is power ed by the auxiliary AC esse ntial bu s. Bleed Air Switch The bleed a ir switch (figure 1-30), on the environ - mental panel, is a two-position lever-locked switCh, with position s placarded BLEED AIR and OFF. BLEED AIR opens both engin e bleed air valves and will provide bleed air to any bleed air system selected by appropriate controls. OFF closes the valves except during engine s tart. The switch is powered by the DC essential bu s. Fire Detect/Bleed Air leak Test Button The fire detect / bleed air leak test button (figure I-o S) is a pu s h-to-te st button, placarded FIRE DET EC T BLEED AIR LEAK TEST. Depr ess ing the swi tch checks the bleed air sensors, fire de tection senso rs, and associated warning lights. If the circuit is i ntact, BLEED AIR LEAK caution light on the ca ution light panel and the MASTER CAUTION, FIRE (L ENG) PULL , FIRE (R ENG) PULL, FIRE (APU) PULL lights will come on. The test button is powered by the auxiliary DC essential bus. Bleed Air leak Caution light The bl eed air lines upstream from the pr ecoo ler are monitored by a leak detection system. Upon se nsing a temperature of 400 °F or more , the system res pond s by activating the BLEED AIR LEA K caution light on the caution light panel (figure 1- (5). ENGINE OPERATE SWITCHES ~ Two engine operate switches (figure 1-4), one for each engine, are located on the engine control panel. These lever-locked switches are placa rded ENG OPER Land R, with each switch havin g two position s placarded NORM and MOTOR . NORM is us ed during normal engine operation and for engine 1-7 T.O. 1A-10A-l starting. MOTOR is used for air-purging of excessive fuel, cooling the engine, or manual starting. When the switch is moved to MOTOR, the following ac- lions are accomplished, provided electrical power and an air source are available: • AT5 valve opens causing the ENG 5T ART CYCLE light to come on • EC5 shutoff valve closes • Both engine bleed air shutoff valves open . NOTE • The throttle must be in OFF or IDLE in order to motor the engine . The engine operate switches are powered by the DC essential bus. ENGINE OPERATE SWITCHES ~ Two engine operate switches (figure 1-4), one for each engine, are located on the engine control panel. These lever-locked switches are placarded ENG OPER Land R, with each switch having two posi- tions pla<.:arded NORM and MOTOR. NORM is used during normal engine operation. MOTOR is used for engine starts, air-purging of excessive fuel, and cool- ing the engine. When the switch is moved to MOTOR, the following actions are accomplished, provided electrical power and an air sour<.:e are available: • ATS valve opens causing the ENG START CYCLE light to come on • ECS shutoff vahe closes • Both engine bleed air shutoff valves open. NOTE • The throttle must be in OFF or IDLE in order to motor the engine. The engine operate switches are powered by the DC essential bus . 1-8 ENGINE OPERATE SWITCHES 1451 Two engine operate switches (figure 1-4), one for each engine, are located on the engine control panel. These switches are placarded ENG OPER Land R, with each switch having three positions placarded IGN, NORM. and MOTOR. The switches are spring- loaded from IGN to NORM positions. The switches must be raised \-vhen moving between NORM and MOTOR. Momentarily placing the engine operate switch to IGN will supply ignition to the <.:orrespond- ing engine for 30 seconds. regardless of the throttle position or engine core rpm. NORM is used during normal engine operation and for engine starting. MOTOR is used for air-purging of excessive fuel, cooling the engine. or manual starting. When the switch is moved to MOTOR. the following actions are accomplished, provided electri<.:al power and an air source are available: • AT5 valve opens causing the ENG 5T ART CYCLE light to come on • ECS shutoff valve closes • Both engine bleed air shutoff valves open. NOTE • The throttle must be in OFF or IDLE in order to motor the engine. The engine operate switches are powered by the DC essential bus. ENGINE START SYSTEM Engine starts require low pressure air to power the A T5 unit mounted on the engine. Air may be ob- tained from the following sources: • APU • Crossbleed air from an operating engine (85070 core rpm minimum) • External pneumatic power unit. ...., .... o ENGINE START SYSTEM r- ______ ~ A~ _______ \ ~ L ENG EN G OPER ENG OPER L ENG INE THROTILE SWITCH NOTE AN ELECTRICAL INTERLOCK ARRANGEMENT IN THE AIR TURBINE CIRCUIT PREVENTS STARTING BOTH ENGINES AT THE SAME TIME R ENG INE THROTTLE SWITCH I GN BUDO N L N OR M L IG N (ON THRODLEI , -."