BIS'ON u :3 ... 0 ,) C, ·.: ......; ~-.. ; · l ~ ·.ll ,~: , - AVIATOR QUALIFICATION COURSE · STUDENT POI NO. 47/48-4740J 4741J 4742J 4743J 4744 HANDOUT 4745, 4746, 4747, 4748, 475o FOR ...... JEILACCOC. HAWOC. U.S. ARMY AVIATION CENTER FT RUCKER, ALABAMA JUN E 1984 CHANGE 1 Written by W. \>J. Smith and SFC(P) RichardT. Christensen. DEPARTMEN T OF AVIATION SUBJECTS SYSTEMS TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama INDEX UH-60 "BLACK HAWK"~ AND IPC Subject a. Pretest Objectives . b. Introduction (4740) c. Power Plants and Related Systems (4743) d. Flight Controls and Hydraulic Systems (4747) . e. Automatic Flight Control System (AFCS) (4748) . . . f. Fuel Systems (4744) . . . g. Electrical Systems (4741) . . . . . . . . . . . . . h. Auxiliary Equipment (4742) . . . . . i • Power Train System (4745) . . . . . . . . . j. Rotor Systems (4746) . . . . . . . . . . . . k. Malfunction Analysis (4750) . . . . . . May 1984 Page 1 2 38 105 . . . 155 . . . . . 206 . . . 216 . . . . . 231 . . . . . 257 . . . . 282 . . . 306 Change 1 NOTES i i TASK CLUSTER ANNEX: A--UH-60 AIRCRAFT SYSTEMS PURPOSE: To provide the student with the knowledge necessary for job performance of a preflight inspection, operational checks, and emergency procedures for the UH-60. TOTAL HOURS: PEACETIME MOBILIZATION 50.0 50.0 POI FILE: 47 EA 1-1 UH-60 AIRCRAFT SYSTEMS PRETEST TYPE OF INSTRUCTION: PEACETIME HOURS MOBILIZATION E3 1.0 1.0 SCOPE: A 1-hour pretest to determine the student's ability to perform the objectives in lessons: Introduction (4740), Flight Controls and Hydraulic Systems (4747), AFCS (4748), Electrical Systems (4741), and Malfunction Analysis, Part I (4750). STANDARD: As stated in each lesson training objective. LESSON RE~ERENCES: TM 55-1520-237-10, chaps 2, 5, 8, and 9. TASKS: As referenced in the above lessons. EQUIPMENT: Classroom support equipment. INSTRUCTIONAL ELEMENT: DOAS, STD, ASB. POI FILE: 47 EA 2-1 UH-60 AIRCRAFT SYSTEMS PRE TEST TYPE OF INSTRUCTION: PEACETIME HOURS MOBILIZATION E3 1.0 1.0 SCOPE: A 1-hour pretest to determine the student's ability to perform theobjectives in lessons: Power Plants and Related Systems (4743 ), Fuel Systems(4744), Auxiliary Equipment (4742), Power Train System (4745), Rotor Systems(4746), and Malfunction Analysis, Part II (4750). STANDARD: As stated in each lesson training objective. LESSON REFERENCES: TM 55-1520-237-10, chaps 2, 5, 8, and 9. TASKS: As referenced in the above lessons. 1 EQUIPMENT: Classroom support 40 students. INSTRUCTIONAL ELEMENT: DOAS, STD, ASB. POI FILE: 47-4740-5 INTRODUCTION TYPE OF INS TRUCTION: PEACETIME HO URS MOBILIZATION c 5.0 5.0 LEARNING OBJECTIVE: Given four written questions co ntaining descriptive situations or operating conditions, from memory, the student will write IAW TM 55-1520-237-10 the- a. Normal crew requirements for UH-60 operations. b. Mi nimum crew requirements for UH-60 operations. c. Normal number of combat troops the UH-60 may carry. d. Optional troop seating capacity of the UH-60. e. Basic structural maximum gross weight of t he UH-60. f. Four steps in setting the parking brake. g. Purpose of the wheel brake puck wear indicator. h. Cockpit indication for the tailwheel in the locked position. STANDARD: Four of four questions must be answered correctly to satisfactorily complete th i s objective. LESSON REFERENCES: TM 55-1520-237-10, chaps 2 and 5. TASKS: This objective supports tasks 03-1402, 1501, 1502, 1506, 3501, 3511, 5001, and 6503. EQUIPMENT: UH-60 composite trainer; UH-60 power train systems trainer, DVC 1-118; UH-60 CDU/PDU display panel, FR 1302; and UH-60 caution/advisory panel, FR 1330. Bl dg 6005 (Yano Hall) classroom support 40 students. INSTRUCTIONAL ELEMENT: DOAS, STO, ASB. 2 GENERAL DESCRIPTION The UH-60A is a twin turbine engine, single rotor, semimonocoque fuselage, rotary wing helicopter. It has a crew of three and carries eleven combatequipped troops, and internal and external cargo. Two tiers of litters may be installed in the cabin, with two litters per tier. The pilot's compartment accommodates the pilot and copilot. There are dual controls and duplicateflight instruments. The crew chief/gunner is located behind the pilot's compartment, at the front of the cabin. The main rotor system has four blades made of titanium and fiberglass with aft swept tips. The rotor head has elastomeric (rubber-mounted) bearings for blade movements. A vibration absorbe~of the bifilar type is mounted on top of the rotor head. The propulsion system has two T700-GE-700 engines operating in parallel. The drive train consists of a main transmission,_ intermediate gearbox and tail rotor gearbox with interconnecting shafts. The transmissions are oillubricated. The tail rotor is a bearingless cross-beam design tilted 20 degrees upward. The nonretractable landing gear consists of two main landing gears and a tailwheel. The long strokes of both main and tailwheel oleos are designed to dissipate high sink speed landing energy. General Data Main Rotor Diameter 53.66ft (16.35m) Hub diameter 5 . 0 ft ( 1 • 52m) Number of blades 4 Blade chord 1.73/1.75 ft. (.52/.53m) Tail Rotor Diameter 11 . 0 ft. (3. 35m) Number of blades 4 (2 double end paddles) Blade chord 0.81 ft. (.24m) Horizontal Stabilator Area 45 sq. ft. (4.18 square meters) Span 14.33 ft. (4.35m) 3 NOTES 4 c :I: I m ~a )> aJ r ~ A :I: ~ A NOTES 6 a::: 0 10 a:::z _.o __. <> l-0.. en 0 LIJ -... z c( CJ " ..J LIJ 11.1 Cl) en a::: ::l 0 La. 1 c c( ..J - m c( 1- CI) ..J < 1 z 0 N -a::: a::: 0 0 1 z :t: 0 0 a::: zz _o c(..J :E~ 7 TROOP SEATS Cabin troop seats will accommodate up to 11 persons including the crew chief and gunner. There are mounting provisions for three additional seats. Each seat has a lap belt and torso harness for restraint in four directions. The seats are designed to protect the occupant in a crash. This is done by an attenuating system consisting of an energy-absorbing telescopic leg brace combined with two rotary attenuators on the seat back support cables. NOTES: 8 PILOT'S SEAT D \0 ' D COPILOT'S SEAT CREWCHIEF GUNNER I Glu 13 II r.1 I 14 I I 11 ~I 7 I I 10 I 3 6 I I 9 [,11 \. 5 I I 8 ( I I U35-394 HI DENSITY SEATING ARRANGEMENT CABIN FLOOR/CARGO TIEDOWN FITTINGS The cabin floor consists of three removable sections. Each section is constructed with a fiberglass bottom skin, honeycomb core, and a fiberglass top skin covering. The floor also has 17 fittings rated at 5,000 lbs. each for cargo tiedown and troop/litter installations. The cargo hook is reached through the cargo hook access panel in the floor. NOTES 10 CARGO RESTRAINT NET RING / ~r ........ CARGO REST.RAINT NET RING 3500 POUND CAPACITY EACH TOP OF CABIN FLOOR - - STA 379.0 STA 308.0 ~: :·: · :; ·~ 0 I I' ' I ' o' ' ' ' ' ' ' I o • '' I '' 't 0 ~:' . ; ij"HHH~~ u>------H' _.-: __ ,....,.._____ _~ EQUIPMENT STOWAGE -~~~-_ COMPARTMENTS CABIN FLOOR \ CARGO RESTRAINT RING LOCATIONS U35-756 67T1175 CONTROL ...... ACCESS N FAIRING OIL COOLER FAIRING (NO STEP AREA) · STEP AREAS APU/FIRE EXTINGUISHER ACCESS PANELS NO. 1 ENGINE FAIRING/ SERVICE PLATFORM OIL COOLER ACCESS PANELS MAIN ROTOR PYLON STEP AREAS AND PANELS 67T1012 U35-396 NOTES 13 EMERGENCY EQUIPMENT Emergency equipment includes two portable fire extinguishers: one on the right bulkhead behind the pilot's seat and one on the copilot's seat. There also are three first aid kits, mounted on pilot's and copilot's seats, and a ~crash axe mounted forward of the crew chief's seat on the aft bulkhead of the cockpit floor. NQIES.: 14 COPILOT'S SEAT FIRST\AID KIT FIRE EXTINGUISHER (COCf(PIT) PILOT'S SEAT FIRE EXTINGUISHER (CABIN) ...... V1 FIRST AID KIT COCKPIT FLOOR FIRST AID KIT (CABIN) CABIN FLOOR CRASH AX (CABIN) U35-774 EMERGENCY EQUIPMENT LOCATIONS 67T2183 en 1 - >< L&l Q z c en L&l (.) z r~ ~ ... c~ _.:\. ·.: >.=:::~··. ... . · D::5 ~:· ..... 1 z L&l > (.) z L&l D:: " L&l ~ Ll.l 16 15 14 ~ ~ 5 13 1D 12.1 1 UPPER CONSOLE SIDE PLATE 8 2 NO. 2 ENGINE FUEL SELECTOR LEVER 8. 1 3 NO. 2 ENGINE OFF/ FIRE T-HANDLE 9 4 NO.2 ENGINE POWER CONTROL LEVER 10 5 WINDSHIELD WIPER 11 6 INSTRUMENT PANEL GLARE SHIELD 12 7 INSTRUMENT PANEL 12 . 1 VENT / DEFOGGER CHAFF RELEASE SWITCH ASHTRAY PEDAL ADJUST LEVER MAP/ DATA CASE PARKING BRAKf L~VER FUEL BOOST PUMP PANEL 17 2 3 \ \~·''" ).\ ' ' 4 \ 5 111 11 8.1 13 STANDBY (MAGNETIC COMPASS) 14 NO. 1 ENGINE POWER CONTROL LEVER 15 NO . 1 ENGINE OFF/FIRE T· HANDLE 16 NO. 1 ENGINE FUEL SELECTOR LEVER 17 FREE -AIR TEMPERATURE GAGE 18 COCKPIT FLOODLIGHT CONTROL 19 UPPER CONSOLE PILOT AND COPILOT SEATS The pilot's and copilot's seats can be adjusted for leg length and height.The pilot's seat is on the right side, and the copil ot's is on the left. Eachseat consists of a one piece ceramic composite bucket attached to two . energyabsorption tubes that are designed to withstand any G-levels incurred insurvivable crash situations. Loads are reduced by allowing the seat andoccupant to move vertically as a single unit through a displacement of 12inches (30.4 em) from any position in the vertical adjustment range of theseat. Each seat is positioned on a standard track with the bucket directlyabove a recess in the cockpit floor. Occupant restraint is provided by ashoulder harness, lap belts, and a crotch belt. The seats provide a tiltfeature that allows the seat to be tilted back into the troop compartment for removal or treatment of a wounded or disabled pilot. Seat tilting can be donefrom the troop compartment. NOTES: 18 FIRST AID KITARMORED WING~ ~ERTIAREEL VERTICAL RELEASE . CONTROL Ill II ~{(j i\(H -I I II ,., ...... \0 TILT BACK RELEASE VERTICAL CONTROL ADJUST CONTROL STOWAGE PANEL INERTIA REEL CONTROL FORE AND AFT ADJUST CONTROL PILOT'S /COPILOT'S SEAT U35-393 67T2011 0 0 o\ OYSY CDNT ,.,II l TS WUN AfllliiC( IJOT CNTOII WU N WUN COlli fill WUN CM"flll CHAH --CPlT ~NO 2-DC U NH Iff AOf CMO CSL TIIIM OISr TUIN AlTM MODI YHf JM COMM,-IIOJI~ IUS 0 : ~~· ·: ~: <0<00<0®~0<00<000<000 ° SU Ul( CTIIO SU(CT JM SCfY SU AlT M WAI N sru NO. 1 CIRCUIT BREAKER PANEL ® 0) NO 2CON'tUTU r tLOT •SHlO 0 cMrnt Ufll!UCP ANT I I(( 0 ® <®> liCHTS0 ,-.JOIIM--AHT I '" ''"' NO.2 CIRCUIT BREAKER PANEL () ,---DC ESN TL BUS~----, () u ~ AUifUhQll aaaOiao 0 A "LOT CO "LOT ~Hf 1M OU CON fill r COMM SC" UT, UHJ C A U TI 8 AC•u• HOI ST o·<0 oos~oo© ---NO 0 r•n CONTII SHUll 0 DC ESSENTIAL BUS CIRCUIT BREAKER PANELS MISSION READINESS CIRCUIT BREAKER 0 0 PANEL (CABIN) Of ICf rWII <®> TAILIIOTOII 0 0 BAnERY AND BAnERY UTILITY BUS CIRCUIT BREAKER PANEL 20 INSTRUMENT PANEL The instrument panel is divided into three sections, two of which are identical. The flight instrument sections in front of the pilot and copilot are identical. In the center section of the instrument panel are located the following: 1) 2) 3) 4) Central display unit (CDU) Caution/advisory panel Ice detector system Ignition switch NOTES: 21 CAUTION/ADVISORY PANEL The caution/advisory panel is located to the left of center on the instrument panel. It contains 82 word legends relating to faults (caution) and operational status (advisory) of selected systems and components. A caution condition is displayed as an amber word legend. An advisory status condition is displayed as a green word legend. When any caution or advisory word legend is provisional (not presently in use) a soljd bar is displayed. Anytime one or more caution indications are detected, a signal is supplied to illuminate pilot and copilot master warning panel caution legends. Both master warning panel caution legends are resettable from either master warning panel. The number one and number two fuel low legends flash when a fuel low condition is detected. In turn, both master warning panel caution legends also flash. If either master caution is pressed to reset, both master caution legends go off--the fuel low legend continues to flash. There are two levels of intensity for all caution and advisory legends, bright o~ dim. If the pi lots flight lights control is rotated to off, all cauti on and advisory legends illuminate bright. Rotation of the control to increase intensity of pilots flight lights causes all caution and advisory legends to dim. PILOT'S DISPLAY UNITS A pilot's display unit is provided on the instrument panel in front of each pilot to provide hi m with rotor speed, power turbine speeds and torque readings. Torque readings are also displayed digitally. Each PDU also includes a lamp test button, three rotor overspeed warning lights, and a photoelectric sensor to automatically control the brilliance of the vertical scale lamps. INSTRUMENT DISPLAY SYSTEM The instrument display system, used in conjunction with engine and subsystem sensors (temperature, pressure, fuel and RPM) provide the pilots with engine and subsystem monitoring. The vertical instrument display system (VIDS) gives continuous indications of these parameters on vertical scales, digital readouts and status lights. The VIDS consists of three units: one central display unit (CDU) and two pilot's display units (PDU). The CDU is located in the center of tile instrument panel and displays fuel quantity, main transmission oil temperature and pressure, dual engi ne oil temperatures and pressures, dual turbine gas temperature (TGT) and dual compressor speeds (Ng). These vertical scales are color coded for easy reading and have no moving parts for high reliability. The CDU is also equipped with a lamp test button, dimming control, digital display ON/OFF switch, and a pair of warning lights 22 to indicate failure of either signal data converter. A photoelectric sensor automatically controls the brilliance of the VIDS. NOTES: 23 NOTES 24 ~~~~ I·;;.~: ~~~c:::Ei[J I:.:: ~:~ lc=:::J~I ;;:~:::I I ·.:.::f;~ II :~.·;:. II :.1;:. II ·~·:f;~ I I ..:~.. II -:;:;g: ~~~ I ..:~::::.. I~I··::..·· IJ .. ::::::.. I ~~~~ 1 •.::.:::... 1~~1...::·.~... 1 I;:.~:::::~~~~ ··:::::"1 I ·.:.-.::~· II :::;:: II ·;: :.·:: II .~·~.~·. I 1-~:""l~~r:::::::::::.:J ~~c=J~IO 1'::.:.·-::'II · ::::.. IQQI ·:;·::· I .... I '~;.~~~.. II ;:::::.~ I~I · ~;.":~"I • ~I·::::.~· lc::::EJQ;] I'".:,"'~;::-~~~~ :.:~·.:; I I ..-.:~':"~ II·::.':".~:.:.· IJ·::::·:.~~:.· II ..·.:~':".. I ~E::JI":::-.-:-· 11 :.:::::I I ~.~·· lc::::=:::J~I ·~ :~::· I @l' ..:.:-:~. n-..···IE::::JI ~~ @~~c=::J .. .. 0 0 • @ CNlCII LIST @ COll FIUCTION Aru CP COU UTI NO IUUS POWU lWII 11r11111 HUNUS fliGHT C~TitOLS CAUTION 111.010 Jill 1 UHf YHf INSfiiUMlNTS fAll WHHl SWNO l 2 J STAIIlAfOII @ US Jr$ 25 f'OWU HIT "" '" : ~ rJj TUT ...OC.lU IUD£ 0( tC( TUT _... :::r:~ a·~ ···· (§jo OAT ~/ lk (O! ~"' ... .ail G 1. RADAR ALTIMETER 2 BAROMETRIC ALTIMETER 3 VERTICAL SPEED INDICATOR 4 MASTER WARNING PANEL 5 VERTICAL SITUATION INDICATOR 6 HORIZONTAL SITUATION INDICATOR 7 AIRSPEED INDICATOR 8 STABILATOR POSITION PLACARD 9 STABILATOR POSITION INDICATOR 10 CIS MODE SELECTOR 11 VSI/HSI MODE SELECTOR 12 RADIO CALL PLACARD 13 PILOTS DISPLAY UNIT 14 CLOCK 15 LIQUID WATER CONTENT INDICATOR 16 BLADE DE-ICE CONTROL PANEL 17 INFRARED COUNTERMEASURE CONTROL PANEL 18 CENTRAL DISPLAY UNIT 19 RADAR WARNING INDICATOR 20 CHECKLIST 21 ENGINE IGNITION SWITCH 22 RADIO SELECT PLACARD 23 CAUTION/ADVISORY PANEL 24 NVG DIMMING CONTROL PANEL Instrument panel front view NOTES 26 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 $ ,---DC ESNTL BUS--;;;--' $aaoaioc::"LOT CO "LOY WHr rill OU CONU lffi\ ~ ,..-----CARGO HOOK [MER RlSE CONTR TEST NORM C~PT ~~(@ ~ SHORT All ~ ARMINd 11iBf SAfE (@>ARMED I'"" COIIIIII SCTY SIT, UHf CAUf/ IA.ClU, HOIST O®SOOOc:~ e -~~~~ e ,...-CABIN DOME l TS-, @ OFF BRT <® COPLT flT INSTR l TS WHITE r~ RED SECONDARY liGHTS ~@). ~~. RED fORMATION LTS CKPT flOOD liGHTS;@ OFF POSIT lOIII liGHTS (@)<® Off WHITE BRT ANTI COlliSION liGHTS DIM STEADY UPPER DAY Off~ @) ~APU~ r--GENERATORS--, ON EXT PWR BAT APU NO . I NO. 2 RESET TEST TEST TEST ~~ ~(@~i~ ~~~~ ON ON f ON ON ~ ON NO I ENG OVSP fiRE OET TEST NO. 2 ENG OVSP TEST A TEST B OPER TEST A TEST B ~~ 012 ~~ fUEl PUMP @) ~~:~S~~~ APU BOOST ENG :~ ~(@ fUEl PRIME APU SAS NO I (NG Ull "'$'iC ~~=:: SCHl 1' c::a1tllll A ~~Q.~.Q~~~~ T I WUN CONU T Ulll B ..-----APU~ I us APU ~~~~~t0 ~~~ -UC Pf CONTI lCONTI Jill GlN $ l INST INST DIT CONTI ~ 0 0 Lower Console 28 TAIL GEAR BOX-~ TAIL ROTOR DRIVE SHAFT N \0 \ INTERMEDIATE GEAR BOX MAIN TRANSMISSION U35-213 TRANSMISSION SYSTEM POWERTRAIN 6804007 FUEL AND LUBRICANT SPECIFICATIONS AND CAPACITIES SYSTEM Fuel Engine Oil Engine Starter Auxiliary Powerplant Transmission Oil Intermediate Gearbox Oil Tail Gearbox Oil Hydraulic First Stage Hydraulic Reservoir Second Stage Hydraulic Reservoir Backup Hydraulic Reservoir SPECIFICATION Primary, Grade JP-4 (NATO CODE F-40) Alternate: Grade JP-5 (NATO CODE F-44) · MIL-L-23699 (NATO CODE 0-156) MIL-L-7808 (NATO CODE 0-148) MIL-L-23699 (NATO CODE 0-156) MIL-L-23699 (NATO CODE 0-156) MIL-L-7808 (NATO CODE 0-148) MIL-L-23699 (NATO CODE 0-156) MIL-L-23699 (NATO CODE 0-156) MIL-L-23699 (NATO CODE 0-156) MIL-H-83282/MIL-H-5606 MIL-H-83282/MI L-H-5606 MIL-H-83282/MIL-H-5606 CAPACITY Gravity refueled: 181 US Gallons (685.15 liters) 1 . 7 US Gallons (6.43 liters) 200 cc. 3 US Quarts (2. 72 liters) 7.5 US Gallons (26 .4 liters) 1.4 Quarts (1.35 liters) 2.45 Quartes (2.21 liters) .86 US Quart (. 94 liters) .86 US Quart ( . 94 liters) .86 US Quart ( . 94 liters) 30 NOTES 31 MAIN LANDING GEAR One main landing gear is mounted on each side of the helicopter forward of the center of gravity. Each individual landing 9ear has a single wheel, a drag beam, and a two stage oleo shock strut. The lower stage will absorb energyfrom landings up to 10 feet per second (fps). Above 10 fps (3.04 m/sec) the upper stage and lower stage combine to absorb loads up to a sink rate of 39 fps (10.66 m/sec). TAIL LANDING GEAR The tail landing gear is mounted below the rear section of the tail cone. It has a two stage shock strut, tailwt1eel lock system, fork assembly, yokeassembly, and a wheel and tire. The fork assembly is the attachment point for the tailwheel and allows the wheel to swivel 360 degrees. The tailwheel can be locked in a trail position by use of a switch in the cockpit indicating LOCK or UNLK. The fork is locked by a lock pin operated by an electrical actuator through a bellcrank. WHEEL BRAKE SYSTEM Main landing gear wheels have hydraulic disc brakes. The self-contained system is operated by the two pedals on the pilot's and copilot's tail rotor pedals. The brakes have a visual brake puck wear indicator. Each wheel brake consists of two steel rotating discs, one stationary disc, 12 brake pucks, and a housing that contains six hydraulic pistons. The parking brake handle, marked PARKING BRAKE, is on the aft end of the lower console. The parkingbrakes are applied by depressing the toe brakes while holdin0 the handle out. An advisory light will go on indicating PARKING BRAKE ON. The parking brakes are released by depressing either pilot or copilot left brake pedal. NOTE: It is recommended that both pilot or copilot pedals be applied to release the parking brakes. NOTES 32 Chan g e 1 ,-,~ , , .,. . D '"': ~ ~ ! RAG , ' " ' . ~ ! ~ ~ ~ ~ BEAM -. . '. . . ' . ' .. ' . ... \ t I I I t ' \ t I I I . . ' "" . . . . . . --. . . . . . . . ~~:---P::~~~8~ ,, ,' ~ -rA - w BRAKE LINE w WHEEL SHOCK STRUT I'-"- BRAKE YOKE "" J ASSEMBLY JACK PAD TAIL WHEEL LOCK~ FORK ASSEMBLY LANDING GEAR INSTALLATION U35-878 1Nl012 PARKING PARKING BRAKE BRAKE SWITCH VALVE ... ' . :·· PARKING ·.·_,. ·····... f t: ~;!l~'t'~'~;:, ~ .... '• •• i -tl~·:t:~,, .· : : .,, /1 \. \ .·····:. ..·······..\.:~:: ········ · ········· ·· · · · ········ · · · ········ · ·········· w ACTUATOR -"' CAM A 1 WHEEL PILOT'S BRAKE MASTER DRAG BEAM CYLINDERS (WT. ON WHEELS) COPILOT'S COPILOT'S 1 MASTER BRAKE r;\ ~---- CYLINDERS PEDALS PARKING BRAKE/DRAG BEAM SYSTEM U35-553 LOCATION DIAGRAM 68F2054 w ln ~ROTATING DISC SECTION A-A WHEEL BRAKE WEAR PIN INDICATOR U35-725 67T2162 ~--------------------------~0~--------------------------~ TAILWHEELLOCK TAIL WHEEL ACTUATOR LOCK ACTUATOR UNLOCK SWITCH ' ' ··-------·------) LOCK /LEVER ' ' ' ' ' .. '-r -----::··~ 1 I "•. \..__) ··········... LOCK SWITCH LOCK PIN LOCK MANUAL SWITCH LEVER TAIL WHEEL LOCK SYSTEM TAIL WHEEL LOCK SYSTEM (ON HELICOPTERS 77-22714 (ON HELICOPTERS 77-22727, THRU 77-22726 AND 78-22960 77-22728 AND 78-22963 AND THRU 78-22962) SUBSEQUENT; AND HELICOPTERS MODIFIED BY KIT 70070-55026) 3£1 DEPARTMENT OF AVIATION SUBJECTS SYSTEMS TRAINING DIVISION UN-ITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama May 1984 File No. 47-4740-5 LESSON EVALUATION INTRODUCTION Complete the following questions by filling in the blank. 