., I' , I', ( \-1 (J> NORM MO TOR l DEENERGIZED AUTOMATICALLY WHEN ENGINE IS SELF SUFFICIENT mrm BLEED AIR (J> OFF 1':" \ 1 ) I ,- / T MOTOR ____ I I CLOSES VALVE REGA RDLES S OF SWITCH POSITION DEENERGIZED TO OPEN VALVE REGARDLESS OF SWITCH POSITION DURING ENGI NE START APU START @----- OFF DEE NERGI ZE D TO OPEN VA LV E REGARDLESS OF SWITCH POSITIO N DURI NG E NG IN E START ~~G~ CLOS ES VALVE REGARDLESS OF SWITCH POSITIO N mrmDEENERGIZED AUTOMA TI CALLY WHEN ENGINE IS SELF SUFFICIENT ~ ENG OPER : : R NORM : ..J, I It"", t ,-;-, I \~~ :~f' I R ENG \ / I IGN BUDON MOTOR _I (ON THROTILEI '-----vr---' '-v--' rnWl rm LEFT ENGINE RIGHT ENGINE [ill aircraft contain a manual sta rting system. This requires that the engine o perate switch be positioned to MOTOR, ignition button depresse d, and throttle in IDLE. [TIl aircraft contain an a utomatic engine starting system . Automatic engine sta rting will be initiated when the throttle is mo ved to IDLE, provided the engine core rpm is below 560-/0, electrical power and a n air so urce are available. The following events occur: • ATS valve opens, allowing engine to rotate • ECS shutoff valve closes • Both engine bleed air shu toff va lves open • Ignition is suppli ed for a minimum of 30 seconds • Fuel is provid ed after engine s tart s to rotate • A TS va lve closes 10 seco nds after engine reaches 56~0 core rpm • ECS valve opens 10 seco nd s after engine reaches 56~0 core rpm • Both engine bleed air s hutoff valves close 10 seco nd s after engine reaches 560:10 core rpm. Engine Start Cycle Caution Light Th e engine st art cyc le caution li ght (figure 1-65), on the caution li ght pa ne l, is pl aca rded ENG START CYCLE and will come on whenever the A TS valve is opened. Engine Core Speed Indicators An engine core speed indicator (40, figure Fa-I) is pr ov ided for each engine. The indicator s display the speed of th e compressor core in percent rpm. The sys tem is independent of th e airc raft electrical system except for in s trum e nt li g htin g purposes . Engine Fan Speed Indicators A fan speed indicator (42, figure Fa-I) is provided for each engine. Th e indicators display the fan speed in perce nt rpm and are powered by the auxiliary AC essential bus . T.O. 1A-10A-1 Engine Interstage Turbine Temperature Indicators An ITT indicator (39, figure Fa-I) is provided for each engine . The indicator s di s pla y the te mp era ture between the high and low pressure turbin e ~ections in degrees C. A warning flag placarded OFF will appear in a window to ind icate power loss . The indicators are powered by the auxiliary AC es~ential bu s. Engine Overheat Caution Lights An engine overheat caution light (figure 1- 65), on th e caution light panel, is provided for each engine. The lights are placarded L ENG HOT and R. ENG HOT and will come on if the ITT indicator exceeds 835 °C / ~ 880° C. FIRE EXTINGUISHING SYSTEM The fire extinguishing syst em is available to both engines and to the APU co mpartment / area . It consists of fire extingu is hing agent st ored in two independently actuated pre ss uri zed bottles l ocated in the fuselage. Either bottle may be di scharged to either engine nacelle or the APU compartme nt / area by pulling the appropriate fire h and le a nd actua t ing the disc harge switch. The system is dearmed by pushing the appropriate fire h and le in. The fire extinguishing system operates on battery bus power . However, fire detection and fu el/ bleed air shu tof f functions require auxiliary DC essential a nd DC essential bus power. NOTE • Extinguishing agent will no t put o ut an engine core fire as it doc s not discharge into the core. Extin gui shing agen t sh ou ld be used if the fire li ght comes o n , indicat- ing fire in the engine nacelle . FIRE DETECTION SYSTEM Fire detection is provided for in both engine nace ll es, and in the APU area by continuous temperature- sensitive elements . The fire warning light in the ap- plicable left or right engine fire h and le will come o n when the entire sensor element is heated to app rox- imately 650°F. The APU fire and overheat sys tem is similar to the engine fire sys tem except th at the warning light is in the APU fire handle. The APU fire detection includes coverage for the adjacent hydraulic , fuel, Change 6 1-11 T. O. 1A-10A-1 electrical, flight control and environmental control subsystems equipment installed in the fuselage be- tween the fuel tank aft bulkhead and the frame aft of the APU. Both systems are powered by the aux- iliary DC essential bus. The system is tested by depressing the FIRE DETECT BLEED AIR LEAK TEST button. Refer to Bleed Air System for test function description. • The fire detection system may not detect an engine nacelle or APU compartment fire/over- heat condition of high intensity and short duration. ENGINE AND APU FIRE HANDLES Three T-shaped handles (8 , 9, 10, figure FO-i), located in the glareshield on the instrument panel, provide fire warning for the engine nacelles or the APU when illuminated. The handles are labeled FIRE (L ENG) PULL, FIRE (APU) PULL, and FIRE (R ENG) PULL. The lights are powered by the auxiliary DC essential bus . By pulling the appropriate fire handle the pilot initiates the following actions: Engine Fire • Arms the fire extinguishing system to respective engine nacelle (provided battery bus power is available) • Cuts off fuel flow to the affected engine by closing the motorized main fuel shutoff valve (pro- vided DC essential bus power is available) • Closes the bleed air shutoff valve from the af- fected engine (provided DC essential bus power is available). APU Fire • Arms fire extinguishing system to APU compart- ment (provided battery bus power is available). • Cuts off fuel flow to the APU fuel control by closing the solenoid operated APU fuel shutoff valve (provided DC essential bus power is available) . • With more than one fire handle pulled, the fire extinguishing agent will be dis- charged into all areas selected. The quan- tity then discharged into the areas selected may be insufficient to extinguish that fire. 1-12 Change 7 Fire Extinguishing Agent Discharge Switch The fire extinguishing agent discharge switch (II, figure FO-I), placarded FIRE EXTING DISCH, is located on the right side of the giareshield above the instrument panel. The switch has three unlabeled positions. When the switch is moved either left or ri2ht, an extinguisher bottle is discharged and agent is directed to the engine or APU compartment selected by the fire handle. The switch will remain in the selected position to indicate which extinguisher bottle was discharged. The fire extinguisher bottles can be armed and discharged if battery bus power is available. AUXILIARY POWER UNIT The APU (figure 1-2) supplies air for engine starting, drives a generator for aircraft electrical power, and can drive a hydraulic pump to pressurize the aircraft hydraulic system for ground maintenance functions. The unit is located in the aft fuselage between the two engines and is provided with safety devices that shut down the APU when certain operating limita- tions are exceeded. Fuel for APU starting is supplied by the DC fuel pump. APU controls are powered by the DC essential bus. APU SWITCH The APU switch (figure 1-4) is a two-position switch, placarded START and OFF. START supplies DC essential bus power to operate the DC fuel pump, open the APU fuel valve, enable APU compartment cooling, energize the APU starter, and enable the APU EGT gauge and APU tachometer. APU GENERATOR SWITCH The APU generator switch (figure 1-11), placarded APU GEN, is a two-position lever-lock switch, placarded PWR and OFF / RESET. When in PWR, the APU generator powers an APU hydraulic pump cooling fan and electrical system busses, provided the busses are not powered by an engine generator or ex- ternal power. If the APU generator drops off the line, the system may be reset by momentarily placing the APU generator switch in OFF I RESET and re- turning it to PWR. APU GENERATOR CAUTION LIGHT The APU generator caution light (figure 1-65) is placarded APU GEN. The light is inoperative when the APU generator switch is in OFF/RESET. With the APU generator switch in PWR, Light on indicates: • Inoperable generator • APU operating with generator switch in PWR but aircraft busses being powered by either external power or engine generator(s) • During this mode of operation the caution light is on regardless of APU generator output. There is no indication that the APU hydraulic pump cooling fan is not receiving power. Overheating of the pump could result from extended operation with a failed APU generator either in the air or on the ground. • APU not running and generator switch in PWR. NOTE • If the APU is operating with the APU generator switch in PWR, and APU is shut down and restarted, the APU generator will remain inoperative until the APU generator switch is momentarily posi- tioned to OFF/RESET, then returned to PWR. Light off indicates: • APU powering aircraft busses. APU TACHOMETER The APU tachometer (44, figure FO-I) indicates the speed of the APU in percent rpm. DC essential bus power is required to enable the APU tachometer through the APU switch. T.O. 1A-10A-1 APU TEMPERATURE INDICATOR The APU temperature indicator (45, figure FO-I) indicates the turbine discharge temperature in degrees C. The indicator is powered by the DC essential bus and is enabled by the APU switch. APU OPERATION APU starting requires only DC essential bus power and a fuel supply. When the APU start switch is positioned to START, the DC essential bus power operates the DC fuel pump, opens the APU fuel valve (aft fuel tank mounted), and energizes the APU starter. The starter rotates the APU compressor and, at approximately lOaJo rpm, the APU fuel valve (APU mounted) opens and fuel and ignition are sup- plied to the APU. Acceleration of the APU continues until at approximately 6OaJo rpm the starter disen- gages. At approximately 95aJo rpm, ignition is ter- minated and the APU is self-sustaining. APU speed and turbine discharge temperature are automatically controlled. The APU will stabilize at lOO( ± 3)aJo rpm in approximately 60 seconds. APU starts can be made up to an altitude of 15,000 feet (most cases up to 20,000 feet) and the APU output will be sufficient to start an engine up to an altitude of 10,000 feet (most cases up to 15,000 feet). The APU will operate during negative g conditions for approximately 10 seconds. On ~, APU will automatically shutdown during ground operation if the APU EGT is excessive, APU rpm is excessive, APU oil pressure is low, or the APU fire warning system is activated. On ~, APU overtemperature shutdown is disabled during ground engine start cycle plus 4 seconds. Once the weight is off the landing gear, the APU will automatically shutdown only if the rpm is excessive or the oil pressure is low. AIRCRAFT FUEL SYSTEM The aircraft fuel supply system (figure FO-4) consists of two internal wing tanks (left and right wing), and two tandem-mounted fuselage tanks (left main-aft and right main-forward). Up to three external (pylon) tanks may be carried; one tank on each wing and one on the fuselage centerline. The