1. For UH-60 operations, ______ ______, and _______ are normal crew requirements. 2. The minimum crew is _______ and _______ required for UH-60 operations. 3. The UH-60 normally transports ____ combat troops. 4. The UH-60 is capable of transporting troops when optional seating is installed. 5. ~-=-=---pounds is the basic structural maximum gross weight of the UH-60. 6. To set the parking brake, the pi 1ot or copi 1ot must ---,.-.,--,----=,....-;-,,--toe brake pedals, pull the parking brake to the fully extended positions, then the toe brake pedals, next release the parking brake ______ 7. The brake puck wear indicators provide a _____ means of checking the brake pucks for excessive _______ 8. The letters appear on the TAILWHEEL switch to provide a cockpit indication when the tailwheel is in the locked position. 37 POI FILE: 47-4743-7 POWER PLANTS AND RELATED SYSTEMS TYPE OF INSTRUCTION: PEACETIME HOURS MOBILIZATION c PE 5.0 2.0 5.0 2.0 LEARNING OBJECTIVES: 1. Given eight written questions containing descriptive situations or operating conditions, from memory, the student will write IAW TM 55-1520-237-10 the- a. Maximim time to obtain TGT indication during engine start. b. Maximum temp for TGT indication before idle speed is reached. c. Maximum engine operating time after starting with NO oil pressure indication. d. Maximum engine operating time after starting with no percent RPM 1 and 2. e. Percent Ng at which ENGINE OUT warning light should be OFF. f. ENGINE STARTER caution light OFF at percent Ng. g. Percent RPM 1 and 2 to AVOID during starting. h. Percent TRQ matching 1 and 2 at 100 percent RPMR. i • Procedure for making engine overspeed system check. j. No. 1 and No. 2 ENGINE ANTI-ICE ON advisory light OFF at percent TRQ. k. Two functions of the engine anti-ice/start bleed valve. 1. Temperature range for the ENGINE INLET ANTI-ICE advisory light illumination when the engine ANTI-ICE switch is moved to the ON position. m. Three functions of the engine alternator. n. Torque matching tolerance of the ECU. o. ECU TGT limitation. STANDARD: Eight of eight questions must be answered correctly to satisfactorily complete this objective. 38 Change 1 2. Given two written questions containing pertinent indications and/or descriptive situations or operating conditions, from memory, the student will write lAW TM 55-1520-237-10 the limitations for- a. Maximum 10 second transient torque limit for single engine operation. b. Maximum 12 second transient percent RPM 1 and 2 limitation. c. Maximum TGT limitation for starting engine. d. Maximum time from Ng indication to idle speed at FAT above -20°C and a pressure altitude of 10,000 feet. STANDARD: Two of two questions must be answered correctly to satisfactorily complete the objective. 3. Given four written questions containing pertinent indications and/or descriptive situations or operating conditions, from memory, the student will write lAW TM 55-1520-237-10 the pilot action for- a. No. 1 ENGINE STARTER caution light , remains on after 65 percent Ng speed. · b. Abort start procedure for an engine TGT exceeding 850°C during start. c. During engine operation no oil pressure is indicated after 45 seconds (engine shutdown). d. Percent RPMR and No. 1 engine RPM increasing to 106 ~1 percent during flight. e. CAUTION to be observed with engine control lever in ECU LOCKOUT. f. No. 1 engine no percent RPM, no percent Ng, no percent TRQ, No. 1 engine TGT 860°C and increasing; No. 2 engine, percent RPMR 102 percent, 85 percent Ng, 20 percent TRQ, 560°C TGT. g. Crossbleed start of No. 2 engine using No. 1 as the air source. h. Compressor stall No. 1 engine in flight. STANDARD: Four of four questions must be answered correctly to satisfactorily complete the objective. 39 LESSON REFERENCE: TM 55-1520-237-10, chaps 2, 5, 8, and 9; student handout. TASK: This objective supports tasks 03-1402, 10-1004, 1005, 1501, 1502, 2002, 2501, 2502, 3006, 3501, 4U10, 4022, 4033, 4042, and 6501. 40 COMMON ABBREVATIONS AC Alternating Current AGB Accessory Gearbox A/I Anti-Icing CDP Compressor Discharge Pressure CP3) CIT Compressor Inlet Temperature (T2) ECU Electrical Control Unit HMU Hydromechanical Unit IGV Compressor Inlet Guide Vane LDS Load Demand Spindle LRU Line Replacement Unit Ng Gas Generator Speed (High Speed Rotor) Np Power Turbine Speed (Low Speed Rotor) Nr Helicopter Rotor Speed p3 Compressor Discharge Pressure PAS Power Available Spindle PTO Power Takeoff Assembly Q1 Power Turbine Torque (one engine) Q2 Power Turbine Torque (other engine) T2 Compressor Inlet Temperature T4.5 Power Turbine Inlet Temperature TGT Turbine Gas Temperature VG Variable Geometry 41 ENGINE CHARACTERISTICS MANUFACTURER: MODEL: TYPE OF ENGINE: OUTPUT POWER: TYPE OF COMPRESSOR: VARIABLE GEOMETRY: COMBUSTOR: GAS GENERATOR TURBINE STAGES: POWER TURBINE STAGES: DIRECTION OF ROTATION: ENGINE WEIGHT (DRY): ENGINE LENGTH: MAXIMUM ENGINE DIAMETER: FUEL: FUEL CONSUMPTION: LUBRICATING OIL: TIME BETWEEN OVERHAUL: Aircraft Engine Group, General Electric T700-GE-700 Turboshaft 1,560 shaft horsepower (SHP) at sea level, standard day conditions at 20,900 RPM Np Combined axial/centrifugal consisting of six stages--five axial and one centrifugal Inlet guide vanes, stage 1 and 2 stator vanes Single, annular chamber with axial flow Two (cooled) Two (uncooled) Clockwise, aft looking forward 415 pounds maximum (190 kilograms) 47 inches (120 centimeters) 25 inches (63 centimeters) MIL-T-5624, Grade JP-4 (Jet B) or MIL-T-5624, Grade JP-5 (Jet A) Grade JP-8 1,251 SHP at sea level, and standard day conditions and Np 100% consumption is approximately 587.97 lb/hr (per eng) MIL-L-7808 or MIL-L-23699 None (on-condition overhead only) 42 CROTCH RETAINING BOLT HOLE NO. 1 ENGINE CROTCH RETAINING BOLT HOLE NO.2 ENGINE ~ w A / FORWARD ENGINE ~ SUPPORT TUBE GIMBAL CROTCH ALIGNMENT PIN HOLES FOR NO. 2 ENGINE CROTCH ALIGNMENT PIN HOLES FOR NO. 1 ENGINE u3s-73o ENGINE FORWARD SUPPORT TUBE 6883051 .t';::.-..---........ ... ? ....... f '•'..·. / ... --..... ',', / .......... ','\ t"'••::... .... .. .. .... --................ ...... ,, 11---.. .... .... .. ' '\' f ........:....:~~.... ....... ':\ /----'"' .. ,, \ ,, ............. "'\\:':\ \, \\ j '·\. \ \\\ \ ,, ~:,:.. ... \ \ ,, \ ,, ---.. .......... \ \ \ '\ \ ,, , ... •::..-.:~-.--~:.. ................ \ ' ,,, \ ,, !"'~'''--'-----... ":Jr ):'.. '....... ',~':'... \ \ \\\ \ It (,-"' ...... .. "'.. \'\\ \ \ ''' ' '' ',~\ \ \\', 1 I Ill I Jj ,, \ ,,, I J \\ \ \',• ~ ,"! I·, '-:-, \ ltj I ))} _.,f/ -.~·--l-!__ ,f.,_,J'_., ....ENGINE MOUNT LINKS 1/ ,'..-_-_-... (., /. ,#··!/-'y•:. ~ "" J ..... ~,.....'.. ..~~::.. -.. .. ! ,___ i!u," -~ " <) . " .. 5'. ' ~ , -_-.. ··. ...,.~..... -~-' ~ .,.-....................,......... ',,.. ............. ........ .. ....... .. .... .... ... \ .. , ...... ...."" ~ fr....~\ \ \\\ ','~.t'. ', -.:.:.: ...... ', \ \\ t .. •, f -.......:... ..:..., \ \\'. ., .... "" ,~~...'.. '> ... .. ,. .. ' -- ,, ""' '•, ,, ' ',,.. ""' ' .... I J ';..... \ ,, \ '' '"" ',' .. (• I II I II I I ' \ ) I , \1 I I II I I f··-'•'.-,., •,' ..._______ENGINE MOUNT STRUT •.( )_ II I ~~ I •;-, ' ,,.l'::' .... -II I J ......,.. , ' ,.,<' I~ / I / '•\ .... .....:...... ' I .,:-... "" I j, \'~ .... ..... /" ,'\ ' • ·----, ) ,."' '• ..... ___ .... , / \~ ......... ,"" ...'.,...., ;!/•;::J ENGINE MOUNT SUPPORT ENGINE COMPAR~:ENT DECK/.., U35-456 T700 AFT ENGINE MOUNTS 6883048 NOTES 45 DESCRIPTION The· T700-GE-700 engine is designed and manufactured by the General ElectricCompany in Lynn, Massachusetts, USA. It is a compact, lightweight, turboshaft engine rated at 1,560 horsepower. It has a combination axial/centrifugal compressor, an annular combustor with central fuel injectors, an air-cooledgas generator turbine, and an uncooled power turbine with a coaxial driveshaft extending forward through the gas generator. The forward end of the drive shaft is connected, via a splined joint, to the engine output shaftassembly. The compressor has variable stator vanes and a starting bleedvalve. An integral inlet particle separator protects the engine from foreignparticles. The engine also incorporates an emergency lubrication system, ahistory recorder, bearing accelerometers, erosion indicators, and otherdiagnostic systems. The engine is constructed in four modules, each comprising the following: 0 Cold Section Module: Inlet section, compressor, and midframe. 0 Hot Section Module: Combustion liner; stage 1 nozzle; and gasgenerator, turbine rotor submodules. o Power Turbine Module: Power turbine rotor and stator and exhaustframe assembly. 0 Accessory Section Module: Accessory gearbox and related accessories. 46 GAS GENERATOR TURBINE STATOR FIRST STAGE NOZZLE COMBUSTOR .c'-1 POWER TURBINE MODULE .. ~ COLD SECTION MODULE ENGINE MODULES SWIRL FRAME ~ co POWER TAKEOFF ~ ~ SCROLL CASE FRONT FRAME INLET SECTION DIRT OVERBOARD BLOWER~- ~ ~ 1.0 ------COLLECTION SCROLL CONTAMINATED INLET AIR T700 INLET PARTICLE SEPARATOR AIRFLOW U35-443 67T3094 VARIABLE GEOMETRY AND ANTI-ICING SYSTEMS Compressor Variable Geometry System The variable geometry system of the T700, high-performance compressor permits optimum performance over a wide range of operating conditions. Use of variable stator vane angles facilitates rapid, stall-free accelerations and optimizes fuel consumption at part-power conditions. The variable geometry components include the stator vanes in stage 1 and 2 of the compressor casing, inlet guide vanes (IGV) in the main frame, lever arms attached to the individual vanes, and three actuating rings (one for each stage). The three actuating rings, levers, and vanes are actuated and synchronized by a crankshaft assembly which is positioned by an actuator within the t~dromechanical unit (HMU). This actuator is, in turn, positioned by a servo system with feedback which responds to compressor or gas generator speed (Ng), compressor inlet temperature (T2), and physical position of the variable geometry actua.tor. Variable Geometry System Operation At maximum power, the variable stators are actuated to their farthest open condition to admit the greatest airflow to the engine. At this time, the starting bleed valve is fully closed, so that al l the compressor discharge air is delivered to the combustor and turbine sections. When less than maximum power is required, the compressor speed (Ng) is less than lOU percent and the pumping characteristics of the individual compressor stages is changed. Air pumping capacity is higher in the fon·Jard stages of the axial compressor than the aft stages. To remedy this condition, the T7UU vari ab 1e geometry system acts by closing down the variable stators in the forward portion of the compressor. Similarly, changes in T2 affect the compressor by closing the variable stator vanes with increasing T2 and opening the vanes with decreasing TL. At compressor speeds below M7 percent (30 percent torque), the HMU actuating system also positions the starting bleed valve in the open position. Anti-Icing and Starting Bleed Valve The anti-icing and starting bleed functions are accomplished in a singlecomponent assembly. The starting bleed valve is a modulating valve actuated by a connecting link to the variable geometry cr ankshaft. Starting bleed modulation is thus controlled as a function of Ng and T2. The anti-icing mode is selected with a cockpit switch. The valving assembly opens as a bleed valve at low Ng and closes when Nq is above 87 percent. However, if anti-icing is selected, the valve rema~ns open above that speed. so Vl ,_.. AFT LOOKING FORWARD FORWARD LOOKING AFT COMPRESSOR STATOR ASSEMBLY Cl) w (!) ~ 1- Cl) ...J ~ X ~ > _J m ~ w en en <( a: 0 t0 a: a: 0 en en w a: a.. ::E 0 (.) (!) 2 a: ~ w c:c M 0 2 BO~ESCOPE PORT OIL INLET TUBE '-<+------MID FRAME CASING DIFFUSER CASING COMPRESSOR CENTRIFUGAL .... ., " , ., DIFFUSER -·-.. ·-. " V1 w 4:.•."':-::'~!', / PRIMER NOZZLE PORT (2) "" IGNITER PLUG PORT (2) ENGINE MOUNTING LUGS DIFFUSER AND MIDFRAME ASSEMBLY STAGE 1 NOZZLE COMBUSTION LINER \.11 ~ STAGE 2 ROTOR STAGE 1 ROTOR GAS GENERATOR TURBINE STATOR HOT SECTION MODULE w ...J ::::> 0 0 ~ w z Ill a: ::::> .... a: w ~ 0 Q. ROTOR ASSEMBLY STAGE 4 TURBINE BLADE VI 0\ DRIVE SHAFT STAGE 3 TURBINE BLADE BLADE RETAINER .......__ -~~ -.........._ __ L~.. _, BLADE RETAINER --STAGE 3 TURBINE DISK ~/' STAGE 4 TURBINENOZZLE STAGE 4 TURBINE DISK & POWER TURBINE ROTOR DRIVE SHAFT COMPONENTS e e e • I..J1 -...J 2 3 4 5 6 MAIN BEARINGS AND SHAFTS RADIAL DRIVE SHAFT COVER [FRONT a (4) V1 00 PARTICLE BLOWER IREAR VIEW I OIL PRESSURE TAP COLD OIL RELIEF VALVE BYPASS ::::::::::::::1-FU EL FILTER FUEL SCAVENGE SCREENS (6) SEQUENCE VALVE f.;:::'J-VG ACTUATOR LINK ACCESSORY SECTION MODULE -~ ENGINE START SYSTEM The engine start system is a pheumatic system using compressed air ducted to - air turbine starters for engine starting. Compressed air is obtained either from (1) the APU, (2) engine crossbleed, or (3) an external air supply. The I APU has a check valve to control APU bleed-air flow to the starter of either engine. The engine crossbleed has a bleed-air manifold and a combination I crossbleed shutoff and check valve for each engine to permit starting the I opposite engine. The external air source supplies air through a combination external connector and check valve to either engine. Electrical power for the No. 1 engine start system is obtained from the DC essential bus through the No. 1 ENG START circuit breaker on the DC essential bus circuit breaker panel. Electrical power for the No. 2 engine start system is obtained from the No. 2 DC primary bus through the NO. 2 ENG START CONTR circuit breaker on the pilot's circuit breaker panel. An ENGINE IGNITION switch, marked ON and OFF, is at the center of the main instrument panel. When it is ON, and when an engine start button is pressed, the engine ignition system operates. The capacitor discharge engine ignition system is a noncontinuous AC powered, low-voltage system. (Ignition is automatically shut off when Ng reaches 52% to 65%.) The system consists of an ignition exciter, two igniter plugs, ignition leads, switches, relays, and advisory lights. Electrical power for the ignition system is obtained from the engine-mounted alternator. NOTES: 59 HEATER INLET 0 0\ ANTI-ICE VALVE---STARTER CROSSBLEED START VALVE CONTROL VALVE PNEUMATIC SYSTEM LINES DIAGRAM U35-150 67T3102 e .a · e . [A) I PACKINGS ENGINE ~ ~ ~ PNEUMATIC STARTER \\ ~ · TUBE ~~ / , ~ 0\ ~ ~~ ~ ~ ' ~ ~ ....... 01 ~ ~ 'I --~' ~ip.,..-------\) ---coNNECTOR ~"'--..ELECTRICAL PLUG U35-420 T-700 STARTER 67T3115 CROSSBLEED VALVE 1 ~ lSTART MANUAL 1L;..;AK vAut. ![ ~To No. 2 ENG SWITCH OFF OVERRIDE -KI.QI SWITCH 0\ N ~~-SPEED SWITCH DC / I ABORT SWITCH ESSEN BUS IENG I I K24 11#1 ENGINE STARTERIH /IJ-~IoFFI -- ~ 1 DC / IAPU I,.. PRI BUS I I I I - APU CONTROL UNIT START BYPASS VALVE NO. 1 ENGINE START SYSTEM U35-145 SCHEMATIC 67T3116 e e ---· e e • TO AIR SOURCE CROSSBLEED HEAT/START ~ VALVE SWITCH . J ~===--=--=-.~r--t-______~ MANUAL START OVERRIDE SWITCH SWITCH 0\ NO. 2 l.V ENG START CONTR I 1 SPEED SWITCH jl.-o-L·---,-J TO APU START_ ABORT NO I • I [#2 ENGINE STARTER I 2 BYPASS VALVE SWITCH PRI. DC BUS FROM APU -:....,__----t 1 .. 1 FROM EXTERNAL START CONNECTION NO.2 ENGINE START SYSTEM U35-146 SCHEMATIC 67T311 7 LUBRICATION AND SUMP SEALING SYSTEMS Lubrication The lube system is a self-contained, pressurized, recirculating, dry sump continued bearing operation after loss of oil pressure from any has been system. It consists of the following subsystems and components: 0 Oil supply and scavenge pump. 0 Seal pressurization and venting. 0 Emergency lube system. 0 Oil filter and cundition monitoring. 0 Tank and air-oil cooler. 0 Fuel-oil cooler. Lube System The oil tank is integral in the main frame. An emergency system for providing cause built into the design. Small integral oil reservoirs, located in each sump, are kept full during normal operation by the oil pressure pump. Oil is always bleeding out of these reservoirs at a slow rate. Air jets, also continuous, act as "foggers•• for this bleed and provide oil mist lubrication at all times. This continues for at least 1 minute, even if the oil supply fails, and requires no triggering device. Each scavenge line has a scavenge screen for fault isolation, and a single chip detector is provided for cockpit warning. Oil Tank The oil tank, fuel-oil cooler, and main frame are an integral aluminum casting. The volume of oil required to fill an empty tank is 1.7 gallons. The tank is filled using the 3-inch, gravity-fill port on the right-hand side. Oil supply to the oil pump is via a screen which is removable through the forward tank wall. This also serves as a tank drain and cleanout port. Visual indication of oil level is supplied by a window on each side of the tank which is used by the ground crew for servicing. Lube and Scavenge Pump The lube and scavenge pump is~ seven-element, gerotor-type pump containing one supply element and six scavenge elements. The pump does not require an integral antileak check valve because the pump is located on the top of the engine precluding a gravity head (pressure) on the pump. 64 Pump shaft bearings are placed to isolate the supply and B-sump scavenge high-pressure pump elements from the other scavenge elements. Oil scavenge elements discharge at the top where air is readily cleared and oil wetting of the pump by another assures priming. Three bolts hold the oil pump cartridge in place in the accessory gearbox. Oil Filter System Oil discharged from the supply pump is routed through a cored passage to the oil filter which is a 3-micron, disposable element. Associated with the filter are a bypass valve which opens at 80 psi minimum, an impending bypass indicator which displays a popout button at 44 to 60 psi, and an electrical bypass switch, actuating at 60 to 80 psi, to provide a signal to a warning light in the cockpit. During cold weather starting, the high oil pressure will cause the bypass valve to open and the bypass switch to actuate. However, the impending bypass indicator contains a thermal lockout to prevent it from tripping. As the oil warms up to over 380 C (1000 F), the thermal lockout is disengaged, and the indicator is ready to warn of filter contamination. A cold-starting relief valve, downstream of the filter, protects the system by opening at 120 to 180 psi and venting the excess oil to the gearbox case. Oil Supply System Oil leaves the pump, is filtered, and proceeds through cored passages in the accessory gearbox where the flow divides to supply oil to the A, B, and C sumps and the accessory gearbox. A plug-in connector brings oil to the A sump through the vane at 3 o'clock and a plug-in connector between the strut and sump outer wall. The A-sump supply and scavenge are accomplished entirely with internal lines. Oil supply to the B and C sumps leaves the aft side of the gearbox in an external tube which tees to the two sump inlets. Oil supply to the B sump enters through a double-walled and insulated tube assembly in the 9 o'clock strut. Within the midframe and sump assembly, there are no joints or 0-ring seals in the lines. This eliminates the potential leaks and 0-ring deterioration in hot zones. A check valve shuts off flow to the B sump to prevent flooding, when the engine is shut down. Oil supply to the C sump enters through a plug-in service tube through the exhaust frame 7:30 o'clock strut. The 0-ring seals at the hub end are immersed in the cool sump environment for long life. Emergency lubrication for each main beari ng is provided in the event of a loss of oil pressure. Oil reservoirs in A and B sumps will adequately lubricate the bearings for 1 minute at 75 percent power. Fourth-stage air pressurizes the emergency air-oil mist jets. 65 Scaver1'ge System Six scavenge pump elements return oil from the sumps to the oil tank. The A sump uses two scavenge elements to ensure scavengi ng at attitudes up to 40 degrees, nose up or nose down. The C sump uses three scavenge elements. The B sump has one scavenge path, because the high speed of the gas generat( r rotor and t he sump configuration force oil into the scavenge line. \ I•The six screens are individually labeled for fault isolation. The common discharge from the scavenge pump is routed to a single chip detector which may be wired to a cockpit warning device. Oil Cooling Two separat e heat exchangers cool scavenge oil bef ore it returns to the tank. Oil from the chip detector passes through the fuel-oil cooler--a tube-and shell device. Thi s cooler is mounted on the front of the AGB and is a line replaceable unit. After passing through the fuel-oil cooler, oil enters the top of the main frame where it is ducted through the hollow separator scroll vanes. This heats the vanes for full-time anti-icing. The vanes then discharge oil into the oil tank. If either of the oil coolers become clogged, a cooler relief valve will opento dump scavenge directly into the oil tank. • Valve cracking pressure is 22 to 28 psi. Seal Pressurization Seal pressurization provides the following: 0 Prevents oil loss from sumps by controlled air inflow. 