fuel supply system 1-13 TO 1A-10A-1 operates as two independent subsystems, with the left wing and left main tank feeding the left engine and the APU, and the right wing and right main tank feeding the right engine. The two subsystems can be interconnected by opening crossfeed valves (con- trolled by a single switch in the cockpit) to allow pressurized fuel flow to both engines and the APU from either subsystem. In addition, the two main tanks can be interconnected by opening a tank gate valve. The main tank sumps are self-sealing bladder cells. Each self-sealing sump contains approximately 900 pounds of fuel. The upper portion of the cells are tear -resistant bladders. The wing tanks are in- tegral within the wing structure and do not have bladder cells. Foam is incorporated in each tank to prevent fuel tank explosion. Boost pressure is pro- vided by boost pumps located in each main and wing tank. A DC boost pump, located in the left main tank is used during engine and APU starts if the left main boost pump is inoperative. For negative g flight, collector tanks will supply the engine with sufficient fuel for 10 seconds operation at MAX power. In the event of a main tank boost pump failure, the affected engine will suction-feed from the failed tank for all power settings up to an altitude of 10,000 feet (most cases up to 20,000 feet). The wing tank boost pumps operate at a higher pressure and override the main tank boost pumps to automatically empty the wing tanks first. The main fuel feed lines to each engine, and to the APU, contain shutoff valves that are controlled by the fire handles. These shutoff valves allow for isola- tion of the fuel feed system outside the tanks. Fuel in the external tanks is transferred to the main or wing tanks by pressure from the bleed air system. Fuel tank sump drains are provided for each tank. Drain valves can be opened externally . Fuel cavity drains are provided in each main tank, and protrude through the aircraft skin to give an indication of fuel cell leaks. The wing tanks have a dual-level refueling shutoff valve. The valve closes when the tank is full and will not reopen unless the fuel level drops approximately 400 pounds or a time delay of approximately 10 minutes has elapsed. Wing tanks cannot be topped I off unless the fuel level is below approximately 1,590 pounds or the fuel manifold has been unpressurized for the time delay period. This assures even fuel transfer from the external tanks. Therefore, during fuel transfer from the external tanks, the wing tank 1-14 Change 8 fuel quantity will drop approximately 400 pounds, then will fill to capacity. This cycling repeats until external fuel is depleted. During air refueling the wing tanks will not accept fuel unless the fuel level in the tanks has dropped approximately 400 pounds or the time delay has elapsed. The total fuel on board after refueling could be approximately SOO pounds less than total capacity. If total fuel capacity is re- quired during air refueling, the external tanks can be turned off sufficiently prior to refueling so that the wing tank quantity drops approximately 400 pounds or the time delay has elapsed. A single-point ground refueling receptacle, located in the leading edge of the left landing gear nacelle, per- mits refueling of each internal and external tank. A control panel, adjacent to the refueling receptacle, provides a means of ground checking the refueling valve shutoff. The panel also permits selective loading of any internal or external tank. Auxiliary DC essential bus power is required for refueling valve checks, selective tank filling, and to enable external tank filling. When the tanks are full the refueling valves are closed by a float valve in each tank. Fuel tank capacities are shown in the usable fuel quantity data table, figure I-S. Fuel grade and specification to be used are covered in the servicing diagram, figure 1-71. In addition to features previously mentioned, the following survivability features are built into the fuel system: • Single-point ground refueling and engine feed lines outside the tanks are self-sealing to prevent leaks • The fuel feed shutoff valves are inside the tanks to keep the engine feed lines dry after shutoff • Fill disable switches are provided to close off a damaged internal tank when air refueling. FUEL QUANTITY INDICATOR AND SELECTOR The fuel quantity indicator (figure 1-9) is provided to monitor the total fuel remaining, or fuel remaining in selected tanks. The digital readout is a continuous display of total fuel remaining including external, in pounds. The pointer display provides an indication of fuel in specific tanks as selected by the rotary selector switch. The left and right pointers indicate for the left and right fuel systems, respectively. The fuel TO 1A-10A-1 USABLE FUEL QUANTITY DATA POUNDS jp-4 jp-5 jp-8 FUEL TANK GALLONS (NOTE 1) (NOTE 2) (NOTE 3) L. MAIN 511 3 , 270 3,475 3,424 R. MAIN 511 3 , 270 3,475 3,424 L. WING 311 1, 990 2,115 2,084 R. WING 311 1, 990 2,115 2,084 TOTAL INTERNAL 1,644 10,520 11,180 11 , 016 CENTERLINE 600 3,840 4,080 4,020 L. WING 600 3,840 4,080 4,020 R. WING 600 3,840 4,080 4,020 TOTAL EXTERNAL 1, 800 11,520 12,240 12,060 TOTAL FUEL 3,444 22 ,040 23,420 23,076 NOTES: 1. FUEL WEIGHT BASED ON 6 .4 LBS PER GALLON. 