0 Prevents hot gases, dust, and moisture from entering sumps. 0 Pressurizes the emergenr.y air-oil mist jets. Compressor stage 4 air is used for the seal and power turbine balance pistonpressurization. To obtain the cleanest air for pressurization of the A and B sump seals, stage 4 air is bled from the rotor. This method of bleed airextraction has the advantage of eliminating external piping. Bleed air entersthe spool through curvic coupling teeth, aft of the stage 4 blisk, and splitsto flow forward and aft. 66 The forward flow is to the A sump aft labyrinth seals. It enters the interseal cavity through the stage 1 blisk front holes; the p~essure is 3 to 5 psi above the sump pressure. Part of the flow enters the sump via the oil seal, some returns to the compressor stage 1 inlet hub, and a small flow from the interseal cavity is used to pressurize the No. 1 carbon seal and the emergency air-oil mist jets. B-sump seals are also supplied from stage 4 air which flows aft into the B-sump forward interseal cavity via holes in the compressor aft shaft. From the forward interseal space, air flows to the aft interseal space through an annular passage which blankets and cools the B sump. Air in each interseal space divides, and part of it flows into the B sump to prevent oil loss. The remainder flows away from the sump to buffer hot leakage air and exclude it from entering the sump. Hot leakage air from the compressor discharge seal is conducted out of the midframe via the 5 o'clock strut and then aft to the exhaust frame. Inner balance piston leakage flow passes through holes in the gas generator turbine forward shaft and flows aft under the turbine wheels to provide cooling. This flow reenters the flow path at the root of the stage 3 turbine blades. A tube at the bottom of the seal pressurizing cavity drains any oil seepage out through the 5 o'clock strut. Air to the power turbine balance piston is brought in through an external pipe from the compressor casing stage 4 bleed port (5 o'clock) to the exhaust frame 4:30 o'clock strut. Pressure level at the No. 5 seal is set by a fixed-air inlet restrictor and provides a forward force on the rotor area inside the balance piston seal. This force reduces the No. 6 bearing thrust load. Sump Vents Venting of air from the sumps is radially inward through holes in the shafts. The center vent concept has proven successful on many GE engines and is an effective means of separating oil from the vent air using the centrifugal field produced by the high-speed shafts. A-sump air flows aft into the annulus, between gas generator and power turbine shafts, through holes in the power turbine shaft and then forward to holes in the torque reference shaft. This path captures oil vapor and pumps it centrifugally into the sump. B-sump, center-venting flow passes inward through a perforated disk, which is a part of the forward rotating seal, and then aft to larger holes in the compressor aft shaft where it flows into the intershaft space. This flow path provides a low air pressure drop but a high oil pressure centrifugal field. Vent air from the B sump flows forward in the intershaft space to pressurize the intershaft seal at the A sump. Surplus B-sump vent flow moves aft to join the inner balance piston leakage flow. C-sump vent airflow passes inside the aft end of the power turbine shaft and into a standpipe connected to the aft sump cover. A small dam in the torque reference shaft traps any liquid oil remaining in the vent air and returns this to the C sump through small weep holes. 67 The oil tank vents through a connector into the accessory gearbox and from there down the radial drive shaft tube into the A-sump center vent. Pressures throughout the seal and sump cavities are a function of the engine cycle. System balance is maintained throughout the ~light envelope; no valves are required. 68 e e e \.0 "' LUBRICATION SYSTEM SCHEMATIC CONNECTOR !scAVENGE SCREENS ) C SUMP C SUMP AFTCOVER ICHIP DETECT(fR] CAPTIVE BOLT -...J 0 At) PREFORMED ~~ B ~~--~A~::__ -~~--~' e • SCAVENGE SCREENS AND CHIP DETECTOR Figure 3-5 e ~g~'1 ~l~ ~ j ~ 1:1~~~ ~~ ~ ~ ~ 1-. p \.. ).. X ' ;u p ~ '1 1\ ):::r- 1-J I-Ii IBi h lr i'v rJ ~ """ ~ -. !\.. p .. H 'J ~ 1 ::> 9 ..J 0 I ..J L. ~ !~P ... \ ... w < Ill ...I ~ h ~~~.., ~~~~~~ 0 ~~L) I~ w Ill 1 ::> ..J ..J ..J w 1 w :J: :;:) VERSPEE I a -swiTCH soLENOID 2 _j I a 1 SIGNAL TO OTHER ENGINE SEQUENCE VALVE ELECTRICAL CONTROL UNIT L.::.: ___ ELECTRICAL CONTROL UNIT SCHEMA TIC Figure 4-8 TORQUE SENSOR TUBE (REFERENCE SHAFT) OUTPUT SHAFT SPLINE OUTSHAFT (Np) TWIST co BRING Np TEETH CLOSER TO REF TEETH (See 4-10) N POWER TURBINE DRIVE SHAFT DISK MOUNTING FLANGE TORQUE SENSOR REFERENCE TEETH POWER TURBINE SENSORS e e ~---__e I I I __ J , ( .... UJ :J 0 cr: 0 1 0 cr: UJ N UJ :J 0 cr: 0 1 ~ :J ~ X <( ~ ,.. -_j ( .... z 0 1 <2:: 0: w a... 0 0: 0 C/) z w C/) w :::J a 0: 0 1 83 NO. 1 ENG OVSP :: t:i'Rf:Di:f.mt::: NO. 2 ENG OVSP TEST A TEST 8 ;:;:;:;: ;:;-t;;~:;:;:;:;:: TEST A TEST 8 &~ttau.f NO.2 ENG OVERSPD SWITCHES UPPER CONSOLE PANEL TEST A TEST B AIRFRAME N0.2AC. ~ l 1 NGPRI BUS ·.---E 0 VSP ," r-------, TO ENGINE (ECU) I TORQUE .p.. .I (X) HISTORY AND RECORDER ~ l ·l OVERSPEEDOVERSPEED OVERSPEED TO I 400Hz SENSOR I SENSING SENSING TORQUE POWER I CIRCUIT CIRCUIT CIRCUITS SUPPLY I A B I L PRIORITY~_j_0 --l I 0 I I CIRCUIT I SEQUENCE i VALVE I ~VERSPEEC ALTERNATOR I SOLENOID POWER L _____ -.J SUPPLY NO.2 ENGINE ENGINE OVERSPEED SYSTEM U35-364 DETAILED BLOCK DIAGRAM AF3031 • e -- FUEL SYSTEM Genera1 The fuel system is designed to provide the proper fuel flow to the engine under all operating conditions including starting, idle, normal flight, and maximum power. In addition, the fuel system operates in conjunction with trre electrical system to provide automatic trim of engine power. Components of the fuel system include an engine-driven boost pump, fuel filter, hydromechanical unit, sequence valve, and fuel injectors and primer nozzles in the combustor. Boost Pump The boost pump is a vane-centrifugal type with an ejector or jet pump at the inlet. It is capable of providing suction to draw fuel uphill from unpressurized tanks. This decreases the fire hazard in case of a damaged fuel line. The pump is mounted on the front face of the AGB and delivers fuel through a cored passage to the fuel filter. Pump discharge pressure is 45 to 90 psi minimum at full power. Fuel Filter -I The fuel filter is a 30-micron, high-capacity filter with bypass valve and bypass warning devices. It is mounted on the forward, left-hand side of the AGB. The filter element is housed in a bowl which is easily removed for element cleaning or replacement. The bypass valve opens at 7.5 psi minimum to assure continued fuel flow with a blocked filter. At the same time the valve opens, an electrical switch is closed to provide a warning signal. In addition, an impending bypass warning is displayed at 6 psi in the form of a popout button on the filter housing. Hydromechanical Unit (HMU) The HMU, mounted on the aft center of the AGB, receives filtered fuel from the gearbox through a cored passage. The HMU contains a vane-type, high-pressure pump to provide pressure to deliver fuel into the combustion system. Some fuel is tapped off to operate various servos in the HMU for the following: 1. Position a metering valve to assure proper flow to the combustor. 2. Position a servo piston which actuates the variable geometry and starting bleed components. 3. Sense various parameters (T2, P3, Ng) which influence fuel flow and variable geometry position. 85 The HMU responds to two separate linkages from the airframe. One linkage is connected directly to the collective pitch mechanism to sense load demand; this input is called load demand spindle (LOS). The pilot actuates the LOS simultaneously when he selects a collective pitch angle. The other input linkage is a separate cockpit control called the power available spindle (PAS). This permits the pilot to manually trim the engineoperation. The PAS also permits the pilot to throttl e the engine to ground idle to minimize noise and fuel consumption and to shut off fuel flow to shut down the engine. A third input is in the form of an electrical signal from the electrical control unit (ECU) which actuates a torque motor in the HMU and automatically trims engine power. The HMU also responds to sensed engineoperating parameters such as- 0 T2 (compressor inlet temperature) vi a a probe in the main frame. o P3 (compressor discharge pressure) via a flexible hose. The HMU also positions the variable geometry linkage by a hydraulic piston extending from the left side of the HMU casing. Sequence Valve The fuel metered by the HMU is delivered to the oil cooler by an external hose and then to the sequence valve. The sequence valve has the following four functions: 1. Provides both primer fuel (to the 2 primer nozzles) and main fuel (to the 12 fuel injectors) during engine starting. 2. Shuts off primer fuel after lightoff and purges the primer nozzles of fuel by blowing compressor discharge air through them. 3. Drains the main manifold when the engine is stopcocked. 4. Cuts back fuel flow to the engine when the ECU actuates the overspeed solenoid. This is accomplished by bypassing a portion of the HMU output back to the inlet of the HMU. 86 e -e FILT CX> '--J SEQUEN OVERSPE SOLENOI PRIME U35-442 COOLER FUEL IN --BOOST PUMP ER-........... n ACCESSORY GEARBox ~ ~~ t t ~ t MAIN MANIFOLD :E VALVE~ --VANE 'D ~ ·~ ) ~ TOI MOTOR I~ P3 t HVDI HAN ICAL RMANIFOLD.._.._ ~ ~ __. T ,......, ,......, r-""1 ,......, ......---, ,......., r-""1 r-'1 ....., ,...., ,...., I "' 1--- T700 FUEL SYSTEM SCHEMATIC 61F3025 Hydromechanical Unit Functions • Fuel Pumping • Fuel Flow Metering • Collective Pitch Compensation Through LOS • Acceleration and Deceleration Flow Limiting (I ncl ud i ng Starting) 00 00 • Ng Limiting • Variable Geometry Positioning • Torque Motor to Trim Ng Governor • PAS Override and Control with Electrical Unit Inoperative • Vapor Vent on PAS Overtravel for Fuel System Priming T700.469 13-761 e e e 00 1.0 P53 SIGNAL INPUT ~---VG ACTUATOR HYDROMECHANICAL UNIT 1.0 0 SEQUENCE VALVE FUNCTIONS e SEQUENCES FUEL TO START FUEL MANIFOLD AND 2 PRIMER NOZZELS. e SEQUENCES FUEL TO MAIN FUEL MANIFOLD AND 12 I,AIN FUEL INJECTORS. e PROVIDES P-3 AIR FOR PURGING START AND MAIN FUEL MANIFOLDS AFTER FUEL HAS BEEN SHUT OFF TO PREVENT COKfNG. e PROVIDES OVERS PEED PROTECTION AT 106 !l% N.P. e e ENGINE ANTI-ICING SYSTEM Engine anti-icing is a combination of hot axial compressor discharge air and neat rejection from the air-oil cooler. Hot air anti-icing is controlled by a solenoid operated air valve. When electrical power is applied to the solenoid, anit-icing is turned off. When the system is turned on, power is removed, the valve spring loads to open and illuminates advisory capsules indicating #1 ENG ANTI ICE ON or #2 ENG ANTI ICE ON. Axial compressor discharge air is bled from stage 5 of the compressor casing, routed through the antiicing/bleed valve, and delivered to the front frame through ducting. Within the swirl frame, hot air is ducted around the outer casing to each swirl vane. Front frame anti-icing flows through a passage in the main frame, then exits to tne main frame scroll and is discharged with inlet particle separator air. Anti-icing is also ducted to the compressor inlet guide vanes (IGVs). Hot scavenge oil passing within the scroll vanes in the main frame prevents ice buildup. Deswirl vanes in the front frame need no anti-icing because of the particle separator action. Engine inlet anti-ice also uses stage 5 engine oleed air. The inlet anti-ice valve (not engine furnished) is opened with the same control switch that deenergizes the engine furnished anti-ice valve and illuminates two advisory lights indicating #1 ENG INLET ANTI ICE ON and #2 ENG INLCT ANTI ICE ON. H.QI£S: 91 Anti-Icing System Anti-icing is accomplished by a combination of hot axial compressor dischargeair and heat rejection from the air/oil cooler integral to ·the main frame. The hot air anti-icing is a system controlled by an external electrical signal which triggers a solenoid-operated air valve. When electrical power is applied to the valve solenoid, anti-icing is turned off. With power interrupted, the valve reverts to the anti-icing mode. Axial compressor discharge air (station 2.5) is bled from the compressor casing at the 7 o'clock position, routed through the anti-icing/bleed valve, and delivered to the front frame and swirl frame via ducting. Within the swirl frame, hot air is ducted around the outer casing to each swirl vane. The hot air is circulated within each vane by a seri es of baffles and then exits from two areas. Approximately 90 percent of this hot air exits at the vane outer trailing edges. The other 10 percent exi ts through a series of circumferential slots in the swirl frame hub at the aft edge. This arrangement also acts as ·a "rain step'' to preclude water from adhering to the hub and flowing into the compressor. Front anti-icing flows through a cored passage in the main frame to the front frame splitter lip; then, it exits to the main frame scroll and is discharged with inlet particle separator air. Hot scavenge oil, passing within the scroll vanes in the main frame, precludesice buildup which could result from moisture-laden inlet particle separator air. 92 MAINFRAME ACTUATOR SHAFT HVDROMECHANICAL UNIT ~ w \.0 ACTUATING RING I ANTI-ICING DUCT ___,}_COMPRESSOR CASING STARTING BLEED VALVE PUSH ROD CRANKSHAFT ANTI-ICING AND STARTING BLEED VALVE T700 STARTING BLEED VALVE INSTALLATION U35-463 67T3103 ~~----------------------~------------------~ NO.1 ENG ANTI ICE WARN N0.1 DC PRI OFF BUS NO.1 ENG #1 ENG ANTI ICE ..-1-----o-'"'" ANTI ICE ON ~ oON ENG ANTI-ICE TO HMU N0.1 \ \ 0 .p- DC hE.~ o ~ANTI ICE ESNTL NO. 1 ENG BUS START K45 (ENTt ~ #1 ENG INLET ANTI-ICE VALVE t NOTE: ~ ~ ~ ~ ALL RELAYS SHOWN ENERGIZED. ANTI ICE/START BLEED VALVE NO. 1 ENGINE ANTI ICE SYSTEM U35·353 SCHEMATIC -ENGINE START 68F3032 #1 ENG INLET ANTI-ICE ON STAGE 5 BLEED AIR ~ ~ ~ NO.1 ENG ANTI-ICE WARN N0.1 DC PRI BUS #1 ENG #1 ENG INLET ANTI-ICE ON ANTI-ICE ON oON ENG ANTI ICE TO HMU N0.1 . -\ 1.0 DC L_~ V1 ESNTL~1 ENGf I ENERGIZED • , BUS START ~ ~ , -'1h K45-- ENERGIZED STAGE 5 BLEED AIR #1 ENG INLET ANTI-ICE VALVE ANTI ICE/START BLEED VALVE NO. 1 ENGINE ANTI-ICE SCHEMATIC U35-354 ANTI-ICE OFF MODE 61F3033 ~~----------------------~------------------~ NO.1 ENG ANTI ICE WARN NO.1 DC PRI OFF BUS NO.1 ENG #1 ENG #1 ENG INLET ANTI ICE I I ()._ ANTI-ICE ON ANTI-ICE ON ~ ON ENG ANTI ICE TOHMU STAGE 5 NO.1 ~ANTI CT\ BLEED AIR DC NO. 1 ENG BUS START K45- \0 ESNTL ~g~~ ICE t (ENT l ~ #1 ENG INLET ANTI-ICE VALVE l ~ ~ ~ ~ ~ ANTI ICE/START BLEED VALV-E NO. 1 ENGINE ANTI ICE SYSTEM SCHEMATIC -ANTI ICE ON MODE U35-355 llf3034 e ·-· ENGINE _______.., ', AIR IN LET ---------- ------1" ~ ,•' ,•' -~ \ / ,•' -----. '\ ,' .·:----\ ', ELECTRICAL PLUG ./ . ~-' \ N. .,' ,· \ . .... .· .,. .... "'\ ' ,/ ' :' /-.-'/ : ' ... .• ~ •, "'-' , ,. \ \I I '' ~ / ' ' \ \ •'I .. 1.0 / '"' \\ \ , \ \~ -·~ ......, . ·'I '.•,• '.. \ \'\\ ' , t I '' l , • • ' \ . I , l l , '-• ' . . ' / .. .-·· ' . ' ' . . -.... I / • 1/ / ~ : ., : ,.. / ~'&. 1 / ~ I ,<·· w ~ \.....-----·-".....:~------·· l~ \ CLAMP CLAMP ANTI-ICING VALVE 7 COVER T700 AIR INLET ANTI-ICE VALVE LOCATION U35-331 67T31 04 ENGINE CONTROL SYSTEM Power Control System Push-pull cables connect the power control levers in the cockpit to the poweravailable spindles (PAS) on each engine hydromechanical unit (HMU). Each power control lever has four detent positions; OFF, IDLE, FLY, and LOCKOUT. The power lever is advanced to FLY for flight. The PAS setting represents the highes~ power that could be supplied if demanded. Powerturbine speed is not governed until the power lever is advanced from IDLE. Load Demand System The load demand spindle (LDS) input is a function of the collective pitch. It provides compensation to reduce transient droop of N . The spindle inputsload demand signals directly into the hydromechanica~ unit. A reduction in collective pitch decreases LDS position, reducing fuel flow and giving immediate and accurate gas generator response. The new setting is trimmed by the ECU to satisfy the Np/Nr and load control requirement set by the electrical control unit tECU). Engine Speed Trim Control System Another input required to fully control the engine is the Np/Nr speed reference signal obtained from the engine speed trim system. This signal is received by the engine ECU. The speed control system allows Np/Nr adjustmentsbetween 96.4 and 100.5 percent. Speed can be increased or decreased by the ENG RPM (beeper trim) switches on the pilot's and copilot's collective grips.The pilot can override the copilot's control. Engine Overspeed Protection System An overspeed protection system protects the power turbine from destructive overspeeds. If a malfunction should cause N to reach 106 percent, the electrical overspeed system will automatical~y decrease fuel flow to the engine as necessary to prevent N /Nr from exceeding 106 percent. Two test switches, in the cockpit, allow ~he pilot to test the overspeed system before takeoff. 98 \ 1 ,\ I ~<-:~ .. ,,, ,....,, '' ,, \ \ \ ,,,.." ,,, ' ' \ ' ,, \ /', \~ I I " \ I 't\ \ ' ', ," ·'\I ,_ . Cl) .J 0 It: 1 z 0 u Ill z - z " Ill I I ~ (.) - ~ U) w > - t; w _, _, . \ 0 (.) en co ~ II) C'l) ::) 99 Np GOVERNING Np GOVERNING NO.1 ENGINE ENGINE NO.2 ENGINE ECU Np REFERENCE ECU REFERENCE SIGNAL --------REFERENCE SIGNAL INCREASE ENG ...... 0 NO. 2 DC. ~SPEED TRIM 0 PRIBUS ~~---r------------------------------------~ 28 VDC PILOT'S COLLECTIVE STICK GRIP COPILOT'S COLLECTIVE STICK GRIP LOG. LT. EXT. LDG. lT. EXT ON@) -PUSH 0 Off ON@) RET ·PUSH 0 OFf RET ENGINE SPEED TRIM SYSTEM U35-224 BLOCK DIAGRAM 67T3109 __. e e ENGINE CONTROL QUADRANT The control quadrant, centered on the upper console, permits either the pilot or copilot to select engine speed, stopcock fuel, start engine, abort start, and control engine fire extinguisher. The ENG POWER CONT lever positions are marked NUMBER 1 ENGINE and NUMBER 2 ENGINE, and identify the OFF, IDLE, FLY and LOCKOUT positions. They are connected mechanically to each engine's HMU and are used to govern engine speeds. The HMU starts to open whenever the power control lever is advanced more the 2 degrees from OFF and increases proportionately with engine speed to FLY. The engine start switch button is in the power control lever handle. The abort switch, mounted on the control lever, is opened when the lever is pulled straight down. The fuel selector levers marked NO 1 ENG FUEL SYS and NO 2 ENG FUEL SYS allow pilot or copilot to select OFF, DIR (direct), or XFD (crossfeed) position for engine fuel supply. T-handles on the outboard side of either fuel selector lever are used to direct the flow of the fire extinguishing agent to either engine compartment. Push-pull cables connect both the power control levers and the fuel selector levers to their components. NOTES: 101 NOTES 102 DEPARTMENT OF AVIATION SUBJECTS SYSTEMS TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama · May 1984 File No. 47-4743-7 LESSON EVALUATION POWER PLANTS AND RELATED SYSTEMS Complete the following questions by filling in the blanks. 1. TGT must be obtained within seconds during engine start. 2. The maximum temp of ooc TGT may be obtained before reaching idle speed. 3. The engine may be operated a maximum of seconds with NO oil pressure during engine start. 4. The engine may be operated a maximum of seconds with NO percent RPM 1 and 2 during engine start. 5. The ENG OUT warning light should be out at percent NG speed. 6. At ___ percent NG, the ENGINE STARTER caution light should be OFF. 7. Percent RPM 1 and 2 ranges of to percent and to ___ ___ percent should be AVOIDED during engine starting. 8. Engine TRQ must be matched within ___ percent at 100 percent RPMR. 9. During engine overspeed system check, the pilot should press test button .,..,-,::---' NG speed should , press test button ____ NG speed should Press both test buttons _ _ _ and , NG speed should _______ 10. At percent TRQ, the NO. 1 and NO. 2 ENG ANTI-ICE ON advisory light will go OFF. 11. Two functions of the engine anti-ice start bleed valve is to provide engine and to allow the engine to accelerate more through the range. 12. The INLET ANTI-ICE advisory light will only illuminate when the engine ANTI-ICE switches are in the position and the engine inlet cowling temperature reaches oc. 103 13. The engine alternator supplies power to the , the ignition _____, and a signal to the SPEED tachometer. 14. The ECU will attempt to keep TRQ 1 and 2 matched within ---percent. 15. The TGT limiter located in the ECU provides protection at ____°C. 16. Engine torque of percent to---· percent is the 10-second transient of ENGINE RPM 1 and 2. 17. Percent RPM of percent to ---percent is the maximum 12-second transient of ENGINE RPM 1 and 2. 18. The TGT limitation of uc is the maximum for engine start beforereaching idle speed. 19. The maximum time of seconds from NG speed indication to idle at FAT above zouc. - 20. Write the pilot action when the NO. 1 ENGINE STARTER light remains onafter 65 percent NG speed. 21. Write the pilot action if engine TGT exceeds 85o uc during start. 22. Write the pilot action for NO oil pressure indication during engine operation. 23. Write the pilot action when the percent RPMR and NO. 1 engine RPM - increases to 106 ~1 percent during flight. 24. Write the CAUTION to be observed when moving the engine power control1ever to ECU 1ockout. 25. Write the pilot action when the NO. 1 engine indicates no percent RPM, nopercent NG, no percent TRQ, TGT is 860 °C and increasing; NO. 2 engine indicates 102 percent RPM, 85 percent NG, 20 percent TRQ, and 560 °C TGT. RPMR is 106 percent. 26. Write the pilot action for a crossbleed start of the NO. 2 engine usingthe NO. 1 engine as the air source. 27. Write the pilot action for a compressor stall of NO. 1 engine in flight. e 104 POI FILE: 47-4747-9 FLIGHT CONTROLS AND HYDRAULIC SYSTEMS TYPE OF INSTRUCTION: PEACETIME HOURS MOBILIZATION c PE 7.0 2.0 7.0 2.0 LEARNING ORJECTIVES: 1. Given six written questions containing descriptive situations or operating conditions, from memory, the student will write IAW TM 55-1520-237-10 the- a. Purpose of the flight control system mixing unit. b. Hydraulic systems that normally supply hydraulic pressure for operation of the pilot assist servo system. c. Purpose of the tail rotor servo control switch. d. Systems pressurized by the No. 1 hydraulic switch. e. CAUTION to be observed when AC electrical power is applied to the helicopter and the backup pump power circuit breaker is out. f. Five caution/advisory lights that will automatically bring on the backup pump in flight. g. Two hydraulic components that are only pressurized by the backup pump. h. Purpose of the hydraulic leak detection/isolation system. i. Purpose of the hydraulic reservoir fill system. j. Feature that prevents the first and second stage primary servos from being turned OFF at the same time. k. Position of the backup pump switch above 30 percent RPMR. 1. Purpose of the pitch boost servo. STANDARD: Six of six questions must be answered correctly to satisfactorily complete this objective. 2. Given three written questions containing pertinent indications and/or descriptive situations or operating conditions, from memory, the student will write IAW TM 55-1520-237-10 the- 105 a. Cockpit indication of a (single) broken tail rotor cable. b. Cockpit indication when a hydraulic leak occurs in the first stagetail rotor servo system. c. Cockpit indication when a hydraul i c leak occurs in the pilot assist servo sys t em. d. Switch that permits the BACKUP PUMP to automat ically come on in f l ight regardless of backup pump switch position. e. Reaction time of the BACKUP PUMP with one or both main AC g~~nerators. STANDARD: Three of three questions must he answered correctly to satis factorily complete this objective. 3. Give n two written questions containing pertinent indications and/ordescriptive situations or operating conditions, from memory, the student will write lAW TM 55-1520-237-10 the- a. Bank angle limitation when a PRIMARY SERVO PRESSURE caution lighti11 umi nates. b. Airspeed 1i mit at ion when one hydraulic system is inoperative. c. Airspeed limitation when two hydraulic systems are inoperative(VMC). d. Airspeed limitation when two hydraulic systems are i nope rat i ve (IMC). STANDARD: Two of two questions must be answered correctly to satisfactorily complete this objective. 4. Give n two written questions containing pertinent indications and/ordescriptive situations or operating conditions, from memory, the student will write lAW TM 55-1520-237-10 the pilot action for- a. Initial approach and touchdown speed for complete loss of tail rotor control (loss of tail rotor servo pressure}. b. Eme·rgency procedure for No. 1 RSVR LOW, No. 1 HYD PUMP, No. 1 PRI SERVO PRESS (servo control switch first stage}, BACKUP PUMP RSVR LOW caution lights ON, then the No. 2 PRI SERVO PRESS caution light illuminates. STANDARD: Two of two questions must be answered correctly to satisfactorily complete this objective. LESSON RE FERENCE: TM 55-1520-237-10. 106 TASK: This objective supports task: 03-1402-10-1501, 1502, 3501, 4005, 4006, 4027, and 6501. EQUIPMENT: UH-60 composite trainer; UH-60 power train system DVC-118; UH-60 hydraulic systems trainer DVC 1-119; UH-60 caution/advisory panel FR 1330. INSTRUCTIONAL ELEMENT: OOAS, STD, ASB. 107 FLIGHT CONTROL SYSTEM Flight Controls (Mechanical) Mechanical flight controls are used to change the pitch in individual blades to control the attitude of the helicopter. In some of the older generationhelicopters, several different movements were needed to control, for example,the hover attitude. In the UH-60A, these combined movements are automaticallydone by the flight control mixer unit. Mixing can he observed when the pilot pulls up collective for a hover. All three primary servos will move in the same direction to increase pitch in all blades and lift the helicopter, but they do not move the same amount or at the same rate. Mechanical mixing of the flight controls reduces the pilot workload and provides a smoother transition between flight modes. Besides the pilot and copilot, flightcontrol inputs are made by the trim and SAS systems. The flight control system is a hydraulically power boosted, mechanical control system. The conventionally arranged pilot and copilot cockpit controls provide cyclic, collective, and directional control of the aircraft from either pilot station. Pilot and copilot control inputs are combined at the forward cabin overhead and then routed through pilot assist servos (boost,trim, SAS) to the mechanical mixer which couples the pilots 1 inputs to provideaircraft response. Mixer outputs are routed to t he three redundant main rotor servos and to the dual stage tail rotor servo. The main rotor servos are mounted on the upper deck just forward of the main transmission. The servos are connected to the stationary swashplate bymechanical linkages. The tail rotor servo is mounted inside the tail rotor gearbox where it attaches to the pitch change shaft. Mechanical linkages are used throughout the system except for the directional controls between the mixer and tail rotor servo which is a cable system. The entire flight control system is designed to be ballistically tolerant to a 7.62mm API threat and includes the following design features: Redundant cockpit controls,ballistically tolerant pushrods and pivots, ballistically tolerant servos, and redundant directional control quadrant. In addition, normal hydraulic flightcontrol power is provided to each servo stage from two independent sources, each of which is automatically backed up by an emergency pump in the event of hydraulic malfunction. 108 COllECTIV£ PITCH SAS BOOST ACTUATOR ASSliiBLY ----YAWSAS ACTUATOR A YAW BOOST ASSEMBLY TAIL GEAII BOX l (!) I ~\:•?x· · ···· ) ' YAW Tltlll ACTUATOR ASSEMBLY cl ~./ .. -::::::-~· ·-::.<~::·<·... .· '• '• ·-- -- 7 ........ /·.......... TAIL IIOTOII CONTIIOl CABLES ',\.. ...... '-•'', ·. ·-------:~e:--.-}=~·.. :.· ..····· .... ··<·"::::::·:::::::.:.::::......~..........,,.......................,...... 8 ........ ,')';~~j) ..•-::.............. PI LOTS TAi l ltOTOII : ..············· CONTROl PEDALS (TYPICAl) l'llOT S Sf.'ITCH CYCLIC '\.····· , ,!)'''~ :~, :; ::::':\:: (1/ // ) .......···· ....... STICK (TYPICAl) ..· . ' ..... ' I ' ... I 0,______ CABLE ,>';,.;,( ~'\(.... )\/,), ,],/ I GUAIIO ,, I II I I.. ' \ .····· ............. ' ' ,,, ' ,,, l .......... c :\ ,,, "'' ,:, : .,..,,·.· \ \\\ 11\\. .•·'.11 I ••••"' · ~··... ·· ;. \ ,,, ·''' ::-··· .: .• ······ I~ It~· ,, ...·;•••\l-" .. • •••• HJ,:}:J_....,..... .. . ...../·:<.':~ ~~ .. ,. .···· ,_,,.· \ ~ \ ' '-· ·. .. ::~) ~ COI'ILOTS P'=" COLLECTIVE PILOT'S/COPILOT'S '<:····· l 0 I STICK PEDAL ADJUSTOR TAIL ROTOR QUADRANT (TYI'ICAL) FLIGHT CONTROL SYSTEM 67T5161 109 NOTES 110 COCKPIT PILOT ASSIST CONTROLS SERVOS I COLLECTIVE J ~{BOOST l ,.... .. f- ....._ FORWARD • 2ND .. h._ .. .. -~ .. 1ST -,, - 0 ..... SASI -...-2ND - J. I I ...., PITCH { TRIM 1P.B.A. 1 ......,--/r;t\ ....J t-' .. f- t-' t-' -... 1ST AFT '~:;;· - ...... ~PITCH CYCLIC BOOST SASl 0 ..... T ... - ROLL - 2ND -~~ .. ~ h ... 1ST . -f--1 -LATERAL '- [TRIM 1 SAsl MAIN ROTOR (PRIMARY) SERVOS 1 PEDALS ~BOOST I MIXERI TRIM FLIGHT CONTROLS SIMPLIFIED U35-518 BLOCK DIAGRAM 67T5032 ...: .. T A I S L E RR ov To 0 R ... ...... 2ND STAGE PRISERVO SWASH PLATE ...... ...... N COLLECTIVE STICK ~ CABLES --~ ~ I I .J 2ND STAGE ,, I .... I L __ wII I Ell IJ lSi STAGE TAIL ROTOR SERVO COLLECTIVE FLIGHT CONTROLS BLOCK DIAGRAM U35-196 67T5066 LANDING LIGHT CONTROL-- ....... ENGINE SPEED ....... w TRIM SEARCHLIGHT CONTROL @ COLLECTIVE STICK GRIP U35-153 67T5027 0 > D:: Ill en ... en 0 0 ali Ill :J -> ... u Ill ... ... 0 (.) > Q. 0.:::..:: ...... _ Oz t (/)....1 :::l:::..:: o.z _...... z 114 PILOT VALVE PIVOT POINT SLOPPY ~ e ) BYPASS VALVE I ...... LINK' ...... Vl RETURN PRESSURE~£, PIVOT POINT OUTPUT INPUT u3s-l!l $ 1 TYPICAL SERVO SCHEMATIC 68H5045 LIJ > LIJ LIJ _. U) LIJ >_. < > ... 0 _. - a. Ill > ..... c >.,_ l-en Ow ........ - A.. i c~ -~ I 1 z c 116 PITCH BIAS ACTUATOR / COLLECTIVE PITCH ~ ~ ~ ......, , TAIL ROTOR SERVO U35-124 FLIGHT CONTROL MIXING UNIT SCHEMATIC &nsoo2 FROM COLLECTIVE COLLECTIVE COLLECTIVE ,...... ,...... 00 YAW MECHANICAL TO YAW LATERAL LONGITUDINAL LONGITUDINAL REASON ANTI-TORQUE LATERAL LEAD (RIGHT TRANSLATION) ROTOR DOWNWASH ON STABILATOR TAIL ROTOR LIFT VECTOR e e INPUT LINK~ ~••, SLOPPY ...,.--' I '1 1 ./ ........-..\ I LINK PRESSURE SWITCHES ~ OUTPUT ...... ...... LINK \.0 QUICK DISCONNECT JAM TEST COUPLINGS BUTTON ---r PRIMARY SERVO U35-217 68H5046 GOOD STAGE -· ...... N 0 ~PROJECTILE DAMAGED CYLINDER DAMAGE TOLERANT BOOST OR PRIMARY POWER PISTONS -DAMAGED U35-499 68H5049 e -• GOOD STAGE ...... N ...... FRACTURED POWER PISTON SEGMENTS FRACTURED BOOST OR PRIMARY U35-500 POWER PISTON 68H5050 0 AROUND K STICK TRIM\ ~NABLE SWITCH\ CARGO HOS~ITCH ...... N N \ ~RELEASE TRIM RELEASE SWITCH-.~ ~ ICS-RADIO RADIO CONTROL PANEL LIGHTS ~ KILL SWITCH~ ~ ) l'L;;;;;;;;;;;;;;;;;;;;;;;;;; CYCLIC STICK GRIP U35-188 6nsoza e --- SAS PITCH BOOST FWD OR AFT SERVO PRI SERVO REF ...... N w PITCH FLIGHT CONTROLS BLOCK U35-197 i DIAGRAM .· 67T5067 SERVO VALVE TRIM VALVE ELECTRICAL CONNECTION ...... N -!"- PITCH TRIM SPRING ASSEMBLY SAS ACTUATOR PITCH BOOST QUICK DISCONNECT COUPLING 68H5057 e e -•- - - ---·· ----- • SAS TRIM ACTUATOR SWASH PLATE L-~ lc b I CYCLIC MIXING UNIT l;ijl:················ · li········O STICK ........ N I Vl ROLL FLIGHT CONTROLS BLOCK DIAGRAM U35-195 67TS065 SAS /ACTUATOR SAS SERVO VALVE ,_. N (]\ ~~OUTPUT LINK QUICK DISCONNECT COUPLING INPUT LINK-- U35-218 ROLL SAS ACTUATOR 61H5055 • e e • 2ND STAGE PRISERVO SWASH PLATE YAW BOOST 1ST STAGE SERVO PRISERVO CABLES --~ ..... N ....... I I L __ I I V: '2ND STAGE I 1: :::-J. ....... !! ' ~ !' m;t I '1 1ST STAGE TAIL ROTOR SERVO YAW FLIGHT CONTROLS U35-516 BLOCK DIAGRAM 35«8005 1-' N co F ~ CONTROL r-J PEDALS YAW SERVO 0 MIXING UNIT TAIL ROTOR CABLE QUADRANT U35-780 NORMAL OPERATION 67T4189 e TAIL ROTOR SERVO • • ----~---~-- 129 PRESSURE SW-ITCH OUTPUT LINK/ ~ l.V 0 PILOT VALVE INPUT--------.._ CENTERING SPRING HOUSING SLOPPY LINK U35 -419 TAIL ROTOR SERVO-EXTERNAL VIEW 67T5120 e -• VIEW A-A INBOARTOION PLATE~ RETEN ~ INPUT CONTROL ROD 6 EACH NUTS 8 WASHERS PITCH CHANGE SHAFT / SEAL AND RETAINER /TAIL ROTOR [ ' • SERVO TAIL ROTO R GEAR BOX M~83248/ 1 · 249 TAIL ROTOR SERVO INSTALLATION BH -872 68·~~07 ' 131 NOTES 132 HYDRAULIC SYSTEMS The hydraulic systems provide hydraulic pressure to operate the flight control servos and to start the APU. There are three hydra~lic systems. 1. No. 1 or 1st stage hydraulic system 2. No. 2 or 2nd stage hydraulic system 3. Backup hydraulic system The No. 1 hydraulic system supplies hydraulic pressure from the No. 1 pump module to the No. 1 transfer module. From the transfer module, pressure is supplied to the 1st stage of the primary servos and the 1st stage tail rotor servo. The No. 2 system supplies pressure from the No. 2 pump module to the No. 2 transfer module. From the transfer module, pressure is supplied to the pilot assist servos and 2nd stage of the primary servos. During normal flight the backup system is not working. The backup hydraulic system automatically supplies hydraulic power to the No. 1 and/or No. 2 hydraulic systems if either one or both lose hydraulic pressure, and 2nd stage tail rotor servo if first stage loses pressure. The backup system is also used to check all hydraulic systems during ground operations. The APU accumulator is hydraulically charged by the backup system. NOTES: 133 HYDRAULIC SYSTEM OPERATION The hydraulic pump modules are combination hydraulic pumps and reservoirs. The No. 1, No. 2, and backup pump modules are i dent ical and interchangeable with each other. The No. 1 pump module is mounted on and driven by the left accessory module of the main t r ansmission. The No. 2 pump module is mounted on and driven by the right ac·cessory transmissi on module. The backup pump module is mounted on and driven by a 3-phase, 400 Hz, 12 3/4-HP AC electric motor. Each hydraulic pump has two filters: a pressure f i lter and a return filter. A red indicator button on each filter will pop out when the filter is dirty. The No. 1 and No. 2 transfer modules connect hydraulic pressure from the pump modules to the flight control servos. Each module is an integrated assembly of shutoff valves, shuttle val ves, pressure switches, check valves, and restrictors. The modules are interchangeable. The shuttle valve is spring loaded to the open or normal position. If 1st or 2nd stage hydraulic pump fails, the shuttle valve automatically transfers backup pump pressure to the failed system. The utility module connects hydraulic pressure from the backup pump to the No.1 and No. 2 transfer modules, the 2nd stage of the tail rotor servo, and the APU accumulator. There are three primary servos : the forward servo , aft servo, and the lateral servo. The servos provide a power boost to the ma i n rotor flight controls. They also reduce feedback forces from the main rotor head. Each servo has two independent stages (1st stage and 2nd stage). Each stage has an independentpiston, valve housing, and hydraulic supply, but the input linkage is common.The servos are interchangeable . The primary servo manifold connects the servos to the No. 1 and No. 2 t ransfer modules. Each stage of a primary servohas a ballistic tolerant feature built in so that i f a projectile should damage one stage, that stage wi ll be inoperative, but will not stop the other stage from operating properly. The collective and yaw boost servos reduce stick f orce and flight controlfriction. The SAS actuators are part of a stablilzation system that gives rate dampeningfor the helicopter in the roll , pitch, and yaw axi s . The pitch trim actuator control s the attitude of t he helicopter. Pitch trim maintains the position of the cyclic stick. The tail rotor servo is in the tail gearbox. It f urnishes a power boost tothe tail rotor controls. The servo has two independent stages, 1st and 2nd stage. 134 NOTES 135 COMPONEtiT LOCATION e: The major components of the hydraulic system are three hydraulic pump modules, two transfer modules, a utility module, a pi lot assist module, three primary servos, three pilot assist servos, three SAS actuators, two tail rotor servos, an APU accumulator, an APU handpump, and a refill handpump. Most of these components are grouped together on the upper deck in front of the main transmission. The servos are connected to the hydraulic modules through manifolds and self-sealing couplings. NOTES: 136 e ---- e BACKUP NO.1 PUMP NO.2 PUMP PUMP + + + NO.1 ..... , .... UTILITY .. _. NO.2 ..... .... .. TRANSFER MODULE TRANSFER MODULE MODULE TAIL ROTOR SERVO • ~, .... ACCUMULATOR ~ w .. -....J ... ,. I ..... , ~, .. .. CJ:J FORWARD r::.... ~, ~ ...... ~, .... .... ~ :1:::::3 AFT t:... ~ ... , ... ... .. ..... 1r:J LATERAL ... ..... I=- ... ..... r-+t . PRIMARY SERVOS PI LOT ASSIST SERVOS HYDRAULIC SYSTEM BLOCK DIAGRAM UJS-222 67T5005 NOTES • SELECTOR NO. Z PUMP VALVE MODULE ·._ ___ _ GTR zr·····..·:······: /(~::::~:;?? '11! ' : : _.-·· ... r PILOT ASSIST I ,,Qv. ' "\<...,\_ MODULE ( I :: ( :' : "\\\ : : •, [~~tt{l:\J r············l j\ 0 oPUMP z ~ ~~ 0 C/) 0:: ~ CHECK VALVES , ••••••••••••••••••••••••••••••••$ ••••••••••••••• NO. 1 PUMP MODULE Figure 5-1. Hydraulic systems schematic diagram (Sheet 1. 1 of 2) 5-2.2 Change 8 147 ADVISORY PANEL BACKUP HYD DC ~CONTR BATT ~ ______________j BUS ~REUEF VALVE • E)- APU ACCUM HANDPUMP P302 J302 l APU : ACCUMULATi>R : PRESSURE : SWITCH : .:·: ..................~··········: NITROGEN SERVICING VALVE - PRESSURE GAGE I ~} . ~L_! t{_~ APU START . . APU START VALVE ~ PISTON ; ... APU 11 POSITION INDICATOR/ (SEE NOTE 2) WINTERIZATION KIT ACCUMULATOR S 45638 L1 (C8) e e e DIM ITlADY UPI'U DAY t~ ~!~t~ a aT I LASH LDWU NIGHT BACKUP HYD HYD PUMP LEU UST Oil RESET x~ i~ ON TUT ~ L (JD\01'~~ @ "~ t-' FUEl TAll G~i! I.A}I ;' ''V~L~~~::-: :\: <'«~ -..--' s 2 "' -·-' \7 ----·· · ·~II TRIM FPS- SAS 1 S ~~~~~~~~ ~@ L-POWER O.N RESET J ~ @) I~//~ I G U35-674 PILOT'S COMPARTMENT DIAGRAM 67Tl141 0 c A u T I 0 N + A 0 y I s 0 R y 0 I #1 FUEL LOW #1 FUEL I PRESS #1 ENGINE OIL PRESS #1 ENGINE OIL TEMP CHIP #1 ENGINE #1 FUEL FLTR BYPASS #1 ENGINE STARTER TAIL ROTOR QUADRANT MAIN XMSN OIL TEMP CHIP INPUT MOL-LH CHIP ACCESS MDL· LH MR DE -ICE FAIL MAIN XMSN OIL PRESS #1 ENG ANTI -ICE ON APU ON #1 GEN #2 GEN #1 GEN BRG #2 GEN BRG #1 CONY #2 CONY AC ESS DC ESS BUS OFF BUS OFF BATT LOW BATTERY CHARGE FAULT GUST PITCH BIAS LOCK FAI L #1 OIL #2 OIL FLTR BYPASS FLTR BYPASS IRCM INOP INT XMSN TEMP • SAS OFF '?I IFF CHIP INT XMSN ~====~ CHIP TAIL XMSN CHIP MAIN MOL SUMP APU FAI L II MR DE-ICE FAULT TR DE -ICE FAIL I(.#1L~~Y:;z ~~~2L~~Y;l #l ENG INLET #2 ENG INLET ANTI -ICE ON ANT I -ICE ON II PRIME BOOST APU GEN ON PUMP ON LOG LT ON l~u~~~.VII CARGO BRT/DIM HOOK ARMED HOOK OPEN @ PARKING EXT PWR BRAKE ON CONNECTED TEST e! 0 #2 FUEL LOW #2 FUEL PRESS #2 ENGINE OIL PRESS #2 ENGINE OIL TEMP CHIP #2 ENGINE #2 FUEL FLTR BYPASS IS:TRIM FAI\ll I ·~:~~T I 0 CHIP INPUT MDL · RH CHIP ACCESS MOL-RH I I 0 144 e -e FILTER INDICATOR BUTTON ------~ r·~·· , ....... • • • II +:- Vl = =~ I I I lJ.I ~---DEPRESSURIZING -~-~-:-~..... R VALVE E F F I U L L LEGEND L L C --_m_____L I\\ \\) PRESSURE •••••• RETURN LEVE.L INDICATOR WINDOW SUPPLY ------DRAIN I ·· 0 I ·HYD·RAULIC PUMP MODULE U35-119 IIH5012 LEAK ISOLATION SYSTEM The No. 1 and No. 2 hydraulic systems each contain a leak detection and leak isolation system. For example, if a leak should occur in the first stage hydraulic system, the "No. 1 RSVR LOW" light will come on. This light will cause the first stage tail rotor shutoff valve to close and the second stagetail rotor shutoff valve to open to isolate the leak. This also turns on the backup pump to supply pressure to the second stage of the tail rotor servo. If the leak is not isolated and continues, the "#1 HYD PUMP" light will come on and open the first stage tail rotor shutoff valve to regain first stagetail rotor servo operation and close the second stage tail rotor shutoff valve. NOTES: 146 LEGEND -PRESSURE (3050 PSI) ~~ PRESSURE (1000 PSI) ... RETURN E:3 DRAIN m:m:D SYSTEM FILL (25 PSI) E!!!] GAS INDICATES MODULES SUPPLY MECHANICAL ;-·- I EFFECTIVITY FOR HELICOPTERS WITHOUT WINTERIZA· TION KIT INSTALLED. I ..... NO. 2 PUMP MODULE CHECK VALVE SERVO NO. 2 DC ~WARN PRI BUS ~ ADVISORY PANEL I NOTE 1. 