2. FUEL WEIGHT BASED ON 6 .8 LBS PER GALLON . 3. FUEL WEIGHT BASED ON 6.7 LBS PER GALLON . FUEL QUANTITIES BASED ON FUEL TEMPERATURE OF 60 ( ± 40) OF (TO 42BI-I - 14) . Figure 1-8 quantit y indicator is powered by the auxi liary AC essential bus . Position s of the selector are as follows: • INT Left and right pointers indicate total internal fu el for respective system • MAIN - Left and right pointers indicate fuel in the respective main tank • WING - Left and right point ers indicate fuel in the respective wing tank • EXT WING • EXT CTR Left and right pointers indicate fuel in the respective wing pylon tank Left pointer indicat es fuel in the fuselage pylon tank . The right pointer will ze ro • TEST INO When button is depre ssed th e lef t and right pointers will read 3,OOO( ± 300) pound s eac h and th e digital readout will read 6,000( ± 400) pounds. When the TEST INO switc h is released the pointers and di gita l readout wi ll return to the normal positions. NOTE • The fuel quantity totalizer will read high if the left main tank quantity is below ap- proximately 500 pounds and a considerable quantity of the fuel remains in the other tanks . Th e percent error will decrease as the fuel remaining decreases. Howe ver, the indi vi dual tank readings obtained by utili z- ing the fuel qu a ntity se lector are not affected. Change 8 1-15 T.O . 1A-10A-1 FUEL QUANTITY INDICATOR AND SELECTOR I - Ior. · j - 33 Figure 1-9 LEFT AND RIGHT MAIN FUEL LOW CAUTION LIGHTS The left and right main fuel low caution lights (figure 1-65) are placarded L-MAIN FUEL LOW and R-MAIN FUEL LOW, respectively. When the L-MAIN FUEL LOW caution light comes on, fuel quantity in the left main fuel tank is 650 (+ ISO, -100) pounds. When the R-MAIN FUEL LOW caution light comes on, quantity in the right main fuel tank is 500 (+ ISO, -100) pounds. This condition can be verified at the fuel quantity indicator. The lights operate independently of the gauge. LEFT AND RIGHT FUEL PRESSURE CAUTION LIGHTS The left and right fuel pressure caution lights (figure 1-65), placarded L-FUEL PRESS and 1-16 R-FUEL PRESS, respect ively, come on to indicate low fuel pressure at the engine fuel feed lines. FUEL TANK VENT SYSTEM Each main and wing tank (figure FO-4) is vented independently to a vent collector tank located in the left main tank . Vent lines from the wing tanks also serve as return lines for a ny fuel collected in the vent tank . Fuel in the vent tank is vented to the wing tanks or over board. Foam is installed in the vent tank to provide fire and lightning protection for the fuel system . MAIN TANK BOOST PUMP SWITCHES Two main tank boo st pump sw itch es (figure 1-10) are placarded BOOST PUMPS , with positions L-MAIN-R and OFF . Land R supply left and right AC bus power to the respect i ve main boost pump. OFF deactivat es the respect i ve boost pump . LEFT AND RIGHT MAIN BOOST PUMP CAUTION LIGHTS The left and right main fuel tank boost pump cau- tion lights (figure 1-65), placarded L-MAIN PUMP and R-MAIN PUMP, r es pectively, come on when fuel pressure at the outlet of the indicated fuel boo st pump is low. FUEL SYSTEM CONTROL PANEL T.O. 1A - 10A -1 WING TANK BOOST PUMP SWITCHES Two wing tank boost pump switches (figure 1-10) are placarded BOOST PUMPS, with positions L-WING-R and OFF. Land R supply left and right AC bus power to the respective wing boost pump. The pumps will automatically stop when the tank float switch senses an empty tank. OFF deactivates the respective boost pump. LEFT AND RIGHT WING BOOST PUMP CAUTION LIGHTS The left and right wing fuel tank boost pump caution lights (figure 1-65), placarded L-WING PUMP and R-WING PUMP, respectively, come on when fuel pressure at the outlet of the indicated fuel boost pump is low. Figure 1-10 1-17 T.O . lA - l0A-l EXTERNAL TANK SWITCHES Two external tank switches (figure 1-10), placarded EXT TKS, are located on the fuel system control panel. One switch is placarded WING and OFF, the other is placarded FUS and OFF. WING and FUS supply auxiliary DC essential bus power to pressurize the external tanks, using bleed air. The fuel is trans- ferred to the main and wing tanks through the re- fueling manifold. When the tanks are empty or not transferring fuel, they are vented to the atmosphere. OFF depressurizes and vents the associated tank(s). However, if the left or right main tank low level switch is actuated due to low fuel, external fuel, if available, will automatically transfer to the main tanks even if the external tank switches are OFF. CROSSFEED SWITCH The cro ss feed switch (figure 1-10) is a two-position switch, placarded CROSSFEED and OFF. In CROSSFEED , two auxiliary DC essential bus- powered valves open to allow any operating boost pump to feed both engines. When OFF, the valves close, isolating the two fuel systems. TANK GATE SWITCH The tank gate switch (figure 1-10) is a two-position switch placarded TK GATE, with positions placarded OPEN and CLOSE . OPEN supplies auxiliary DC essential bus power to open the gate valve linking the left and right main fuel tanks. Fuel in the main tanks will be below the tank gate valve and will not transfer in level flight when the fuel level is below 1,300 pounds in