1000 PSI PRESSURE WHEN PITCH/ TRIM OPERATING. RETURN WHEN PITCH/TRIM NOT OPERATING. MANUAL BACKUP PUMP MODULE EXTERNAL HYD PWR MANUAL RESERVOIR FILL PUMP MANUAL SELECTOR VALVE UTILITY MODULE "" z ,g ~~-_~~n .... "" 0 PUMP FILTER BYPASS VALVE CHECK VALVES ,................................................ NO. 1 PUMP MODULE 148 Section 5 TM 55-1520-237-23-2 P302 J302 ADVISORY PANEL BACKUP HYD APU : ACCUMULATOR : DC ~CONTR PRESSURE BUS ______ BATI ~ _________j : SWITCH : NITROGEN SERVICING VALVE ~RELIEF ~ PRESSURE GAGE ~VALVE I t;;T ~ I ······-~~i APU APU START VALVE ACCUM HANDPUMP I =... IWWl S 45638.1 (CB) Figure 5-1. Hydraulic systems schematic diagram (Sheet 1 of 2) Change 8 5-2. 1 SERVO o-~ . 1 11 2 HYD PUMP I .....• WARN ~: N0. 2~ PRI DC I ~ CAUTION PANEL BUS I!! ~ I'll~ 11Pil0T- ASSIST SHU TOFFj ~ VALVE • QJ( I ~ l [lq J~ ~ ....... .... ·-·-·-·-...... .....1!! ·-.. ----T r T ,1 I ~ Jl I Ill ~ ·--·~ ._..__. L.. 1... ·~ ~· 1.--J L l i ~ A ~---· ~ vOJGPM ~ ~ --~ 1 I! I ~ .J jJ 2ND STAGE PRIMARY ,..._. I SERVO SHUTOFF ~ VALVE-. ; I ~ ~ NO. 2 TRANSFER MODULE I 1 ; I r{}-o • ~ ; f/1 -=2ND STAGE TAIL 1-~ ROTOR SHUTOFF ~ VALVE ~ ~ y ~ 1.... L..-JI ~I J ~-1.... t-. ·• L-. L-. L-. _; ~~ ·~ ._. I· I n I I [I ' II I I I ~l ~ SERVO I I NO. 1 TRANSFER MODULE N0. 1 ~WARN ~I PRI DC ~ BUS 1 11 1 HYD PU;;-, o----A CAUTION PANEL •••••....•..............................................................., . -I #2 PRI 1 ---~a-------- SERVO PRESS ._-----------.-------------- ~ ~ ~ ~ ' ~ ~ ~ ~ ~:+ ~:+ ~I+ ~:+ #l PRI I I I SERVO PRESS! ~ 6 6 '-' 6 "V 6 1 ---------= CAUTION PANEL LAT FWD f01 rol o o ol o o ol 11 I I I I ._ 1ST 2ND .._ ,__ 1ST 2ND ,__ -C STAGE STAGE b,_.. -[ STAGE STAGE b,_.. JAM JAM JAM JAM i TEST TEST ~ TESTTEST BUTTON !!I BUTTON BUTTON ill BUTTON : I! ! I! i 1 • I I PRIMARY SERVO MANIFOLD 1STSTAGE ~ "t........!........t' .. ~--~----------------------~~~~----~ ~1 0.6 0 I#1 ::~t~TR I 1••••• 1, , GPM .(I r-,.. tel 1-~----CAUTION PANEL II 1·----1... (... -r-~- "' + TAIL ROTOR SERVO 149 Section 5 TM 5~ 1520-237-23-2 ~-----] ~ f-- I/ I ~ BOOST SHUTOFF VALVE ~ ~ - ~ ~ PRESSURE ~I+ ~:+ REGULATING .-(SEE NOTE 1)l VALVE I 'VI =---....~~ -1 '"'6 6 PITCH/ TRIM TURN-ON AFT VALVE fOl I ~~rs ~ I o 0 0 l -y-- T ~ I ~ 1ST I 2N~ STAGEJAM STAGE L-_-JAM TEST --TEST SAS ····==SHUTOFF________, BUTTON BUTTON r ~VALVE ' I ; I • 9 n I I i + A--o ll!i ! PILOT-ASSIST r.l MODULE l-SAS OFF J ~ CAUTION PANEL PILOT•ASSIST MANIFOLD ! - ~ ; --. ;···::·-----···;··-···-·a····-~----······ ., , •: ~--··············· = ; = = = I t t· ! III I l I te~ 1 I . 1 I 1 I eeI e1I I e 1 I et 1tI .----·--- ROLL YAW COLL SAS PITCH/TRIM SAS SAS BOOST BOOST + ACTUATOR SERVO ACTUATOR ACTUATOR SERVO SERVO •~>-J +I I ~[)-~-I I 1 1 BOOST .._j SERVO OFF S 45638 2 (CB) CAUTION PANEL Figure 5-1 . Hydraulic systems schematic diagram (Sheet 2 of 2) Change 8 5-3 NOTES 150 *2 HYDRAULIC PUMP i ~tKVU VALVt PILOT •2 - r-TRANS 1\....1 --2ND Sl AGE ASSIST t MOD TAIL ROTOR SERVO r MOD I 1 FORE t-PITCH TRIM ., , en .._ SAS/PJTCH BOOST ~· =;::ao L 1 T AFT ar:: ., , t-YAW BOOST .... en ,__. ..._ COLL BOOST L _l I LAT \..11 ar:: ., 1 ,__. Q.. UTILITY · +1 1--1 ...-MOD TRANS 1ST STAGE TAIL r ... I ~ MOD \... ROTOR SERVO •1 HYDRAULIC BACKUP PUMP PUMP -ACCUMULATOR HYDRAULIC SYSTEM NOTES 152 DEPARTMENT OF AVIATION SUBJECTS SYSTEMS TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama May 1984 File No. 47-4747-9 LESSON EVALUATION FLIGHT CONTROLS AND HYDRAULIC SYSTEM · 1. The flight control mixing unit is provided to eliminate 2. The number 2 hydraulic system supplies hydraulic pressure to the and 3. The first stage tail rotor servo can be manually turned off by the switch marked 4. The components that are pressurized by the number one hydraulic system are the and the -------------------------------- 5. When the backup pump power circuit breaker is out, AC power must be shut off prior to resetting the backup pump power circuit breaker. Otherwise, it is possible to damage the 6. The five caution/advisory lights that will automatically bring on the backup pump in flight are: 1. ; 2. 3. ; 4. ; and 5. 7. There are two components in the aircraft hydraulic systems that can onlybe pressurized by the backup pump. They are and the S. The leak detection/isolation system the flight control -----:=--: hydraulic system by preventing of hydraulic fluid. ~. A . and valve are on the right side upper deck to serv1ce the three hydraulic systems. 1U. The primary servo shutoff switch \'li 11 not shut off both stages of primary servos at the same time because of an 11. When RPMR is above 3U percent, what position will the backup pump switch be in? 153 12. The pitch boost servo is provided to cockpit control forces. 13. The caution light marked will i11 umi nate when a ta....i.....l_r_o-=t_o_r_ca.....,b.....l.-e~b-re_a_,.k-s-.--------- 14. The following is a list of caution/advisory lights that will illuminate when a leak occurs in the #1 tail rotor servo. 15. A leak in the pilot's assist servos will cause the following caution lights to illuminate: 16. The will allow the backup pump to come on automatically regardless of the backup pump switch position wh il e i n f 1i g h t • 17. The backup pump will react within seconds when a main generator is supplying aircraft (airframe) power. 18. The UH-60 is limited to a bank angle when a PRI servo pressure caution light is on. 19. The airspeed of the UH-60 is limited to ____ whenever one hydraulic system is inoperative. 2u. After the failure of two hydraulic systems while VMC, the UH-60 is limited to airspeed. 21. Wh i le flying on an IMC flight plan (actual), you lose two hydraulic systems. What must you limit your airspeed to? _______ 22. You are making a roll· on landing after a total loss of tail rotor control (dual cable breakage). Your initial approach speed is and touchdown speed is _____ 23. You have already experienced a leak in the #1 primary servos and have taken the proper corrective action. The following is a list of caution/advisory lights illuminated: #1 hyd pump, #1 pri servo press, #1 rsvr low, backup rsvr low, backup pump on. If the #2 pri servo press light illuminates, what is your pilot action? 154 POI FILE: 47-4748-5 AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS) TYPE OF INSTRUCTION: PEACETIME HOURS MOBILIZATION c 4.5 4.5 PE 0.5 0.5 LEARNING ORJECTIVE: Given five written questions containing descriptive situations or operating conditions, from memory, the student \'lill \'Jrite lAW TM 55-1520-237-10 the- a. Purpose of AFCS. b. Five switches on the AUTO FLIGHT control panel. c. Procedure for FPS operation. d. Procedure for yaw trim operations hel 0\'1 60 KIAS. e. Procedure for yaw trim operations above 60 KIAS. f. Procedure for making heading hold trim below 60 KIAS. g. Switch used to neutralize stick forces after establishing the desired trim condition. h. Three switches on the stabilator control panel. i. Stabilator WARNING to be observed before applying AC electrical power to the helicopter. j. Procedure for use of the POWER ON RESET switches located on the AFCS panel when a malfunction occurs in AFCS. STANDARD: Five of five questions must be answered correctly to satis factorily complete this objective. LESSON REF ERENCE: TM 55-1520-237-10, chaps 2 and 9. TASK: This objective supports tasks 03-1402, 1501, 1502, 4021, 4027, 6011, and 6501. EQUIPMENT: UH-60 composite trainer; UH-60 power train system panel, DVC 1-11 8 ; UH-60 caution/advisory panel, FR 1330. INSTRUCIONAL ELEMENT: OOAS, STD, ASB. 155 AUTOMATIC FLIGHT CONTROL SYSTEM GENERAL DESCRIPTION The UH-60A automatic flight control system (AFCS) improves the helicopter's flight characteristics in the pitch, roll, and yaw axis. It providesattitude, heading, and airspeed hold; a positive cyclic stick gradient; andautomatic turn coordination. The AFCS consists of three major subsystems. 1. An analog stability augmentation system (SAS 1). 2. An automatic stabilator control system. 3. Digital AFCS. Analog SAS 1 provides improved aircraft handling qualities and stability in the pitch, roll, and yaw axis. Tne analog SAS and digital AFCS interface with the helicopter's mechanicalflight control system by using trim and SAS actuators for a mechanical input to the flight control system. The stabilator control system operates electromechanical actuators thatposition a movable horizontal stabilator. The stabilator provides improvedcontrol in pitch axis and level hover attitude. NOTES: 156 AUTOMATIC FLIGHT CONTROLS SYSTEM The automatic flight controls system (AFCS) is an electro-hydromechanical system that provides inputs to the flight controls system to assist the pilot in maneuvering and .. handling the heli copter. AFCS consists of three major subsystems, analog stability augmentation system, digital AFCS, and the stabilator system, that provide ocillation damping and maintain desired attitude, speed, and heading . The analog stability augmentation system, identified as SAS 1, provides short term correction and rate damping in the pitch, roll and yaw axis and also provides limited attitude hold in the roll axis. The SAS amplifier is the central controller for SAS 1 operation. The SAS amplifier processes input signals and provides SAS actuator drive signals to the pitch, roll and yaw SAS actuators. The SAS actuators provide a mechanical input to the flight controls system. The digital AFCS system provides the following functions: digital stability augmentation system, flight path stabilization, pitch bias actuator control (for positive cyclic stick gradient in the pitch axis), and trim control system. The digital stability augmentation system, identified as SAS 2, provides short term correction and rate damping in the pitch, roll, and yaw axis and also provides limited attitude hold in the roll axis. The flight path stabilization system (FPS) provides for: airspeed/pitch attitude hold; roll attitude hold; heading hold; cyclic (pitch and roll) stick and pedal (yaw) position trim system; and at airspeeds greater than 60 knots, coordinated turn feature that automatically coordinates tail rotor input to the banking angle selected through use of the cyclic stick. The trim system provides a cyclic (pitch and roll) stick and pedal (yaw) position trim reference to maintain the cyclic stick and pedals at the desired trim position. The SAS/ FPS computer is the central controller for all digital AFCS functions. The computer processes input information and provides output to the pitch trim actuator, roll trim actuator, yaw trim actuator, pitch bias actuator, pitch SAS actuator, roll SAS actuator, and yaw SAS actuator. These in turn provide mechanical inputs to the flight controls system. The stabilator system provides the helicopter with stability in the pitch axis by controlling the position of the horizontal stabilator. The stabilator system can operate in either the automatic or manual mode. In the automatic mode, the stabilator amplifiers produce position sensor signals, airspeed sensor signals, pitch rate gyro signals and lateral accelerometer signals. The outputs from the stabilator amplifiers are applied to two electro mechanical actuators which control the position of the stabilator. In the manual mode, the pilot's manual slew switch command will drive the stabilator actuator to any position selected. As in automatic mode, the outputs from the stabilator amplifiers are applied to the actuators which control the position of the stabilator. 157 D:: w > 0 :I: I ~ 0 .... II. D::~ c ... e - ... D:: 0 .... c .... - m c .... en e e • ,_. Vl \0 FORWARD FLIGHT RELATIVE WIND u35-473 STABILATOR AIRFLOW-FORWARD FLIGHT 35KI025 STABILATOR SYSTEM e Aerodynamics The stabilator is a horizontally mounted symmetrical airfoil attached to the tail pylon of the helicopter. Its angle of incidence is controlled by an automatic, electromechanical, servo system. Its angle of attack is determined by angle of incidence and relative wind direction. The relative wind that acts on the stabilator is the result of helicopter airspeed and main rotor downwash. In forward flight, the direction of the relative wind that results from helicopter airspeed depends on pitch attitude. Changes in pitch attitude cause changes in stabilator angle of attack. The stabilator produces either a lifting force or a downward force. Forces developed at the stabilator provide pitch stability in addition to providing lift at the tail. When the helicopter hovers, there is no forward airspeed. Main rotor downwash produces a downward force on the stabilator. This would cause a nose high hover attitude. The automatic control system drives the stabilator trailing-edge-down at low forward airspeeds, to provide an aircraft hover attitude that is closer to level. When the helicopter moves into forward flight, the stabi lator is driven trailing-edge-up~ allowing it to provide pitch stability. Increasing collective pitch would produce a downward force at the stabilator. The effect is minimized by driving the stabilator trailing-edge-down when collective pitch is increased. Th~ automatic control system provides further increases in pitch stability, i :e., rate signals cause the stabilator to move in the direction that develops a force to oppose aircraft motion. The stabilator control system also responds to l ateral acceleration signals.While making a turn in flight, this response is necessary because lateral motion of the helicopter produces .a relative wind that changes angle of attack of the tail rotor blades, thus, a relative wind change affects tail rotor thrust. Changes in tail rotor thrust affect the lift and result in a tendency of the helicopter to change pitch attitude whenever it "slips" or "skids." The stabilator is automatically positioned to provide more or less lift at the tail, as required to maintain pitch attitude. 160 e STABILATOR SYSTEM Positio~ Indicating System Pilots and copilots position indicators provide instrument panel displays for stabilator position. Instrument panel placards list limit airspeeds never to be exceeded for stuck stabilator positions. Stabilator posit i on is controlled by a dual electromechanical servo system. T.he control system is operated in automatic mode, with manual mode serving as "aH////////////////////1 0 < < 0 b: ~ .. ACT #2 POS ~ I II COMMAND ACT #2 POS •I I STAB I FAULT 1 FAULT STABCOMMAND ....... AM PL. : MONITOR MONITOR : AM PL. Q\ V'1 #2 I ACT #1 POS ACT #1 POS #1 L _____ _ ..,_ _____JI ACCEL ACCEL COLL #2 COLL #1 #2 #1 PITCH RATE PITCH RATE A/S #2 A/S #1 #2 #1 STABILATOR SYSTEM U35-906 SIMPLIFIED BLOCK DIAGRAM 351(8180 STABILATOR SYSTEM System Operation Components of the No. 1 Stabilator Control System receive AC and DC electrical power from essential buses. No. 2 system components are powered from the No. 2 primary AC and DC buses. Automatic mode engages when power is applied to the aircraft busses. The control panel auto mode ON light goes on, the STABILATOR caution light remains off and the stabilator drives to about 39 degrees down. The auto~atic mode positions the stabilator to the best position for existing f light conditions without any inputs from the pilot. The control panel TEST button can function only at airspeeds less than 60 knots. It should be used only as a GROUND test. When it is pushed, the stabilator should move up by about 5-l~ degrees and the AUTO mode is disengaged. The control panel MAN SLEW switch disengages the AUTO mode and allows stabilator control between about 9 degrees up and about 39 degrees down (trailing edge). Auto mode may be reengaged by pushing the auto mode reset ON button. NOTES: 166 ANALOG STABILITY AUGMENTATION SYSTEM (SAS 1) SAS improves helicopter handling qualities and stability in the pitch, roll, and yaw axis. SAS 1 operates independently of the SAS/FPS computer to provide redundancy of SAS functions. System Components Actuators Pitch, roll, and yaw SAS actuators are mounted on the transmission deck, attached to the pilot-assist servos. The three electro-hydromechanical actuators link SAS electronic components to the helicopter's mechanical flight control system. They receive hydraulic pressure from a common source. Actuator electrical inputs are supplied by the analog SAS amplifier and the digital SAS/FPS computer. The actuators respond to electrical inputs by moving control linkages that change rotor blade angles without moving the cockpit controls. Amplified The SAS amplifier is located on the floor of the electronics compartment "tunnel", below the instrument panel. It processes aircraft sensor signals to develop command signals that are applied to the SAS actuators when SAS 1 is engaged. It contains a rate gyro that serves as sensor for yaw SAS. Amplifier electrical outputs are limited to allow analog SAS a maximum of plus and minus five percent authority. Gain control circuits double each channel's gain when SAS 2 is switched off and authority remains at five percent. Sensors The No. 1 stabilator control amplifier, supplies the SAS amplifier with pitch rate, lateral acceleration, and airspeed discrete signals. The pitch rate signal originates from a rate gyro inside the stabilator amplifier. The lateral acceleration signal is produced by the No. 1 lateral accelerometer. The airspeed discrete signal is develped by the stabilator amplifier. The airspeed transducer supplies the signal that determines polarity of the airspeed discrete signal. The pilots (No. 2) vertical gyro supplies the SAS amplifier with a signal that represents helicopter roll attitude. The rate gyro, inside the SAS amplifier, supplies SAS with signals that represent helicopter yaw rate. Controls The auto flight control panel SAS 1 and SAS 2 switches provide engage control for SAS. Engaging either or both switches turns on SAS actuator hydraulic pressure. SAS 1 signals are supplied to the pitch, roll, and yaw actuator coils when SAS 1 is on. 167 Indications The caution/advisory panel SAS OFF capsule lights if SAS actuator hydraulic pressure is switched off or lost. 168 AS SHUTOFF VALVE ROll SAS ACTUATOR ~ YAWSAS ACTUATOR I -----YAW BOOST SAS ASSEMBLY PRESSURE SWITCH PILOT ASSIST MODULE TRANSMISSION/UPPER DECK I ~ J I 0 ~···· .... ......... __ ,~.:....-·· . ..-~· \( ··\:::.:::·)·...'!.:....· STABILATOR CONTROLS/AUTO FLIGHT CONTROL PANEL ·< ..... .....( \ ·· ..· ""'---.-ro-------,.J ...··· I .•'.· ~-··-.;:·'::::-· G @ c @ c @: @ ~ ·cs a lft,..f'OIUU"'~n LATERAL - ACCEL lLJl ' tJo~ c @ c @ c ·o·· ® 0 0 " @ 0 0 0 0 @ 8 1AS ACC H 0 0 ... N0.1 STABILATOR AMPLIFIER 0 I I SAS AMPLIFIER I ---~-- SAS 1 COMPONENT LOCATIONS BH-91~ 35K8 189 169 NOTES 170 e FLAPPER VALVE t-' -....1 NOZZLES . t-' ORIFICE FILTER ASSEMBLY ACTUATOR - SAS 2 ----... CENTERING LOCK PIN CENTERING LOCK ADJUSTMENT SERVO VALVE : SPOOL r-----r~ACTUATOR FEEDBACK CENTERING LOCK HOLE SAS ACTUATOR SCHEMATIC U35-557 61H5056 ANALOG STABILITY AUGMENTATION SYSTEM (SAS 1) System Operation Electrical Supply The SAS amplifier receives 115 VAC power from the AC essential bus, throughthe SAS AMPL circuit breaker. The SAS amplifier and auto flight control panel engage circuits are supplied with 28 VDC essential bus power. Hydraulic Supply Hydraulic pressure, required for SAS actuator operation, is supplied by either the No. 2 or the backup hydraulic pump. Pressure, at 3000 psi, is supplied through the SAS shutoff valve on the pilot assist module to the three SAS actuators. The valve may be shut off by turning off SAS 1 and SAS 2 on the automatic flight control panel. Pitch Channel The SAS pitch channel causes the main rotor tip path to tilt in the direction required to oppose pitch attitude changes. This provides the helicopter with "dynamic stability" by minimizing oscillations in the pitch axis. Roll Channel The SAS roll channel causes the main rotor tip path to tilt in the direction required to oppose roll attitude change and keep the helicopter level. This provides the helicopter with dynamic stability and a tendency to remain level in the roll axis. Because an attitude proportional signal is used, a jump in tip path position and roll attitude can be expected if SAS is engaged or disengaged with the aircraft at any attitude other than level. Roll rate and lateral acceleration signals are used to provide automatic turn coordination. The roll rate signal is supplied by the SAS roll channel. The No. 1 stabilator amplifier supplies filtered lateral acceleration and airspeed discrete inputs to the SAS amplifier. The airspeed discrete signal will control switch circuits that disable automatic turn coordination below 60 knots. At airspeeds above 60 knots, the discrete voltage operates switch circuits that apply roll rate and lateral acceleration signals to yaw SAS. Yaw Channel The SAS yaw channel causes the tail rotor to change pitch to oppose changes in helicopter heading. This provides dynamic stability by minimizing yaw oscil lations. At airspeeds above 60 knots, yaw SAS also provides automatic turn coordination. A yaw rate gyro, inside the SAS amplifier, produces an electrical signal that is proportional to rate of heading change. 172 e -e AFCS 47/48-4748-4 -----------t•2 VERi GYRO SAS •t AlP AJS FROI •t USED TO TELLIF Att IS ABOVE &OK STAB AlP USED IF A/C IS ABOVE 60K TO ~~ ].. I IPITCHIATECIIGI c ROIUCOIPENSATE FOR SLIP OR SKID UTAC ELF FROPl STAB AMP •t STAB AlP OPERATION OF ROLL SA$: RAW SIGNAL FROI •2 VERT OPERATION OF PITCH SAS: GYRO IS PROCESSED TO COlE USES PITCH RATE GYRO SIGNALOPERATION OF JAW SA$ BELOW 60 KIA$ YAW RATE GYRO IS USEDUP WITH DERIVED RATE WHICHABOVE 60 KIAS, LATERAL ACCEL SIGNAL IS DOES THE SAME THING AS RATE GYRO,COMPARED TD YAW RATE GYRO AND GREATEST 1-' ANY DIFFERENCE 1$ TAKEN OUT. -....J w SIGNAL IS USED. ~ t YAW ROLL PITCH SA$ us SAS ACT ACT ACT NOTE : RATE GYRO TELLS WHAT DIRECTION A/C IS - GOING AND HOW FAST ' IT IS GOING IN THAT DIRECTION. LATERAL ACCEL RELAYS WHEATHER THE AIC IS IN ASLIP OR SKID DURING ATURN YAW RATE GYRO LOCATED IN NOSE OF A/C A/S FROM~2 USED TO TELL IF THE "'• ...... 11 .. -...,J STAB. AMP. A/C IS ABOVE GOK ~ LAT ACCEL FROM USED TO COMPENSATE *2 STAB AMP FOR SLIP OR SKID ABOVE GOK AFCS 47/48-4748-4 SAS/SPS COMPUTER SAS *2 ... ... .....,. ..... ,._. .,. .....,. ......,. *1 VERT GYRO PITCH RATE GYP, •2 STAB AMP ROLL RATEl GYRO YAW ROLL PITCH SAS SAS SAS ACT ACT ACT e DIGITAL AUTOMATIC FLIGHT CONTROL SYSTEM (AFCS) Digital AFCS provides the following control functions: a. Stability Augmentation for SAS 2 identical SAS 1. to and redundant with b. Stick trim which holds the cyclic stick at position. a pilot selected c. Pedal trim which holds the tail rotor position. pedal s at a pilot selected d. Flight path stabilazation (FPS) which maintattitude and airspeed in the pitch axis, atand heading in the yaw axis. In addition, turn features. ains the helicopters titude in the roll axis, it provides coordinated e. Pitch bias actuator control to provide the airspeed to cyclic stick gradient in pitch. pilot with a positive The digital AFCS central controller is the SAS/FPS digital computer. It is programed to scan, monitor and compare input sensor signals on a periodic basis, and store this information into computer memory for AFCS calculations. The output drive signals are used for trim, SAS, FPS, and pitch bias actuators. NOTES: 175 NOTES 176 STICK SAS SHUTOFF PITCH TURN ON ~ TRIM VALVE VALVE SWITCH TRIM RELEASE SWITCH SA$ ROll SERVO VALVE SAS PRESSURE SWITCH STABILATOR CONTROLS AUTO FLIGHT CONTROL PANEL P'ITCH BIAS PILOT ASSIST MODULE ACTUATOR SA$ PITCH0 SERVO VA LVE l I PILOT'S AND COPILOT'S I ~ 1 CYCLIC STICK I (!)~----------- r( ~ 1r ·,·<'';r::o. ·.. ~ COllECTIVE BOOST PRESSURE SWITCH ROll TRIM ACTUATOR ~~:",,~;:::<;.....~ ~. ( · ,...> _.,,~lJt;,;~""'"' .... ::\ ;;-~~·-~ ~ ..,;.;;;;,?;;;:::::;r::.::·--· YAW BOOST P'RESSURE 0 LOWER CONSOLE I SWITCH YAW TRIM ACTUATOR I ,·. TRANSMISSION/UPPER DECK AIR DATA TRANSDUCER .,... ~----------------------------~0~----------------------------~ 0 () ~·,-~ IAI eeow UTI Wllf .... YAU'I. c c c c 0 @ 00000 CJ • 9~9.9.2 : ..... o.. snc• •11/Cc';.:::C" 0 11110 act Cl ~ 0~::, ©>::. L.__ ... NO( •• · ~a0 fuo. ~ ' © 7\1 c .. ·o.. 0 " c c c c c '"' ~ c 0 0 0 0 0 ~.. '" " 0 @ 0 0 ... JU I J ill COPILOT'S VERTICAL GYIIO N0.1 AND N0.2 STABILATOR AMPL~F'IERS CN -1314A SAS/F'PS COMPUTER An CABIN OVERHEAD 0 0 I I I DIGITAL AFCS COMPONENT LOCATIONS IH-111 35K8190 17'? NOTES 178 e e • AFCS 47/48·4748·4 SAS/FPS COMPUTER I A/S I PROCESSED :; 1-11 ......~ COMMANDS ....... PITCH RATE ,...... PITCH BIAS -....! 1.0 . ACTUATOR VERT GYRO HORITY 180 K ~ ~ en ~~· 'c ~ ~~ A/S VS. CYCLIC STICK POSITION ~~ 80 K STICK POSITION DIGITAL AFCS MALFUNCTION INDICATIONS/EFFECTS Indicator lights on the caution/advisory panel and on the automatic flight control panel provide the pilot and maintenance personnel with information necessary to locate and repair AFCS malfunctions rapidly. The computer will disable only the affected functions during flight. NOTES: 180 e e e SAS/FPS COMPUTER FAILED PITCH, ROLL OR YAW OR PITCH TRIM SERVO VALVE SAS ACTUATOR VALVE COIL CIRCUIT OPEN OR FAILED OR PITCH, ROLL OR YAW PITCH, ROLL OR YAW SAS ACTUATOR VALVE RATE GYRO FAILED CIRCUIT OPEN RESET ....... R 00 ....... E PITCH TRIM SERVO COIL s SHORTCIRCUIT:OR E T PITCH, ROLL OR YAW TRIM ACTUATOR FAILED OR .... POWER ON RESET ___J ROLL STICK FORCE *COULD BE AFFECTED CHANGING TOO FAST OR STABILATORPEDAL FORCE CHANGING SYSTEM FAILURETOO FAST NOTE: WIRING FAILURES CAN*PITCH, ROLL OR YAW CAUSE ANY FAILURERATE GYRO FAILED INDICATION STABILATOR/FLIGHT CONTROL PANEL - FAULT ADVISORY LIGHT$ (LEFT) U35-792 ~=~NO. 1 OR NO. 2 ~=~ N0. -1 OR NO. 2 LATERAL ACCELEROMETER COLLECTIVE STICK FAILED TRANSDUCER FAILED R NO. 1 OR NO. 2 VERTICAL E GYRO OR COMPASS 5 ....... E A/S IGYRO SYSTEM FAILED 00 N T POWER ON RESET _J ~=~coULD BE AFFECTED ~=~AIRSPEED OR BY STABILATOR AIR DATA SYSTEM FAILURE TRANSDUCER NOTE: WIRING FAILURES CAN FAILED CAUSE ANY FAILURE INDICATION STABILATOR/FLIGHT CONTROL PANEL - FAULT ADVISORY LIGHTS (RIGHT) U35-793 .::.. e e e e -e FAILURE ADVISORY LIGHTS FAILED COMPONENT REF CAUTION LIGHTS NO. 1----: PITCH BIAS ACCL CLTV TRIM FAIL II FAIL I ;!f;~ CPTR SAS 2 ~·-,; ~ ~ II II '~: :1. "! TRIM RGYR A/S GYRO "' ;· ~ FLT PATH !J. II STAB II s·,."" .;w, l ~ ~ ~~ •" ~ ~~ -~ ~i ·..<%a ·1", "' J ~ COMPUTER (POWER SHUTDOWN) 1 X X X lX X ifi i. " X -~ ; l<; ~i :~ VERTICAL GYRO (PITCH) 5 X X X "' e PITCH RATE GYRO 6 ~ X X X . . "4; .,f,' ...... J; ~ ~~ X w AIRSPEED OR AIR DATA TRANSDUCER 7 X 00 ,'"';; w X ., PITCH STICK POSITION TRANSDUCER 8 VERTICAL GYRO (ROLL) 9V X . ., X 9D DIRECTIONAL GYRO ~ X e LATERAL ACCELEROMETER 10 X rx X ' rx COLLECTIVE STICK POS. TRANSDUCER 11 <~ · r, PEDAL FORCE TRANSDUCER 12P X lX LATERAL STICK FORCE TRANSDUCER 12L X lX PITCH TRIM ACTUATOR (Fb) 13P lX X lX ROLL TRIM ACTUATOR (Fb) 13R lX X . lX 13Y YAW TRIM ACTATOR (Fb) lX X lX COMPUTER (ROLL TRIM D/A) 14R [X X lX COMPUTER (YAW TRIM D/A) 14Y [X X lX COMPUTER (WEIGHT ON WHEELS) 15 lX XX ~ ROLL RATE GYRO 16R YAW RATE GYRO 16Y X PITCH SAS VALVE/COMPUTER DRIVER 17P [X ROLL SAS VALVE/COMPUTER DRIVER 17R X YAW SAS VALVE/COMPUTER DRIVER 17Y DIGITAL AFCS -COCKPIT MALFUNCTION INDICATIONS USM ·12 TRIM FPS REF. PITCH BIAS SAS NO. ACTUATOR PITCH ROLL YAW PITCH ROLL YAW PITCH ROLL YAW 5 INOP. · CENTER INOP. NO RATE GYRO FAILURE DETECTION 9V INOP. NO TURN COORD NO RATE GYRO FAILURE DETECTION 9D " INOP. NO RATE GYRO FAILURE DETECTION 7 A/S SIGNAL-120 KTS NO COLL/YAW NO A/S HOLD NO TURN COORD TURN COORD ON6 INOP. · CENTER INOP. INOP.16R INOP. NO TURN COORD 16Y NO TURN COORD INOP. 10 NO TURN COORD 11 ' NO TURN COORDNO COLL/YAW 8 . NO A/S INTEG 12L NO BEEP ON GND NO INTEG NO TURN COORD 12P " NO INTEG 2 INOP. · LAST POSITION 3 INOP. · LAST POSITION '• 13P INOP. INOP. 13R . - INOP. INOP. 14R ' INOP. INOP. 13Y INOP. INOP. 14Y •; INOP. INOP. 17P INOP. 17R INOP. 17Y 15 ' INOP. INTEG ON GND INTEG ON GND INTEG ON GND RATE GYRO FAILURE DETECTION ON GND 4 INOP · CENTER INOP. INOP. INOP. NO RATE GYRO FAILURE DETECTION 1 INOP · LAST POSITION INOP. INOP. INOP. INOP. -----·-INOP. INOP. INOP. I INOP. I INOP. ----· DIGITAL AFCS-EFFECTS OF MALFUNCTIONS USM -13 e e e e -e SJ.ill.lill STABlLATOR CONTROL PANEL STU. MODE SILECT IIUTO J STU. IU. SlE W. - ~ ~ SUI. TEST SAS/ FP SYSTEM SIGNALS ouT M\uoRS !~PI PITCH UTE CYIO ITCH ~ SIGNALS IN • 1 PITCH CYIO ___.,..-:viO _ 8- PILOTS • 2 VUTIC _ __, ,_. VEITICAL CYIO IOU s '1 ACCl. co t l i S l · D ICU -----,....,....d-_.... I ln IYI,.~T.l--I ~ l ' 1 ACCl. _ liS l · DOCEI OUT CYIO YAW 8 :L •1 110•2 ACCL. IFILTEIEI ) S ARE Ill DATA HICU >TAIILATOII YAW lATE CYIO AliP SIGNALS IN I-OUCEI TRIM ACTUATORS .. I m!" I I un ~ fJ I I I ~ ___________ , ,..-----J I , f PIA SIC NILS II A1S l·DUCE • 2 PITCH UTE CYID c II •2un. n1o. • I Till , l SIGUlS II I SISIFPS DMPUTOI IQ I PIA Till I ~ T f:;:F==-! I ClTV '--~AJS l · DUCEI PilOT'S '----+-----+++-~men~: CDITIOL SWITCHlS SIS 2 FPS FPS L~COPilOTS) •t VUTIC CYIO I 1-J.--liS l·DICEI 1 Lt--·• ·cm rJ.--'1 PITCI lATE GYIO '----+--•tacci. I r --!---lSI 43 COMPASS ' IIUDIIIC ) 1-----....J ... ----------•---------J t l__ •z PITCIIITECYIO L lOLl UTE CYIO SENSOR LOCATION SIGNAL TO PURPOSE ...... NUMBER 2 FLIGHT 1. NUMBER 2 AUTOMATIC CONTROL CIRCUITS ()\ COLLECTIVE CONTROL STABILATORSTICK MIXER AMPLIFIERPOSITION UNITSENSOR CXl 2. SAS/FPS COMPUTER YAW FLIGHT PATH STABILIZATIONCOMPARE CIRCUITS AFCS COLLECTIVE STICK POSITION U35-502 SENSOR -NUMBER 2 e e -e -e • SENSOR LOCATION SIGNAL TO PURPOSE NUMBE~ ~ FLIGHT 1. NUMBER 1 AUTOMATIC CONTROL CIRCUITS t-' COLLECT, VE CONTROL STABILATOR I 00 STICK MIXER AMPLIFIER '-1 POSITION UNIT SENSOR 2. SAS/FPS COMPUTER YAW FLIGHT PATH STABILIZATION CIRCUITS AFCS COLLECTIVE STICK POSITION . SENSOR -NUMBER 1 U35-501 SENSOR LOCATION SIGNAL TO PURPOSE ...... 00 00 NUMBER 2 NOSE NUMBER 2 STABILATOR a. AUTOMATIC CONTROL CIRCUITS LATERAL ELEC-AMPLIFIER ACCELERO-TRONICS b. FILTERED AND NULLED SIGNAL TO: METER COMPART-SAS/FPS COMPUTER COMPARE MENT LOGIC RIGHT HAND· SIDE AFCS LATERAL ACCELERATION SENSOR U35-504 NUMBER 2 e e e -e • SENSOR LOCATION SIGNAL TO PURPOSE NUMBER 1 NOSE NUMBER 1 STABILATOR a. AUTOMATIC CONTROL CIRCUITS LATERAL ELEC-AMPLIFIER ...... CXl b. fiLTERED AND NULLED SIGNAL TO: \.0 ACCELERO-TRONICS 1.) SAS 1 AUTOMATIC TURNMETER COMPART-COORDINATION MENT LEFT HAND 2.) SAS/FPS COMPUTER AUTO-SluE MATIC TURN COORDINATION SAS AND FPS AFCS LATERAL ACCELERATION SENSOR - NUMBER 1 U35-503 SENSOR ROLL RATE GYRO ..... 1.0 0 NO. 1 YAW RATE GYRO NO.2 YAW RATE GYRO LOCATION NOSE BRACKET SAS AMPLIFIER NOSE BRACKET SIGNAL TO PURPOSE SASiFPS COMPUTER SAS 2 ROLL STABILITY SAS AMPLIFIER YAW STABILITY SAS 1 SAS/FPS COMPUTER YAW STABILITY SAS 2 -------------·· ----------------- U35-505 AFCS SENSORS-ROLL AND YAW RATE e e e e SENSOR LOCATION SIGNAL TO PURPOSE NUMBER 2 NUMBER 2 1. NUMBER 2 PITCH STABILITY ..... PITCH RATE STABILATOR STABILATOR SYSTEM \0 ...... GYRO AMPLIFIER CABIN 2. SAS/FPS COMPUTER a. PITCH BIAS ACTUATOR CONTROL OVERHEAD b. NUMBER 2 SAS PITCH STABILITY RIGHT HAND c. PITCH FLIGHT PATH SIDE STABILIZATION AFCS PITCH RATE SENSOR -NUMBER 2 GYRO U35-507 SENSOR NUMBER 1 ,..... PITCH RATE N \0 GYRO LOCATION NUMBER 1 STABILATOR AMPLIFIER CABIN OVERHEAD LEFT HAND SIDE SIGNAL TO PURPOSE 1. NUMBER 1 PITCH STABILITY STABILATOR SYSTEM 2. SAS AMPLIFIER NUMBER 1 SAS PITCH STABILITY AFCS PITCH RATE SENSOR -NUMBER 1 U35-506 GYRO e -e e SENSOR LOCATION SIGNAL TO PURPOSE AIR DATA COCKPIT 1. COMMAND a. AIRSPEED REFERENCE TRANSDUCER RIGHT HAND INSTRUMENT b . ALTITUDE REFERENCE SIDE SYSTEM c. ALTITUDE RATE REFERENCE ...... \0 CANTED w BULKHEAD 2. NUMBER 2 a. AUTOMATIC CONTROL CIRCUITS STABILATOR b. BUFFERED SIGNAL TO NUMBER 1 AMPLIFIER STABILATOR AMPLIFIER 3. SAS/FPS COMPUTER AIRSPEED COMPARE U35-514 AFCS SENSOR -AIR DATA TRANSDUCER SENSOR LOCATION SIGNAL TO PURPOSE AIRSPEED COCKPIT 1. NUMBER 1 a. AUTOMATIC CONTROL CIRCUITS TRANSDUCER LEFT HAND STABILATOR b. TEST CIRCUIT DISABLE SIDE AMPLIFIER c. BUFFERED SIGNAL TO NUMBER 2 I-' \0 +:-CANTED STABILATOR AMPLIFIER BULKHEAD 2. SAS/FPS COMPUTER a. AIRSPEED HOLD b. PITCH BIAS ACTUATOR CONTROL c. AUTOMATIC COORDINATED TURN (LOGIC) AFCS SENSOR -AIRSPEED TRANSDUCER U35-513 · e e e e -e SENSOR LOCATION SIGNAL TO PURPOSE AN/ASN-43 NOSE 1. DOPPLER NAVIGATION HEADING REFERENCE COMPASS ELEC-SYSTEM TRONICS 2. VHF OMNIDIREC· HEADING REFERENCE GYRO COM PART-TIONAL RANGE CN-998 MENT RIGHT HAND 3. COMMAND HEADING REFERENCE ..... 1.0 SIDE INSTRUMENT SYSTEM V1 CANTED BULKHEAD 4. PI LOT'S AND HEADING INDICATION COPILOT'S HSI'S (COMPASS CARDS) 5. SAS/FPS COMPUTER a. HEADING HOLD b. TURN COORDINATION c. SAS 2 COMPARE (RATE) U35-512 AFCS SENSORS -HEADING INFORMATION SENSOR LOCATION SIGNAL TO PURPOSE NUMBER 2 NOSE 1. PILOT'S VSI (NORM-) a. ATTITUDE INDICATION VERTICAL ELEC-AND COMMAND INST b. ATTITUDE REFERENCE GYRO TRONICS SYSTEM (NORM) CN-1314/A COM PART ...... 0' \0 MENT 2. COPILOT'S VSI (ALT) ATTITUDE INDICATION RIGHT HAND SIDE 3. SAS/FPS COMPUTER ROLL ATTITUDE COMPARE 4. SAS AMPLIFIER ROLL STABILIZATION (DERIVED RATE FOR SAS 1) AFCS ROLL SENSOR -NUMBER 2 GYRO U35-511 e -e • SENSOR LOCATION SIGNAl TO PURPOSE NUMBER 1 NOSE 1. DOPPLER NAVIGATION ATTITUDE REFERENCE VERTICAL ELEC-GYRO TRONICS 2. COPILOT'S VSI (NORM) ATTITUDE INDICATION ..... \.0 CN-1314/A COM PART -...J MENT 3. PILOT'S VSI (ALT) AND a . ATTITUDE INDICATION LEFT HAND COMMAND INST b. ATTITUDE REFERENCE SIDE SYSTEM (ALT) 4. SAS/FPS COMPUTER a. ROLL ATTITUDE HOLD b. SAS 2 COMPARE (RATE) . c. TURN COORDINATION AFCS ROLL SENSOR -NUMBER 1 GYRO U35-510 SENSOR LOCATION SIGNAl TO PURPOSE ~ \.0 co NUMBER 2 NOSE 1. PILOT'S VSI (NORM) a. ATTITUDE INDICATION · VERTICAL ELEC-AND COMMAND INST. b. ATTITUDE REFERENCE GYRO TRONIC.S SYSTEM (NORM) CN-1314/A COM PART- MENT 2. COPILOT'S VSI (ALT) ATTITUDE INDICATION RIGHT HAND SIDE 3. SAS/FPS COMPUTER PITCH ATTITUDE COMPARE AFCS PITCH SENSOR -NUMBER 2 GYRO U35-509 e e • - e e SENSOR LOCATION SIGNAL TO PURPOSE NUMBER 1 NOSE 1. DOPPLER NAVIGATION ATTITUDE REFERENCE VERTICAL ELEC ...... GYRO TRONICS 2. COPI LOT'S VSI (NORM) ATTITUDE INDICATION "' "' CN-1314/A COM PART-MENT 3. PILOT'S VSI (All) AND a. ATTITUDE INDICATION lEFT HAND COMMAND INST b. ATTITUDE REFERENCE SIDE SYSTEM (AlT) 4. SAS/FPS COMPUTER a. PITCH ATTITUDE HOLD b. PITCH BIAS ACTUATOR c. SAS 2 COMPARE (RATE) AFCS PITCH SENSOR -NUMBER 1 GYRO U35-508 NOTES 200 AS I GAIN J.: ' WITH SAS 2 ON ~-4 MA/V VALVE YAW RATE ------1 80 MV/ •/SEC WASHOUT RATE { SAS I GAIN . ].: WITH SAS 2 OFF ~8 MA/ V YAW SAS I YAW SAS2 SERVO SERVO Ll~~r.ER I VALVE, l: 0E +AND-REFERENCE 51, SAS/ FPS COMPUTER LAGGED '-- RATE ~ 1.4 MAIV IF SAS SAS I ENGAGE NO.I LATERAL 2 IS ON ~ACCELEROMETER I -FROM NO.I 2.8 MA/V IFSAS ,....: STA81LATOR 2 IS OFF AMPLIFIER FILTER 8-- ANY(OF 4) PEDAL r SWITCHES ACTIVATED AND AIRSPEED r>60 KTS (NO.I) TRANSDUCER YAW TRIM ACTUATOR YAWSAS PEDAL ____.J @----o (. ACTUATOR 0 TRIM FPS 1 +AND-10% c._,. YAW BOOST TAIL ROTOR TAIL ROTOR SZRVO SERVO PITCH I / _..... CEILING I COPILOT 7 / / //~ ( P_1_DAL } INO .I (LH) ~SWITCH LATERAL SAS I AMPLIFIER ...L___..l.._ ACCELEROMETER ' --...,....,...-------------- YAW RATE GYRO AERODYNAMIC COUPLING SAS I YAW AXIS BLOCK DIAGRAM BH-372 35K8037 201 AS I GAIN WITH SAS2 ON SASI GAIN { WITH SAS 2 OFF LIMITED ROLL PROPORTIONAL ROLL TRIM ACTUATOR CYCLIC ROLL f.:\_ __ _ ~(j TRIM/FPS & (_,"'."""'- /--~CEILING //:/ / ~CYCLIC STICK PILOTS/ RH ROLL AKIS VERTICAL ( \. G'f'_R()_ _10.4 MA/V RATE . ~9 MA/V PROPORTIONAL Soa MA/V RATE ~. 8 MAl V PROPORTIONAL SAS I ENGAGE ~TORQUE TUBE" ROLL ROLL SASI SAS2 SERVO SERVO VALVE +ANDs•t. ~ ~- REFERENCE SAS/FPS COMPUTER . HYDRAULIC COUPLING 2ND STAGE ( A I < ' l•l l•l £ > < ROTATING : SWASHPLATE STATIONARY SWASHPLATE ROLL SAS ACTUATOR +AND-10% AERODYNAMIC COUPLING BH-370 SAS I ROLL AXIS BLOCK DIAGRAM 351<8035 202 AS I GAIN WITH SAS 2 ON {o.4 MA/V PITCH RATE 125MV/ •/SEC ~HOUT I • I RATE { SAS I GAIN WITH SAS 2 OFF {o.eMA/V PITCH SASI PITCH SAS 2 LIMITER I 4MA SERVO VALVE I SERVO VALVE LAGGED RATE +AND5')(, SASIENGAGE ~E REFERENCE SAS/FPS COMPUTER TIP PATH ROTATING SWASHPLATE STATIONARY SWASHPLATE / / ~-" / (, ____ /~-/+---4----+ ~ CYCLIC STICK ~ PITCH AXIS ------- CEILING COPILOT. 0 PITCH RATE ELECTRICAL SIGNAL PITCH RATE GYRO AERODYNAMIC COUPLING SAS I PITCH AXIS BLOCK DIAGRAM BH-371 351<8033 203 NOTES 204 DEPARTMENT OF AVIATION SUBJECTS SYSTEMS TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama May 1984 File No. 47-4748-5 LESSON EVALUATION AFCS Complete the following questions by filling in the blanks. 1. AFCS enhances the _ ---··------------ ·-and ________ _ qualities of the helicopter. t. The five on/off switches located on the AFCS control panel are ·------ ---------·-·-------··-__________, ---------' and 3. The minimum functions that must be activated for the proper FPS operation are -·--- ------ _ _________, _______ and/or ______ and 4. At airspeeds below------ -----' the yaw axis of FPS will provide ~. At airspeeds above , the yaw axis of FPS will provide ------a-n-35 % 228 DEPARTMENT OF AVIATION SUBJECTS SYSTEMS TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama May 1984 File No. 47-4741-3 LESSON EVALUATION ELECTRICAL SYSTEM Complete the following questions by filling in the blanks. 1. When the APU generator is the only source of AC power, all equipment may be operated with the exception of the and at the same time. 2. In the event one main generator fails, the other main generator is capable of supplying the helicopter AC power requirements. 3 • . Three positions of the main generator switches are ____________ _____,and I 4. If a main generator becomes disabled or disconnected from its load, the control switch is manually placed to I and then back to the ______ position in an attempt to reset. 5. The battery can provide ______ APU starts at 35 percent capacity. 6. The provides voltage regulation over and under voltage protection, over and under frequency protection, and feeder fault protection. 7. Three positions of the external power switch are ______, and _____ 8. The position of the main generator control switch provides a check to the generator AC output without connecting the generator into the helicopter bus system. 9. The unit that restores the battery charge and determines the condition of the battery is the _________!..I_______ 10. The two 200-ampere ACIDC convertors convert voltage to -----voltage. 229 List the pilot action when the #1 and #2 CONVERTER caution l i ghts are e on. a. b. I List the pilot action when the #1 and #2 CONVERTER and AC ESS BUS OFF caution lights are on in flight. a. b. C·. Caution lights remain lit. d. e. f. g. 11. i . 23 0 POI FILE: 47-4742-3 AUXILIARY EQUIPMENT TYPE OF INSTRUCTION: PEACETIME HOURS MOB Ill ZATION c 3.0 3.0 LEARNING OBJECTIVE: Given nine written questions containing descriptive situations or operating conditions, from memory, the student will write IAW TM 55-1520-237-10 the- a • . Four functions of the APU. b. Two APU controls located on the overhead console. c. CAUTION to be observed during APU coastdown. d. Minimum APU accumulator pressure for starting the APU. e. Two methods normally used in recharging the APU accumulator. f. Position of the AIR SOURCE HEAT/START, HEATER, and VENT BLOWER switches for maximum heat, utilizing the engines as the air source. g. Switch that sends test signals through the fire detection system to put on the FIRE WARNING lights and verify system operation but not including the photo cells. h. Number of FIRE WARNING lights on with TEST switch in the number 1 position. i • Number of FIRE WARNING lights on with TEST switch in the number 2 position. j. Procedure for arming the FIRE EXTINGUISHER LOGIC MODULE. k. Positions on the FIRE EXTINGUISHER switch. l. Caution to be observed when using the windshield wipers. m. CAUTION to be observed for windshield anti-ice when flying into an icing condition. n. Positions of the cargo hook release (EMER RLSE) switch that permits checking the. emergency circuits. o. Location of the cargo hook emergency release (EMR/REL) switches. p. Procedure for normal release of the cargo hook from the cockpit with the controller (CONTR CKPT ALL) switch in CKPT position. q. Cockpit indication when the cargo hook is in the OPEN position. 