each main tank. The sump fuel will not flow between the tanks. CLOSE closes the gate valve. If the tank gate is used in other than relatively level unaccelerated flight, monitor fuel quantity for exces- I sively large cg shift due to fuel transfer. Fuel venting may be noted in this situation. LEFT AND RIGHT TANKS UNEQUAL CAUTION LIGHT The left and right fuel tanks unequal caution light (figure 1-65), placarded L-R TKS UNEQUAL, comes on when an imbalance of 750( ± 250) pounds in fuel 1-18 Change 1 quantity is sensed between the two main fuselage tanks. This condition ma y be verified by checking the fuel quantity indicator. FUEL SYSTEM OPERATION Normally, fuel system operation is automatic except for selecting external tanks. The main and wing tank boost pump switches are positioned to Land R. The tank gate switch is positioned to CLOSE. The cross- feed switch is positioned to OFF. With the battery switch in PWR , the DC boost pump is energized when the APU switch is positioned to START, or if either throttle is forward of OFF, and the left main boost pump is inoperative. The DC boost pump supplies fuel to the APU and the left engine. When the left and right AC busses are energized, the left and right main and wing tank boost pumps will operate. External tank fuel will be transferred to the internal tanks as fuel is used, until the external tanks are empty. The wing boost pumps will then supply the respective engine with fuel until the wing tanks are empty, at which time the wing tank boost pumps will automatically shut off. The main boost pumps will then supply the respective engine with the remainder of the fuel in the airplane . In the event of a wing tank boost pump failure, wing tank fuel will gravity feed to its associated main tank. Gravity feed of a full wing tank will not occur until the main tank fuel level is below approximately 600 pounds. Gravity feed of a partially full wing tank will occur at lower main tank fuel level. Dual check valve units in each wing tank gravity feed line prevent reverse fuel flow from the main tanks back into the wing tanks. When carrying external tanks, fuel sequencing will be as follows: • External wing tanks • External fuselage tank • Internal fuel. WARNING I • Feeding fuel simultaneously from external wing and fuselage tanks will cause cg shift that may exceed allowable limits. AIR REFUELING SYSTEM The aircraft can be refueled inflight from a boom- equipped tanker. The aircraft is equipped with a UARRSI (figure 1-2), located forward of the cockpit. By positioning a lever on the fuel system control T.O. 1A-10A-1 panel, a flush (slipway) door, powered by the right hydraulic system, folds down into the fuselage to expose the air refueling receptacle and to provide a slipway to guide the tanker boom. When the tanker boom is inserted in the receptacle, the nozzle latch rollers are actuated to the locked position, and Change 1 1-18A/(1-18B blank) I refueling transfer commences. Fuel transfer through the receptacle is distributed to the main and wing tanks, and to external tanks if carried. Through use of the fill disable switches, located on the fuel system control panel, the pilot can prevent fuel from enter- ing any specific internal tank suspected of being damaged. As each tank is filled, float-operated fuel shutoff valves within each tank will close, preventing overfill. When refueling is completed, the disconnect of the boom nozzle will normally be accomplished by a signal from the tanker or by the receiver pilot depressing the air refuel disconnect/reset button on the control stick grip. An automatic disconnect will occur when both receiver and tanker systems are completely operational and one of the following occurs: • Excessive fuel pressure occurs in the receiver fuel manifold • The tanker boom limits are exceeded (see T.O. I-IC-I-26). Refer to T.O. I-IC-I for basic flight crew air refuel- ing procedures and T.O. I-IC-I-26 for A-IO flight crew air refueling procedures. If the right hydraulic system fails, the spring-loaded slipway door will open when the air refuel control is set to OPEN. The time for the door to open sufficiently to expose the receptacle is improved by reducing speed and will occur within approximately 3 minutes at 150 KIAS. Aerodynamic effect will open the door sufficiently to expose the receptacle lights and permit emergency "stiff boom" refueling with or without a READY light. Applying boom nozzle pressure on the slipway door should result in the slipway door downlock engaging and a READY light. The LATCHED and DISCONNECT lights will not come on in this case. Air Refuel Control The air refuel control (figure 1-10) is placarded RCVR, with two positions OPEN and CLOSE. When OPEN, the left DC bus powers the signal amplifier and the hydraulic control valve. The slipway door opens and the READY light comes on when the door is locked open. When the boom nozzle is inserted in the receptacle, the latches close, securing the nozzle, the LATCHED light comes on and the READY light goes off. After the nozzle is removed from the T.O. 1A-10A-1 receptacle the DISCONNECT light comes on. CLOSE directs hydraulic pressure to close the slipway door and the DISCONNECT light goes off. In the event of loss of hydraulic pressure, OPEN