231 r. NOTE to be observed to prevent unintentional discharge of the cargo ~ hook explosive cartridge when making the emergency release test. STANDARD: Nine of nine questions must be answered correctly to sat isfactorily complete the objective. LESSON REFERENCE: TM 55-1520-237-10, chaps 2, 4, and 8; student handbook. TASK: This objective supports tasks 03-1402, 10-1501, 1502, 4012, 4043, 6011, and 6501. EQUIPMENT: UH-60 fire detection, fire extinguisher trainer, FR 1344; UH-60 caution/advisory trainer, FR 1330. INSTRUCTIONAL ELEMENT: DOAS, STD, ASB. • 232 APU LEADING PARTICULARS Rated Engine Speed t·1aximum Steady State Exhaust Gas Temperature \~eight {dry) Output Shaft Horsepower Fuel Consumption at rated pm<~er Reduction Gear and Accessories Input speed {rated) Output speed {rated) Fuel Control Assembly Fuel Lubricating Oil Oil Pump Press at rated speed Oil Sump capacity Components and Systems Compressor Turbine Combustor 61,565 RPM {monitored for control, but not rea dab1 e) 6490 c 92 lbs. {41.73 Kg) hp 90 shp {45Kw) @ Zero Bleed 40 shp {30Kw) @ 61 ppm Bleed 150 pph {approx) {from No. 1 Fuel Tank only) {68 Kg hr) 61,565 RPt~ 12,000 RPM {APU AC generator drive) 4,235 RPM JP-5, JP -4 MIL-L-23699 or MIL-L-7808 15-4u psi {1.0-2.8 Kg/cm2) Monitored f or lo w pressure but not readable) 3 q t s • { 2 • 7 1 i te r s ) Single-stage, centrifugal fl O'IJ Single-stage, radial-inflow Annu 1 a r type 233 BATTERY BUS OR , ... I BATTERY UTILITY BUS LOWER CONSOLE APU ELECTRONIC UPPER CONSOLELEGEND SEQUENCING UNIT BACKUP __ ELECTRICAL ----111( ACCUMULATOR HYDRAULIC ~ I WIRING SYSTEM ~ELECTRICAL CONTROL --HYDRAULIC Jll':.o ~FUEL CAUTION/ADVISORY PANEL FUEL PRIME APU SHUTOFF VALVE APU SYSTEM • SIMPLIFIED BLOCK U35-170 DIAGRAM ---e - - - AUXILIARY POWER UNIT START SYSTEM The hydraulic accumulator and hand pump, in the aft cabin ceiling, provide the hydraulic pre~sure for driving the APU starter. If the APU fails to start, the hydraulic accumulator can be recharged by the hydraulic hand pump. When the APU is operating, the AC generator mounted on the APU provides electrical power to the helicopter systems. The hydraulic utility module and backup pump on the left forward deck within the main rotor pylon will automatically recharge the depleted hydraulic accumulator for the next APU start. The APU controls are in the cockpit on the upper console. Indicator lights on the caution/advisory panel provide cockpit monitoring of the APU system. 235 --STARTER OFF N w 0'1 START AUTOMATICALLY, \BORTED IF E.G.T. <50°F -.... MAIN FUEL /ON (14%) DISABLE START CKT >14% IGNITION ON START FUEL ON ARM 40 SEC TIMER APU %_RPM E.G.T. LIMIT 1375°F · E.G.T· :,:-:·;-;,.,,__ ) Ll MIT ~~L..., ~~il~:l~ 100 OVERSPEED SHUTDOWN IGNITION OFF START FUEL OFF ACTIVATE LOW OIL PRESSURE CKT. ARM 1.5 SEC. TIMER DISABLE SEQUENCE FAIL (40 SEC. TIMER) ACTIVATE UNDERSPEED CKT. (90% + 1.5 SEC.) II II APU ON MAX FUEL ON LOADED SPEED UNLOADED SPEED (103%) U35-428 APU SEQUENCE DIAGRAM 67T3058 -- e APU START SEQUENCE During a normal APU start, ignition, fuel flow, precent APU RPM, acceleration, exhaust gas temperature, starter engagement/disengagement; and oil pressure overspeed must follow a timed sequence of events. The electronic sequence unit (ESU) monitors and controls these events during start and operation. If a start sequence fails, then the ESU will automatically shut down the APU. AUXILIARY POWER UNIT BUILT-IN TEST EQUIPMENT An indicator panel in the cabin monitors APU faults and will indicate the reason for APU shutdown on built-in test equipment (BITE) indicators. The BITE indicators are incorporated in the APU electrical sequence unit (ESU) and will indicate 13 distinct reasons for APU shutdown. Those indicators can be monitored during APU operation without interrupting normal operating systems. During a start, each major sequence step will have a visual indication of go/no-go. During operation; major predetermined parameters (exhaust temperature, turbine speed, and oil pressure) are sampled at least every 50 milliseconds. If any one of the predetermined values is exceeded, the APU shuts down and appropriate BITE indication is made. 237 0 0 I #l FUEL LOW Ill GEN #2 GEN #2 FUEL LOW J II -I #1 FUEL c #1 GEN BRG #2 GEN BRG #2 FUEL PRESS PRESS I II I #1 ENGINE A Ill CONV #2 CONV #2 ENGINE OIL PRESS OIL PRESS I u #l ENGINE AC ESS DC ESS #2 ENGINEOIL TEMP BUS OFF BUS OFF OIL TEMP I T CHIP BATT LOW BATTERY CHIP #l ENGINE CHARGE FAULT #2 ENGINE iI I #1 FUEL GUST PITCH BIAS 112 FUELFLTR BYPASS LOCK FAIL FLTR BYPASS 0 Ill ENGINE Ill OIL #2 OIL #2 ENGINE STARTER FLTR BYPASS FLTR BYPASS STARTER N Ill PRJ Ill HYD 112 HYD 112 PRJ SERVO PRESS PUMP PUMP SERVO PRESS II TAIL ROTClR IRCH Ill TAIL RTRQUADRANT I NDP II A SERVO MAIN XMSN INT XMSN TAIL XMSN OIL TEMP OIL TEMP OIL TEMP BOOST SERVO STABILATOR OFF SAS OFF TRIM FAIL II II I vD 0 I LFT PITOT FLTPATH 0 I RT PITOTHEAT STAB IFF HEAT II II I CHIP INPUT CHIP CHIP CHIP INPUT ll s - MOL -LH INT XMSN_. TAIL XMSN MDL -RH c I CHIP ACCESSA CHIP MAIN APU ~ CHIP ACCESS 0 uT MDL -LH MDL SUMP ~ FAIL __ MOL -RH R MR DE-ICE TR DE -ICE 0 ICE N FAULT FAIL DETECTED I II II y MAIN XMSN LIll RSVR 112 RSVR BACK -UP OIL PRESS LOW LOW RSVR LOW II Ill ENG [ NG. INLET I #2 ENG INLET 112 ENG p ANTI-ICE ON -ICE ON ANTI -ICE ON ANTI -ICE ON A D PRIME BOOST v wAPU ON~~APU GEN ~ PUMP ON sI A 0 gAPUL~~UM~I wow LDGH ON #2 TAIL RTR N R SERVO ON y EBRTIOIM CARGO HOOK ARMED HOOK OPEN I PARKING EXT PWR @ L BRAKE ON CONNECTED TEST I 0 U35-377 '238 • HEATER/VENTILATION SYSTEM The system consists of a heated air source, cold air source, m~x~ng unit, temperature sensing unit, overtemperature sensor, controls, ducting, and registers. The heating system is a bleed air system with bleed air supplied in flight by the main engines or the APU. The external air connector allows use of an external ground source which can also provide heat. The helicopter cabin and cockpit are ventilated by an electrically operated blower motor. The VENT BLOWER control switch is marked OFF and ON. When ON, the blower pulls in ambient air, forcing it through a flapper-type check valve into the caoin ducts. The blower is not used for heater operation. Ram air vents, for cooling the cockpit area, are on each side of the upper console and at the front of the lower console. NOTES: 239 . . . . . ·. "'·..... ~..--...._ z z - . . .. . . . •········... ~..... ·.. ·.. . \. .·... -"' -r...... ........-. ... ... . •·... ·.... ·~ I ~-..'-, ·. : ~ ,....:... ' ... •. ~ ·....:...._. ; ...... . .. -..-..~~.' ' •, . , ...................._ , -;-___..,. ~ ~ ~ ~ ~ .... ... .. . ... . .. .. . . . " . .-. ·.. ~ . ... 1\\ \ r ~~ •• . .. . ......................... . .\ ·. ····.·...................,~-\~,... . .•:··::................~ ............-..-.,."' ~ : ., .. ·~~·~·~~;~~;~·~ ' .," .. . . ----.. ~ '~-.. ... ·~ ,....... ·~ \ '-.._[ ~ . ·. •. ., •.. . . . . . . ... , . . .. . ~ . ,.., 'I I ,_ \ \ 1'1\ \\\-'.a\ .:, '.. . . . •... "':. ..-11-......_ """: ,. ' .........___.,-: ,, ~ --~~. . , ' ,/ :. .: ,, .' ,, ,' ; ; ,, , ' ,. , ' , I I I I I I I I 1 1 I I I II : I ~ II : ;; ... ... .......... . . ~---..·· --.. -~---· 0 - 1 c ..I ..I c 1 en z - 0 1-... N N (.) 1::: ::::» 10 Q D:: - c D:: Ill ... c Ill z e .- N -'="" 1-' I SHUTOFF VALVE REGULATION VALVE BELLOWS ~ ~ AMBIENT AIR - OVERTEMP ~~.,~. ~COCKPIT SENSOR ~·I·Efil CONTROL \ -$$$$$$$$$$$~-- TO CABIN=> MIXER VALVE (MOVABLE) HEATER OPERATION SCHEMATIC u3s-166 67T2026 FIRE DETECTION SYSTEM The fire detection system provides fire warning to the cockpit in case of fire in either main engine compartment or in the APU combustor compartment. The system consists of five light (not heat) flame detectors, control amplifiers,and a test panel. The detectors are mounted strategically in the potentialfire zones. Two detectors are installed in each main engine compartment. Master fire lights are on the glare shield located in front of the pilot and copilot. Both will illuminate for any fire or test condition. A fire control handle is located on the left and r ight side of t he engine quadrant for the number one and number two engine compartments. Pulling back on th~ f i re control handle (T handle) shuts of f fuel and arms the logic module for use of the fire extinguishing bottles for that system. A fire control handle for the APU is located on the upper console. Pulling out on the APU fire control handle shuts off fuel and ignition and arms the logic module for use of the fire extinguishing system for the APU. A system TEST switch is located on the upper console. The switch is used to test system electrical circuits. NOTES: 242 FIRE EXTINGUISHING SYSTEM The MAIN and RESERVE engine fire bottles are located in the main rotor pylonaft of the main gearbox. Each fire bottle is fill~d with 2.5 lbs (1.13Kg) of liquid CF3Br (monobromotrifluoromethane) and charged with nitrogen pressurized at 600 psi (42 Kg/cm2). The control circuitry is designed so one or both bottles may be used on one or both main engines or the APU compartment. Each bottle has a thermal discharge valve that will automatically open if the bottle pressure builds up excessively. If ~he aircraft should crash, an omnidirectional inertia switch (impactswitch) will cause a fire bottle to fire into each main engine compartment. NOTES: 243 LOWER CONSOLE LEGEND OUTPUT CIRCUIT INPUT CIRCUIT ENGINE #2 DECK UPPER CONSOLE DETECTOR 0. 1 ENG NO. 2 ENG FIREWALl ---rll ANDLE IIIIHANDLE MER OFF "T" rf ""')), EMER OFF "T" FIREWALL N DETECTOR .p.. .p.. FIREWALL DETECTOR ~~~ ~ FIREWALL DECK DETECTOR ENGINE #1 FIRE DETECTOR SYSTEM-FUNCTIONAL U35·169 BLOCK DIAGRAM 67T3030 e e -e THERMAL RELIEF OUTLET NO.2 ENG _/--\ ~~ ~-.// ) ~ ~-~ ~-?! DIRECTIONAL ·· : -· : ~-.u CONTROL --~--_) / ', ,-4 N .j> -~ .r::;" __, t--/ \.{_.-VALVE J ~Fri___J../ --<~-;~\""APU I.Jl • ,-f "" --'\.·) . / __./ ,/ ;' ,___ . -- _,--/, c _____/ ~---~ / "->~ _.r' ( }"l ~·;\": ...r< ~'-~··"'·· ) ~~--/ "'-No.1 ENG ,• ~ \_.'. _/'--J ~ ..........-"' ... FIRE EXTINGUISHING SYSTEM U35-160 INSTALLATION &n30l7 WIPER MOTOR ~C)_ / , ' ,,. 0 ....,,..::-.. ..... ,/ ,,'' .............. :' I •' .. , _j /<:)_~ COUPLING// / N p.. ~ FLEX SHAFT I ·~ J \ ,,,WINDSHIELD WIPER \ ,,'',,,'' ' ' "·, ,' ' ...... ---_____,,,,,""CONVERTER ----~ WIPER MOTOR· FLEX SHAFT WINDSHIELD WIPER SYSTEM U35-109 LOCATION DIAGRAM &n2ool --e NOTES 247 WINDSHIELD ANTI-ICE SYSTEM The windshield anti-ice system prevents the format ion of ice on the pilot's and copilot's windshields. This is done by applying three-phase 115 VAC power to heater elements within the windshield. This maintains a temperature of 70oF (210C) to 115oF (46°C) on the windshield surface. The AC power is supplied to the windshield heater elements by anti-ice controllers that are activated by two temperature sensors within each windshield. NOTES: 248 e _,..·.-" .:;;.-··:~-•'\\·;<---'\\\ \\' ' 1.0 ·~ .. .0 ·-. ·········· ..· ·-..... ... ... . .. ........:·...... ..:::::/ /::::...-··········; /::..-····· \ \'}\ \\ \ \ \\ \ J\ • • 0 •• • • Q, '• '• "' '" ' 'I ' '• , l. ··--...___ -..___ :-~·',:::·:::~:-~~~:~(.:"f,·..P/f:!"~:::i:'_..J_u'~\~~~!_,>::;~::-~~~:::::~{;c'~~--· \~ ':~··. ··<:·.--.. ····<:::::::::::.. .....--NO.1 JUNCTION BOX COPILOT'S WINDSHIELD WINDSHIELD ANTI-ICE SYSTEM LOCATION DIAGRAM U35-435 61F6058 CARGO HOOK SYSTEM The cargo hook system consists of an 8,000 pound (3,628 Kg) capacity hook, electrical control circuits, and two advisory lights, (HOOK ARMED, CARGO HOOK O~EN) on the caution/advisory panel. The hook is located in the cargo hook well underneath the cabin floor. The electrical controls of the cargo hook system consist of the following: a CARGO HOOK ARMING switch labeled SAFE and ARMED; a CARGO HOOK CONTR switch labeled CKPT and ALL; CARGO HOOK EMER RLSE switch labeled NORM, OPEN, and SHORT; a TEST light; CARGO HOOK BEL switches on the pilot's and copilot's cyclic stick; an EMERGENCY HOOK BEL button on the pilot's and copilot's collective sticks; and a crewman's control pendant with NORMAL and EMERGENCY RELEASE switch and button. The system has three modes of load release: (1) a normal release system, (2) a manual release system, and (3) an emergency release system. The cargo hook contains ~n explosive cartridge known as a squib cartridge, which is used to release the load during emergency conditions. When 28 VDC is supplied to the cartridge it explodes, driving a piston and pin inside the squio at the hook into the load arm lock to release the hook. The load beam will drop the load, and the CARGO HOOK OPEN light will stay on until the old squib has been replaced. NOTES: 250 COVER MANUAL (EXPLOSIVE CARTRIDGE RELEASE /COVER) LEVER LOAD BEAM CONFIGURATION B EXPLOSIVE CARTRIDGE MANUAL RELEASE KNOB LOAD BEAM CONFIGURATION A CARGO HOOK (TURNED 180°FOR CLARITY) 251 COPILOT'S COLLECTIVE \·.....· ·\~: PI LOT'S STICK · · <~. :· CYCLIC STICK COLLECTIVE STICK GRIP N Ul I ------0 I N CARGO HOOK I 0 I I 0 I I 0 I CYCLIC STICK GRIP UPPER CONSOLE CAUTION/ADVISORY PANEL CARGO 0 0 0 0 0 REL. 1 ~ CARGO HOOK ~ :!llflll~i~:llll1\!llii~;,f;,.. . SWITCH ~ EMER RlSE CONTR ARMINd Q, ~li ~ ~ SHORT All ARMED ~-:·x·: · : ·:·:·:·:·:·: ·:·:·:·:·:·:·:·:·: · :· : ·: ·: ·:·:·::::: : :: :·:·:·:·:·:·:·:·:·:·: ·: CARGO HOOK SYSTEM LOCATION U35-411 DIAGRAM 67T2112 • e - -~- AUXILARY EQUIPMENT LEFT RIGHT SIDE FILE NO 47/48-4742-5 SIDE Sl AIRSPEED S2 TRANSDUCER COPILOT PILOT ~------------------------------------, Ln N I • I AIRSPEED AIRSPEED I w I : ~ ~: I VERTICAL SPEED INDICATOw~ I ALTIMETER I • I L-------------------------------~ • AIR DATA TRllt-sDUCER TO AFCS & CIS ~ VSI VERTICAL SITUATION INDICATOR HSI HORIZONTAL SITUATION INDICATOR PITOT HEAT SWITCH #2 PRI AC BUS SA #1 PRI AC BUS SA #1 PRI DC BUS LFT PITOT ....___ r----Ill RT PITOT HEAT 11 " ' HEAT N A2 Al·· Al A2 U1 ~ X2 Xl Xl X2 X4 X3 X3 X4 - - - - PITOT STATIC HEATER SYSTEM SCHEMATIC U35-110 I8FIOI5 DEPARTMENT OF AVIATION SUBJECTS SYSTEMS TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama May 1984 File No. 47-4750-3 LESSON EVALUATION AUXILIARY EQUIPMENT Complete the following questions by filling in the blanks. 1. List the other three functions the APU provides. a. b. c. d. Emergency in flight electr ical operations. 2. The two APU controls located on the overhead console are the APU and APU handle. 3. Cycling of the BATT switch during APU coastdown may cuase possible or shutdown. 4. The APU accumulator pressure m st be a minimum of psi for starting the APU. 5. The pump and the pump are the two normal methods of recharging the APU accumulator. 6. Utilizing the engines as an ai r source, for maximum heat the AIR SOURCE HEAT/START switch must be in t he position, the HEATER switch must be in the position, and the VENT BLOWER switch must be in the position. 7. The switch sends test signals through the fire detection system to put on the FIRE WARNING light and verify proper operation to, but not including, the photo cells. 8. The number of FIRE WARNING lights that will be illuminated is when the FIRE DET TEST switch is in the NO. 1 position. 9. The number of FIRE WARNING lights that will be illuminated is when the FIRE DET TEST switch is in the NO. 2 position. ------- 255 10. The FIRE EXTINGUISHER LOGIC MODULE is armed by -----the 11 T11 handles. 11. The positions of the FIRE EXTINGUISHER switch are ' and 12. To prevent possible damage to windsh i eld surface, do not operate wipers on windshields. 13. Windshield anti-ice should be turned on prior to entering icing conditions to prevent which can cause engine 14. The cargo hook EMER RLSE test switch has three positions. The emergency re l ease circuit is.tested by placing the switch in the position and the position. ---- 15. The EMER HOOK RLSE switches are located on the pilot's and copilot'ssticks and the crew members' 16. When the cargo hook control switch marked 11 CKPT/ALL 11 is placed in the CKPT position, the load will be normally released by pressing the ---:-;--..,......----------,--:--o switch 1 ocated on the pilot's and copilot s sticks. 17. When the cargo hook is in the OPEN position, the cockpit indication will be illumination of an light marked 18. To prevent unintentional discharge of the cargo hook explosive cartridge, the pilot shall of the emergencyrelease circuit test before that is done. 256 POI FILE: 47-4745-1.5 POWER TRAIN SYSTEM ~ . TYPE OF INSTRUCTION: PEACETIME HOURS MOBILIZATION 1.5 1.5 c LEARNING OBJECTIVE: Given four written questions containing descriptive situations or operating conditions, from memory, the student will write lAW TM 55-1520-237-10 the- a. Special feature of the power train chip detect system. b. Two purposes of the intermediate and tail gearbox chip/temperature system. c. Effects of excessive temperature over the intermediate and tail gearbox chip detect fuzz burner. d. Operating limitations of the main transmission. e. Minimum transmission oil pressure for continuous flight. f. Maximum time limit to operate the transmission with an oil temperature above 120°C. g. Location of the main transmission oil temperature warning system sensor. STANDARD: Four of four questions must be answered correctly to satis factorily complete this objective. LESSON REFERENCE: TM 55-1520-237-10, chaps 2, 5, and 9; student handbook. TASK: This objective supports tasks 03-1401, 1402, 10-1501, 1502, 4006, 4060, and 6501. EQUIPMEN; : ~H-60 power train system trainer, DVC 1-118; UH-60 composite trainer; UH-60 caution/advisory panel, FR 1330. INSTRUCTIONAL ELEMENT: DOAS, STD, ASB. 257 POWER TRAIN The primary function of the power train system is to take combined power from - two engines and transfer it to the main transmission, main rotor, and tailrotor. The secondary function is to provide a mec hanical drive for electricaland hydraulic accessories. The power train consists of two high speed engine drive shafts, a main trans mission, an oil cooler, six tail rotor drive shafts, an intermediate gearbox, and a tail gearbox. The tail rotor drive shafts are mounted with vi cous dampened bearings so that drive shaft alignment is not required following shaft replacement. The oil cooler drive is an integral part of the tail rotor drive shaft system. The intermediate gearbox, located at the base of the tail pylon, changes driveshaft angle of drive 58 degrees and reduces tail drive shaft speed.Lubrication of the gears and bearings is accomplished by .an integral splash oiling system. The tail gearbox supports and drives the tail rotor at a reduced RPM of approximately 1190 RPM. A splash oiling system also provides tail gearboxlubrication. The tail gearbox assembly houses a two-stage hydraulicallypowered servo that is used to change pitch of tail rotor blades for aircraft yaw control. A gust lock prevents the blades from rotating when the helicopter is parked. The gust lock system consists of a locking handle at the rear of the cabin, aGUST LOCK caution light on the caution/advisory panel, a locking device, andteeth on the tail takeoff flange. The lock reacts wind loads on the mainrotor blades and torque from the main rotor shaft during blade fold. The ·lock should be applied or released when the rotor system is stationary. NOTES: 258 TAIL GEAR BOX-~ TAIL ROTOR DRIVE SHAFT OILCOOLER \ N Vl 1.0 INTERMEDIATE GEAR BOX MAIN TRANSMISSION TRANSMISSION SYSTEM POWERTRAIN U35-213 6804007 MAIN TRANSMISSION The main transmiss£on assembly consists of five mo~les: two input modules, two accessory modules and one main module. The transmission assembly receives power from both engines through the input modules and provides gear reduction within the main transmission to change direction of rotation to decrease RPM, and to change angle of drive to the main rotor. The main transmission is mounted with a 3 degrees forward tilt on top of the cabin fuselage between station 320 and station 360. The input modules change the angle of drive from the engine drive shafts to the main transmission. The accessory modules provide drive for two generators, two hydraulic pumps, and a mount for the rotor speed sensor. NOTES: 260 INPUT MODULE N ()"\ ,..... ~')'' CJ HYDRAULIC MODULE Ill IT~ ~ACCESSORY MODULE MAIN TRANSMISSION ASSEMBLY U35-211 6804008 ~ N SPEED SENSOR -~-~~ (~IGHT MODULE) ~ ~ACCESSORY ()"'~~ MODULE N LOW OIL ~ PRESSURE SWITCH LE) ......______ (LEFT MODU ~~-~ ~ CHIP ~ DETECTOR --------------~ U35-230 ACCESSORY MODULE EXTERNAL VIEW 6104011 e e • - TAIL TAKE OFF ENGINE INPUT !'..) 0'\ w STATIONARY RING GEAR OIL PUMP U35-129 MAIN MODULE DRIVE SCHEMATIC 6804014 MAIN TRANSMISSION LUBRICATION SYSTEM The transmission incorporates an integral wet sump lubrication system that provides 7.5 gallons of cooled, filtered oil under pressure to lubricate all bearings and gears. The lubrication system consists of two combination pressure and scavenge type pumps operating in parallel, one oil cooler/blower assembly, a two-stage main oil filter·, and internally cored oil passages in the transmission sump housing. The only external oil lines are pressure and return lines to the oil cooler/blower assembly. Oil quantity is measured with an oil level dipstick located at the right rear corner of the transmission. The dipstick ADD mark indicates about 1/2 gallon of oil should be added. .N.Q.I..E.S : ~ 264 LEFT ACCESSORY GENERATORS RIGHT ACCESSORY MODULE'-......_ ~ MODULE CJ;:-s ~I I ...::IT1--~···~ 1/ I ' ' I 18! ....,.....-§ I I RIGHT INPUT LEFT INPUT MODULE i.... I MODULE -~ I ~ - I § ! ,.....,..... I N I 0'> Vl I § I i § •• •• ~_J MAIN MODULE~ I I LEGEND .-.~...-..r-..,...,...-PRESSURE u•••••••••u RETURN I I S S ........,....;..........,...,....-..r'..,...,....-..r'..oll':oi ~.1...--r'.l.l..... SUPPLY MAIN TRANSMISSION LUBRICATION U35-127 BLOCK DIAGRAM 6104019 ,...........~ .---- CAUT ION/ AfNISORY PANEL NR SENOR N "' "' OFF RE~T ! ROTOR OYERSPEEO RESET SWITCH NOSE AVIONICS COMPARTMENT soc ... SOC#Zl,"r-----,------r_J MAIN TRANSMISSION INDICATING SYSTEMS BLOCK DIAGRAM 81·H~90 e • 6104114 - ~ TRANSMISSION CHIP DETECTOR SYSTEM The main transmission chip detector system consists of five magnetic chip detector plugs for the left and right input modules, left and right accessory modules, and the main module. These chip detectors incorporate a "fuzz-burn" feature which will remove small particles of metal debris from the detector contacts. "FUZZ-BURN" eliminates nuisance warning lights caused by normal collections of metal fuzz. There is an associated chip caution light for each module of the main transmission which will illuminate when a large metal chip bridges the gap in the chip' detector. When a chip warning light illuminates, the affected chip detector must be removed and cleaned and electrical power must be cycled off and on to clear the chip caution light. N.QIE.S: 267 NOTES • 268 RIGHT ACCESSORY MODULE RIGHT INPUT MODULE MAIN GEAR BOX DC ESNTL FCH~JET !···-v····· ······~·-··· ·-·~ FROM BUS ~ rL.•.•.•~ TRANSMISSION OIL COOLER GENER~ LfoANIFOLD .-II CHIP ACCESS MDL RH II I ~~ r.-~·~iI ~..oill':. .,,. .,...,...., ~------~~--~----~~~~ ! I ~ I CHIP INPUT MDL RH I LEFT ACCESSORY MODULE LEFT INPUT MODULE - ~ ····~····· ·······~·-··· ...~ J . II CHIP ACCESS MDL fH] I -~ - ~ ~..oill':.--· GENERATOR II II I CHIP INPUT MDLLH CHIP ~ + CHIP DETECTOR DETECTOR TO TRANSMISSION TO TRANSMISSION OIL COOLER OIL COOLER I I ~ \. I I ) Jl CHIP MAIN MDL SUMP II I -=- OIL HOT OIL HOTI CHIP INT XMSN JI I A B A B r---cHIP.TAilXMsN It-----' FUZZ FUZZ BURN t--+-- 0\ FWD, AFT, OR LAT, INPUT ROTATING SCISSORS e U35-299 SWASHPLATE AND UNIBALL CUTAWAY e • 67T4071 e - MAIN ROTOR -M 1~ ~ SPLIT N CONES-------..! "'-J 00 --~~~-SHAFT EXTENSION PITCH CONTROL ROD I MAIN TRANSMISSION SWASHPLATE MAIN ROTOR HEAD AND MAIN U35-237 TRANSMISSION ASSEMBLY 6804041 en 1 z - 0 a. z - A. - e FLAP RESTRAINER EXPANDABLE PIN "" _J 2\ SPINDLE N 00 \0 \ DROOP-RESTRAINER ELASTOMERIC BEARINGS U35-208 SPINDLE MODULE CUTAWAY 6104043 (.) -a:: UJ z ::Ee, 0 .... -::E (1)0:: (/)-< (.) /e~ << 0 ....IUJ A. A. UJal (.) 0 .... -.... (/) < A. I .... c z (/) 0 0 0 0:: 1 c D: Ill A. 0 A.i 0~ 0 A. 1 .... (/) en A. I 0 ' ' ........ ...._ ___, 0 D: ' ...... Q ...... ...... ...... ---1 \ ...... ...... , , ___ _l.-__.1' UJ c < ..... m C") 0 .... ,;, C") :l 290 -(.) - en 0 Q. Q. < ..... a.. - z ~ < z 0 ~ J: z 0 - - ~ en 0 Q. u.l > 0 ~m 291 MAIN ROTOR BLADES Each of the four main rotor blades has a titanium spar for its main structural member. The trailing edge is a one-piece fiberglass pocket internally supported by Nomex honeycomb. The swept-back tips reduce noise level and improve blade performance. The leading edge of each blade has a titanium abrasion strip. The outboard one third of the abrasion strip is covered with a nickel abrasion strip. The blade spar is pressurized with nitrogen to detect cracks. Pressure loss in the spar is indicated through the use of a blade inspection method (BIM) indicator which continously monitors spar pressure. The BIM indicator is installed on each blade at the root end to visually indicate loss of blade spar structural integrity. The blades are attached to the rotor head by two quick-release expandable pins that require no tools to remove or install. All blades can be easily folded to the rear and downward along the tail cone of the aircraft for storage or shipment. NOTES: 292 3 SEGMENT CG OF BLADE TITANIUM ABRASION STRIP CENTER OF GRAVITY NICKEL ABRASION STRIP TIP CAP OVER TITANIUM I ----------------~---------------------~-~-----------~-~-------c;--------- BALANCE TRIM TAB TIE DOWN STRIPES ATTACH POINT BOTTOM ONLY CUFF DE-ICE CONNECTOR ~------------------------.tA I BH-813 UH-60A MAIN ROTOR BLADE 67T~02 293 ABERGLASS SKIN t NING BOLT INBOARD TORQUE ( '-__j -----1 Llwmm~mWrrmtm=m:mmrMmfdl~m~m*mWJ~m~m;tmJ~mmmtm~mmrm*'m~:m:wm;mmm:mWMm~m~:mtrrm~m!tntH!r~~M IP CAP TAlL ROTOR BLADE-FRONT VIEW CUTAWAY BH-742 68G7005 301 NOTES 302 - e DE-ICE HEATER MAT· VIEW B-B SPAR ANTI-ABRASION STRIP/ FIBERGLASS SKIN DE-ICE CONNECTOR POLYURETHANE NICKEL ABRASION STRIP ABRASION STRIP OUTBOARD SIDE C8 w ! r~/ 1 0 w ·A VIEW A-A TIP CAP U35-736 TAIL ROTOR BLADE 117T41H FIBERGLASS SKIN NOM EX HONEYCOMB CORE w 0 ~ ALUMINUM HONEYCOMB ANTI-ABRASION STRIP____., DE-ICE HEATER MAT~~~( FILLER AND LEAD WEIGHTS/ SPAR U35-734 TAIL ROTOR BLADE CONSTRUCTION 68G7004 DEPARTMENT OF AVIATION SUBJECTS SYSTEM TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama May 1984 File No. 47-4746-1.5 LESSON EVALUATION ROTOR SYSTEMS Complete the following questions by filling in the blanks. 1. The gust lock will only be applied when the rotor system is 2. The operating limits with the gust lock engaged are 3. During runup, the droop stops should be out at percent RPMR. 4. During shutdown the droop stops should be in at percent RPMR. 5. What caution should be observed to prevent damage to the main rotor blade tip caps? 6. What action is required if one droop stop fails to seat during shutdown? for --------------- 305 POI FILE: 47-4750-6 MALFUNCTION ANALYSIS TYPE OF INSTRUCTION: PEACETIME HOURS MOBILIZATION PE3 6.0 6.0 LEARNING OBJECTIVES: 1. Given five questions containing pertinent indications and/or descriptive situations or operating co nditions, from memory, the student will write lAW TM 55-1520-237-10 the pilot action for- a. No. 1 PRI SERVO PRESS caution light on in flight . b. No. 1 HYD PUMP caution light ON, BACKUP PUMP ON, advisory light OFF (three steps, backup pump on, advisory light remai ns off) . c. No. 2 HYD PUMP caution light ON, BACKUP PUMP ON, advisory light OFF (three steps, backup pump on, advisory light remai ns off) . d. No. 2 TAIL ROTOR SERVO ON, advisory light goes off after being on. e. BOOST SERVO OFF caution light ON resulting in a high cockpitcontrol force in the pedals {250 pnunds). f. Helicopter exhibits erratic motion without a corresponding SAS FAILURE advisory light indication. g. STABILATOR caution light ON during takeof f (auto mode failure). h. No. 1 GEN caution light on IMC. i. No. 1 GEN BRG caution light comes on in f light. j. BATTERY FAULT caution light illuminates i n flight. STANDARD: Five of five questions must be answered correctly to satisfactorily complete this objective. 2. Give n seven written questions containi ng descriptive situations or operating conditions, from memory, the student wil l write lAW TM55-1520-237-10 the pilot action for- a. Percent TRQ split between engines 1 and 2 (No. 1 engine 100 percentTRQ: TGT 850°C; No. 2 engine 40 percent TRQ) TGT 560°C. b. Percent TRQ split between engines 1 and 2 (No. 1 engine 100 percentTRQ, TGT exceeding 850°C; No. 2 engine TRQ 40 percent, TGT 620°C). c. No. 1 ENGINE OUT warning light illuminates, audible tone over theheadset, NO Ng, Np, and TRQ; TGT above 850°C; No. 2 engine TRQ 20 percent, Ng87 percent, Np 102 percent, Nr 106 percent, TGT 650°C. 306 I d. No. 1 ENGINE OIL PRESSURE caution light illuminates in flight, I. I engine oil pressure indication, ze~o pressure. e. No. 1 fuel pressure caution light on (aircraft fuel system modified). f. Ground operations APU OIL TEMP HI caution light illuminates. g. Ground operations APU FIRE "T" HANDLE and MASTER WARNING FIRE lights illuminate (battery source of DC power). h. In flight No. 2 ENGINE FIRE "T" HANDLE and MASTER WARNING FIRE lights illuminate. i. Ground operations cargo hook emergency release (switch openposition) test lamp fails to illuminate at pilot test. j. In flight MAIN XMSN OIL PRESS caution light illuminates XMSN oil pressure low. k. In flight CHIP INPUT MOL LH caution light illuminates, high squealnoise and vibration is noted. 1. TAIL ROTOR QUADRANT caution light on both tail rotor cable broken. m. PEDAL BIND with BOOST OFF caution light on. n. NOTE to be observed after the emergency hook release has been used. STANDARD: Seven of seven questions must be answered correctly to satis factorily .complete this objective. LESSON REFERE]CE: TM 55-1520-237-10, chaps 2, 4, and 9; student handbook. TASK: This objective supports ta~ks 03-1402 and 4010. EQUIPMENT: UH-60 T700 engine simulator, OVC; UH-60 fire detection, fire extinguishing trainer, FR 1344; UH-60 caution/advisory panel, FR 1330. INSTRUCTIONAL ELEMENT: DOAS, STD, ASB. 307 NOTES 308 DEPARTMENT OF AVIATON SUBJECTS SYSTEM TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama May 1984 File No. 47-4750-6 STUDENT OUTLINE MALFUNCTION ANALYSIS 1. Master warning panel. a. #1 or #2 engine out. 309 b. Fire (engine fire). c. Low rotor RPM (dual engine failure). • 310 2. Caution/advisory panel. a. Advisory lights. (1} #1 and/or #2 engine antiice on. (2} #1 and #2 engine inlet antiice on. (3} APU on. (4} APU accumulator low. (5} APU generator on. (6} Prime boost pump on. 311 (7) Backup pump on. (8) LOG LT on. (9) #2 tail rotor servo. (10) Cargo hook open. (11) Hook armed. (12) Parking brake on. (13) External power connected. (14) Weight on wheels (WOW) (added light). -312 b. Caution lights. (1) Engine and related systems malfunction. (a) Ill and 112 fuel low. 1. 2. 3. (b) lfl or 112 fuel pressure. 1. 2. 3. (c) Ill or 112 engine oil pressure. (d) Ill or 112 engine oil temperature. (e) Ill or 112 engine chip. (f) Ill or 112 fuel filter bypass. (g) lfl or 112 engine starter. (h) Ill or 112 engine oil filter bypass. (2) Hydraulic and flight control system malfunction. (a) lfl or 112 primary servo pressure. (b) Ill or 112 hydraulic pump. 1. Ill hydraulic pump failure. a. Backup pump on. Advisory light on. b. Backup pump on. Advisory light off. 2. 112 hydraulic pump failure. a. Backup pump on. Advisory light on. b. Backup pump on. Advisory light off. 313 3. #1 and #2 hydraulic pump failure. Backup pump on. Advisory light on. (c) Tail rotor quadrant. 1. One cable broken. 2. Both cables broken. (d) 1st stage tail rotor servo (jam). (e) Boost servo off. (f) SAS OFF. (g) #1 reservoir low. 1. Leak in the first tail rotor servo system. 2. Leak in the #1 hydraulic pump. 3. Leak in the #1 primary servo system. (h) #2 reservoir low. 1. Leak in the pilot assist system. 2. Leak in the #2 hydraulic pump. 3. Leak in the #2 primary servo system. c. Stabilator malfunctions. (1) Auto mode. (2) Stuck stabilator. d. SAS malfunctions. (1) SAS 1. (2) SAS 2. e. Trim/FPS malfunction. (1) Trim actuator fail. (a) Pitch hardover. ib) Roll hardover. 314 (c) Yaw hardover. (2) Trim act jam (para). f. FPS malfunction. g. PBA malfunction. h. Electrical systems malfunction. (1) AC (a) #1 or #2 generator. (a) VMC. (b) IMC. (b) #1 and #2 generators. (c) #1 or #2 GEN BRG. (2) DC (a) #1 or #2 CONV. (b) #1 and #2 CONV. (c) Battery low charge. 1. 35-40%. 2. Below 35%. (d) Battery fault. i. XMSN--system malfunction. (1) Temperature (high). (2) Pressure (low). (3) Chip. j. INT XMSN or TAIL XMSN gear boxes. (1) Temperature. (2) Chips. 315 k. APU. (1) Oil temperature. (2) APU failure. 1. Pitot heat. m. Main and tail rotor deice. (1) Main rotor deice fault. (2) Main rotor deice fail. (3) Tail rotor deice fail. n. ICM inoperative. o. Auxiliary fuel. P• IFF. 316 DEPARTMENT OF AVIATION SUBJECTS SYSTEMS TRAINING DIVISION UNITED STATES ARMY AVIATION CENTER Fort Rucker, Alabama tttay 1984 File No. 47-475U-6 LESSON EVALUATION MALFUNCTION ANALYSIS Complete -the following questions by filling in the blanks. 1. List the emergency procedures for NO. 1 PRI SERVO PRESS caution light ON in flight. i\. b. 2. List the emergency procedures for NO. 1 HYD PUMP caution light ON, BACKUP PUMP ON, advisory light OFF (three steps, backup pump on advisorylight remains OFF). a. b. Advisory light remains OFF. c. 3. List the emergency procedures for NO. 2 HYD PUMP caution light ON, BACKUP PUf•lP ON, advisory 1 i ght OFF (three steps, backup pump on advisorylight remains OFF). a. b. Advisory light remains OFF. c. 4. List the emergency procedures for NO. 2 TAIL RTR SERVO ON advisory light goes OFF after being ON. a. b. Advisory light remains OFF. c. 5. List the emergency procedures for BOOST SERVO OFF caution light ON, resulting in a high cockpit control force in the pedals (250 pounds). a. b. c. d. 6. List t he emergency procedures when the helicopter exhibits erratic motion without a corresponding SAS FAILURE advisory light indicator. a. b. c. d. 7. List t he emergency procedures for STABILATOR caution light ON duringtakeoff (auto mode failure). a. b. c. ~. List t he emergency procedures for BATTERY FAULT caution light ON in flight. a. Light remains ON. b. 318 9. -List the emergency procedures for NO. 1 GEN caution light ON if IMC. e a. Light remains ON. b. c. d. 10. The pilot action for a percent TRQ split between No. 1 engine and No. 2 engine (No. 1 engine 100 percent TRQ, TGT 850°C; No. 2 engine 20 percent TRQ, TGT 560°C). 11. The pilot action for a percent TRQ split between No. 1 engine and No. 2 engine (No. 1 engine 100 percent TRQ, TGT exceeding 850°C; No. 2 engine 4u percent TRQ, TGT 620°C). 319 12. The component failure and the pilot action for NO. 1 ENG OUT warninglight illuminated, audible tone over headset, no NG, no NP, no TRQ, TGTabove 85U°C; No. 2 engine 20 percent TRQ, 87 percent NG, 102 percent NP, 1U6 percent NR, TGT 55u°C. 13. The pilot action for the NO. 1 ENG OIL PRESS caution light illuminated in flight; engine oil pressure indication is zero pressure. 14. The pilot action for the NO. 1 FUEL PRESS caution light on (aircraft fuel system modified). 15. The pilot action for illumination of the APU OIL TEMP HI caution lightduring ground operations. 320 11 T11 16. The pilot action for illumination of the APU FIRE HANDLE and master warning FIRE light during ground operations. 17. The pilot action for illumination of the NO. 2 ENG FIRE 'T' HANDLE and master warning FIRE light in flight. 18. The pilot action when the EMER RLSE circuit TEST light fails to illuminate with the EMER RLSE switch in the OPEN position during ground operation test. 19. The pilot action for illumination of the MAIN XMSN OIL PRESS caution light in flight, XMSN oil pressure zero. 321 20. The pilot action for illumination of CHIP . INPUT MDL LH caution light inflight, accompanied by high squeal noise and vibrations. 21. The pilot action for illumination of TAIL ROTOR QUADRANT caution lightin flight with complete loss of tail rotor control. 22. The pilot action for a PEDAL BIND with no accompanying caution light inflight; after turning BOOST servo off, bind is not eliminated. 23. The CARGO HOOK OPEN advisory caution light will remain ON and the hook will remain open after the EMER RLSE has been used. The hook cannot beused until is replaced. 322 FLIGHT LINE SUPPLEMENT HANDOUT MATERIAL NOTE: Take this handout with you to the flight line. 323 MIXING UNIT Eliminates inherent control coupling by automatic mixing of the controls. NOTE: Mixing unit is designed to work at designed grass weight of 16,825pounds. Four mixes: 1. Collective to YAW. 2. Collective to pitch or longitudinal. 3. Collective to roll or lateral. 4. Yaw to pitch or longitudinal. See the four mechanical mixes for explanation, next page. -1 I I I_ _j C.oc.t.l!CT1111! 1 N por.5 324 THE FOUR MECHANICAL MIXES 1. COLLECTIVE TO PITCH or Longitudinal Mixing. a. The main rotor forward and aft tilt is adjusted to assist in streamlining the downwash on the stabilator as the collective is moved. b. COUNTERS THE EFFECT OF ROTOR DOWNWASH ON THE STABILATOR. 2. COLLECTIVE TO ROLL or Lateral Mixing. a. Because of translating tendency, a single rotor helicopter has the tendency to translate or draft in the direction of the tail rotor thrust. In the UH-60 this translating tendency is compensated for by the collective to roll mixing, a function of this mixing unit. b. PROVIDES LEFT LATERAL MAIN ROTOR INPUT WITH AN INCREASE IN COLLEC TIVE TO COUNTERACT THE RIGHT ROLLING OR RIGHT DRIFT TENDENCY DUE TO TAIL ROTOR THRUST AND VICE VERSA. 3. COLLECTIVE TO YAW MIXING. a. On conventional helicoptrs, when you increase collective, you must add left pedal to increase the tail rotor thrust to compensate for torque. When you decrease collective, say to enter an autorotation, you must apply right pedal to counter torque. b. In the UH-60 the collective to yaw mix AUTOMATICALLY PROVIDES INCREASED TAIL ROTOR PITCH WITH AN INCREASE IN COLLECTIVE, AND VICE VERSA, and automatically provides a decrease in tail rotor thrust with a decrease in collective. 4. YAW TO PITCH or Longitudinal. a. The tail rotor is tilted 20 degrees and provides 2 1/2 percent lift. b. When a pilot applies left pedal the tail rotor has the tendency to lift the tail resulting in a nose low altitude. c. When a pilot applies right pedal the lift is taken out of the tail rotor resulting in tail low and nose high. THE YAW TO PITCH MIXING COMPENSATES FOR THIS TENDENCY BY AUTOMATICALLY IN THE TAIL ADJUSTING THE FORWARD AND AFT TILT WITH AN INCREASE OR DECREASE ROTOR LIFT TO MAINTAIN THE AIRCRAFT IN A LEVEL ALTITUDE. Remember: The mixing unit is designed to work at designed gross weight of over16,285. If the aircraft is at a lighter weight the mixing unit will compensate, at a heavier weight it will undercompensate. 325 ELECTRICAL MIXING ! I TAtC..~onuc 4Ff&C'T'cv•AJtUs ll'd'ba. I!.~~UT•IfAJU.S \ I i ! . ! ~AM8£ tlEo ~ll{ 0~ ~~IL\MI}£1ltu F-JN MA~. f!. FF~c.T•ve.N ~ss ! i ' e.u~·c,.t. n •)' ,,,Jii 0~ ~UC!\1\~A_I. rw\IICIAIQ I £t.L.cT~Ic_J(t /11 t'tlulr. _IDI, ' IVI£CIIAPICA' HJ'/:/~r, ' MI!C.CANtTAR SEIC.."'O AtA o.. r .. Tl(ttiiSDucatc -'•tiNA£. THE UH-60 was designed around a mission gross weight of 16,825 pounds. Based upon this weight, the mechanical mixing was incorporated to compensate for various aerodynamic forces that are present as flight controls are moved and lift, torque, thrust, and drag are changed. Collective to yaw mixing was designed to compensate for "torque effect." This is a mechanical linkage and is not adjustable. When collective is increased or decreased the pitch in the tail rotor is changed accordingly a proportional amount. Below 100 knots collective to yaw mixing is not sufficient to compensate for the torque effect. Electrical mixing is incorprated to help compensate for torque effect by increasing/decreasing the tail rotor pitch i n additon to the collective to yaw mixing. When the airspeed is below 40 knots, 100 percent of the electrical mixing capacity is utilized. The mechanical mixing, collective to yaw, is operating at a constant ratio throughout the entire airspeed spectrum. As the airspeed increases above 40 knots, the tail rotor becomes move efficient and the cambered fin on the vertical tail pylon starts to become effective. With the increase of airspeed, the increased effectiveness of the tail rotor and cambered fin, tail rotor pitch must be decreased. Mechanical mixing wo~'t change without control movement. The digital com puter serves airspeed and as airspeed increases electrical mixing decreases until 100 knots. Beyond 100 knots, there is no electrical mixing being incorporated due to the capacity of· the tai l rotor and cambered fin to compensate for torque effect. 326 ,r~ TAl~ ~Dib~ S£~~0 -P~tes~u~~ SHvrorf ........ VIILIII! ~000 PSI ; 2 .., p~ y 0 p II rl p 3000 PSI PILDT A '59 ttl TJ