releases a lock allowing the spring-loaded slip way door to open. NOTE • Fuel in external tanks will not feed with the air refueling control in OPEN . Fill Disable Switches The four fill disable switches (figure 1-10) are similar to circuit breakers. Two switches are placarded L-MAIN-R and two are placarded L-WING-R . if a main or wing tank is damaged, pulling up the respec- tive switch prevents that tank from being refueled. The switches are powered by the left DC bus. Signal Amplifier Switch The signal amplifier switch (figure 1-10) provides for emergency refueling. The switch is placarded SIG AMPL and has two positions placarded NORM and OVERRIDE. During the normal refueling cycle, the switch remains in NORM, and air refueling system power and actuating signals function automatically. If a failure occurs, fuel may not be transferred or the tanker boom may not stay latched. In this case, the override switch should be placed to OVERRIDE. In OVERRIDE no signals are passed to the tanker, and the tanker cannot actuate the disconnect cycle. Disconnect is accomplished by depressing the air refuel disconnect/reset button on the control stick. The signal amplifier switch is powered by the left DC bus. Air Refuel Disconnect/Reset Button An air refuel disconnect/reset button (time shared with Maverick Track and AIM-9 11OV is provided on the control stick grip (figure 1-17 . Setting the air refuel control to OPEN activates this button. With the boom nozzle inserted in the receptacle and the LATCHED light on, a disconnect may be accomplished by depressing the air refuel discon- nect/reset button. If the DISCONNECT light is on, depressing the air refuel disconnect/reset button recycles the system to the ready mode. Change 8 1-19 I T.O. 1A-10A-1 Air Refuel Line Check Button ~ The air refuel line check button (figure 1-10) is a push- button switch placarded LINE CHECK . Momentarily depressing this button checks the air refuel manifold integrity through a time delay relay. If the manifold is damaged, in flowing fuel will be discharged overboard with the possibility of fire and explosion. When the button is depressed before operating the air refuel door, the internal tank shutoff valves are closed and the air purge valve opens allowing air to pressurize the air refueling manifold. One engine must be operating at 85070 core rpm or the APU must be operating to supply sufficient air pressure for the READY light to come on if the manifold is intact. The READY light comes on when the air pressure builds up in the manifold (approximately 1-3 minutes). The light will go off approximately 3 minutes afer the line check but- ton is depressed. However, the light will remain on as long as the wing tanks are above approximately 1625 pounds. If the READY light does not come on within 3 minutes after the line check button is depressed, the refuel manifold is damaged. In this case, air refueling should not be attempted unless absolutely necessary. Air Refuel Status Lights The air refueling status indication is provided by three lights (6, figure FO- I) placarded READY, LATCHED, and DISCONNECT. When the slipway door is fully open and locked, the READY light comes on. Once the tanker boom nozzle and the refueling receptacle are connected, the READY light goes out and the LATCHED light comes on. When the boom nozzle and refueling receptacle are disconnected for any reason, the LATCHED light will go out and the DISCONNECT light will come on. The DISCONNECT light will remain on until the air refuel control is moved 10 CLOSE or the air refuel disconnect / reset button is depressed. Air Refueling Exterior and Receptacle Lighting Aircraft lighting is provided for night refueling oper- ations. These systems are described under Lighting System, Exterior Lights, in this section . Air Refueling Intercommunications With some tankers, secure interphone is available between aircraft when the signal amplifier switch is in NORM and the LATCHED light is on. The intercom system is powered by the DC essential bus. Controls for this system are described under INTERCOM con- trol panel, in this section. 1-20 Change 6 ELECTRICAL POWER SYSTEM The electrical power system (figure FO-5) provides DC and AC power. The battery produces DC to power essential equipment which provides the aircraft with a limited instrument flight capability. The in- strument inverter changes DC from the battery to AC to power essential equipment. DC produced by the battery is adequate to start the APU. When operational, the APU generator produces sufficient AC and DC (through the converters) to power all electrical busses indefinitely, provided electrical load is minimized. With engines running, two generators take over production of AC and DC (through the converters) to power all busses. External power can also be used to power all AC and DC (through the converters) busses. The cockpit battery switch is shown in figure I-II and the cockpit circuit breaker panel is shown in figure 1-12. DC SYSTEM Battery The battery is a 24-volt nickel cadmium type, and supplies DC to the battery, DC essential, and aux- iliary DC essent ial busses. External Battery Switch The external battery switch (figure 1-71), has two positions placarded ON and OFF. The switch is spring-loaded to ON, and is held in OFF by a safety pin. When ON, this switch allows the battery to supply DC to the battery bus. Battery Bus The ballery bus (figure FO-5) provides DC so that certain equipment can be operated when the cockpit battery switch is off. Cockpit Battery Switch The cock pit battery switch (figure I-II), placarded BATTER Y, has two posit ions placarded PW Rand OFF. The switch is located on the electrical power control panel. PWR allows the battery to supply the DC essential and auxiliary DC essential busses. and OFF disconnects the battery from these busses. DC Essential Bus and Auxiliary DC Essential Bus The DC essential and auxiliary DC essential busses (figure FO-5) provide DC to equipment deemed essential for flight . Converters Left and right converters convert AC from the APU generator, external source, or generators to DC . When operational, the converters automatically replace the battery as the DC power source for the --- ------------------------ - T.O. 1A-10A-1 battery, DC essential, and auxiliary DC essential busses; and also supply the left DC, right DC, and DC armament busses . The converters normally provide DC to their respective left or right busses . However, either converter can supply both busses if one converter is not operating. Change 8 1-20A/(1-20B blank). T.O . 1A- 10 A - 1 ELECTRICAL POWER CONTROL PANEL Fig ur e I-II Converter Caution lights Th e left a nd right c onv erter ca ution li ght s (fi g ur e 1-65) are placard ed L C ONY a nd R CO N Y. Th ese lights will come on to indicat e failure of the associ- at ed converter. If either gene rator fail s , the associat ed convert er caution li ght should remain off, indicatin g automatic transfer to the operatin g sys tem. Left DC Bus, Right DC Bus, and DC Armament Bus The le ft D C, ri g ht DC , a nd DC ar mam ent bu sses (fig ure FO-5) pro vid e DC to mi ss io n s upport e qu ip - ment and tho se sys tems no t dee med essential to fli ght. AC SYSTEM Instrume nt Inver t er Th e in stru men t in verte r cha nges D C su pplied by th e batter y to AC. A C fro m the in verte r p ow ers th e A C essent ia l, a ux ili ary AC csse nt ia l, an d AC in str um ent tra ns fo rm er bu sses , wh en th e le ft and ri g ht A C bu sses a re n ot energized o r th e in st rum ent in ve rt er switc h is set to T E ST. AC Essential Bus, Au x iliary A C Es sen t ial Bus , and AC Instrument Transformer Bus Th e AC essenti al, au x il ia ry AC esse nti a l, a nd in - st ru me nt t ran s fo rm er busses po w'e r e qu i pm ent (fig ur e FO-5) needed fo r st art ing engincs a nd o pcra t ing engin e in st rum en ts . 1-21 T . O. 1A-10A -1 CIRCUIT BREAKER PANEL CIR CU IT BREAK ER IS PLACAR DED RUDDER AU TH LIMIT ON ~ r--AILERON---, SPSr. Ru DDER r--ELEVATOR---, l OISC R AUTH LIMIT l OISC R ® 0 @ 0 (5} r--AILERON ---, l TAB R o (5) ,-----ENGINE ---, l START R {D @ rEMER FUEL SHUTOFF-, l ENG R (D @ ,-EXT STORES---, ' 1 JElT ' 2 @ @ IFF @ UHF CDMM ill CDNVERTIR l @ EMER HAP @ APU CDNT @ DC FUEL PUMP (D STB Y ATT INO <D INTIR CDMM @ EMER TRIM ® LANO GEAR ID ~ENG IGNITOR--, LlR ' ! LlR·2 · @ACQ) ,-BLEEO AIR---, .th CONT R ~ <D MA STER CAUT @ PilOT HEAUC @ r- GENERATOR--, l CONT R @ ill ,----- AU X ESS BUS-------, OA 9B 0 C ill <D ® r--INVERTER-----, PWR CONT AUXESS BUS TIE BATTERY BUS TRANS ~ ® (6E) @ ® \, CB RESULT OF PO PP ED CB CB RESULT OF POPPED CB CB RESU LT OF POPPE D CB AILERO N DISCON NEC TOR STAYS IN LAST SPS & RU DD ER TR AV EL REMA INS ± 25° ELE VA TOR DISCO NN ECTOR STAYS IN LAS T DISC LlR POS ITIO N. RUDDER ABOVE 240 KIAS . LI MIT ER INOP. DISC LlR POS ITI ON. AILERON AIL SH IFTER IN LAS T POS ITI ON. AUTH LI MIT SPS TONES INOP . EMER TRIM EMER TRIM INOP. TA B LlR TAB LI GH T INOP. ASSOC IAT ED EMER FLAP EMER FL AP RETRAC T INOP. LA ND GEAR NORM LD G GEAR EXTE NDI HYD SHUTOFF VALVE - OPEN . APU CONT APU AND DC FUE L PUMP INO P. RETRACT. NORM WHEE L ENG INE ENG STARTERS AN D IG N INO P. STAR T LlR ASSOC IATED MAIN FUEL LOW AND MA IN PUMP LTS INOP. AU TO FEED ING OF EXT FUE L AT LOW FUEL INOP. DC FU EL DC FUEL PUMP INOP. PUMP STBY An SAl AVAILABLE FOR 9 MINUTES IND MAX IMUM . BRAKES , NOSEWHEEL STEER IN G, AND ANTI-SK ID IN OP. EN G IGNITOR A SET OF IGN IT ORS (ONE IN EAC H L' R-1I -2 ENG I IN OP WITH EACH CB THAT IS OU T. EMER FUEL FIRE HAND LEFUEL SHUTOFF - SH UTOFF INOP. ENG LlR INT ER COMM ALL INTERCOM PANE L AN D RWR AU DI O INOP (LD G GEA R AND AOA TO NES REMA IN!. BLEED AIR ASSOC IATED BLEED AIR VA LVE CO NT LlR OPEN. MASTER MASTER CAU T AND CAUT LI GHTS EXT STORES ONE OF DUA L Jm SYSTEMS INVERTER INVERTER DI SCO NN ECTED CAUTIO N INOP . Jm ill#2 INOP WIT H EACH CB TH AT PWR FROM BUS ES . NO EFF EC T IF PITOT HTR IN PI TOT TUB E ONLY INOP. IS OUT. GEN- ON LINE. HEAT AC IFF IFF INOP - ALL MO DES . INV ERTER INVERTER INOP. NO EFFECT GENERA TOR IF GEN-G N LI NE - NO EF FECT . UHF COMM UHF RAD IO INOP . CON T IF GEN - ON LI NE. CONT LlR IF GEN-OFF LI NE - WILL NO T COM E ON OR RESET . CONVER TER BAnERY OR R CO NVERTER L MUS T SUPP LY DC. AUX ESS AUX AC ES SEN TI AL AND 26V INS T BUS TRANSFORMER BUSES INOP BATIERY BATIERY ONLY DC SOURCE (LOS S OF ENG INSTR UMENTS!. BUS TRANS AVA ILAB LE TO BATIER Y BUS. AUX ESS AUX DC ES SENTI AL BUS INOP. BUS TIE ( 1- 10:. -1- J6 Fig ur e 1-12 1- 22 Change 6