Definition of propulsion system for V/STOL research and technology aircraft

TABLE OF CONTENTS ICont'd) SECTION DESCRIPTION PAGE 8.0 APPENDICES A . INTERFACE DEFINITION FOR LIFT/CRUISE TURBOFAN ENGINE COMPONENTS LIST OF SYM80LS LIST Of TABLES DESCRIPTION PAGE

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Wind Tunnel Test Results PD370-254 Insta 1 led Performance PD370-254 Insta l led Performance, Customary Units PD370-2% instal led Performance PD370-25E Insb l led performance, Customary Units RTA ~ o t e r / ~ l c o h o \ Flow C o p c i t y Estimated weight Beakdown of Water Alcohol System for RTA T56 Planetary Parameters T56 planetary Parameters in PD370-25 RTA V/STOL Potential Gear Ratios Control led Variables LIST OF FIGURES DESCRlPTION PAG E

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Research and Technology Aircraft Propulsion System XT701-AD-700 Compressor Inlet Distortion Limit PD370-2% Insto 1 la tion Drawing PD370-25D Insta 1 lation Drawing PD37S25E Insfa l lation Drawing XT701-AD-700 Turboshaft Engine PD370-2% ~ i f t / ~ r u i s e Engine PD370-2% ~ i ! ! / ~ r u i s e Engine Normal Power ~istribution, A l l Engines Operating Normal Power Distribution, One Engine Inoperative Without ~ a t e r / ~ l c o h o I Normal Power Distribution, One Engine Inoperative 3% wo ter/A I coho1 RTA Duty Cycle Maneuver Loading Criteria Effects of ~ o t e r / ~ l c o h o l lniection on RTA Engines ~ a t e r / A i c o h ~ l System Schema t i c Compressor Inlet W i t h ~a ter/A lcohol Modifications LIST OF FIGURES (Cont'dl FIGURE DESCRIPTION PAGE PD370-25D Lift/Cruise Gearbox Geirercll Arrangement PD370-2% Lift/cruisc Gearbox General Arrangement T56 Reduction Gear Train Schematic T56 Safety Coupling Schematic T56-A 18 Safety Coupling RTA Propulsion System Control Sc:hcrnatic Pitch Change Control Schematic RTA Control A c t i v i t i e s RTA Propulsion System Development Plan Schedule Back to Back Test Rig vi.

The program objective was to provide engineering services in the definition of a turboshaft propulsion system for a V/STOL Research and Technology Aircraft.

Theso services included engineering support of a large scale variable pitch, lift/ cruise fan engine test in the 40 x 80 wind tunnel a t NASA Ames Research Center, technical support to b e i n g Aerospace C o m p n y and McDonnel l Aircraft Company, preparing a propulsion system performance computer card deck, preliminary design studies of system components and the preparation of a program plan for propulsion system development.

The propulsion system consists of two lift/cruise turbofan engines, one furboshaft engine and one t i f t fan connected with shafting through a combiner gearbox.

The turbofan engines are modified Detroit Diesel A l l ison XT701 -AD-700 turboshafts integrated with Hamilion Standard variable pitch funs. The turboshhft engine i s also an XT701-AD-700.

Very low core inlet distortion parameter levels were calculated from the limited

40 x 80 test du ta . The levels were we! l within the established XT701 -AD-700

limits and were not affected by nacelle angle-of-attack changes.

Aircraft contractors were provided propulsion system performance and physical characteristics for use in their research aircraft studies. Three engine-three fan system card decks were delivered to NASA Lewis Research Center, Nava I A i r Development Center, b e i n g Company and McDonnell Aircraft Company, System steady state performance can be calculated for either vertical or cmventional fjight.

A Iso the card deck w i l l calculate either installed or uninstalled performance. he user can determine system performance at the following optional modes: o Pitch and roll attitude control o A I1 engines operating o O n e engine inoperative, either a turbofan or the turboshaft o Water-alcohol injection o Contingency power level above in terrnedia te Detroif Diesel Allison and Hamilton Standard completed an agreement on the interface between the HS variable pitch fan rotor and the DDA turboshaft engine, p a r b o x and fan frame assembly.

The VTOL mode of opera tion w a s selocted as the design point using the more severe conditions of 3 engines and 3 funs operating or 2 engines and 3 fons operating with one l ift/cruise engine inoperative, P D376-254, the liFt/cruise o n ~ i n e designed for a fixed nacelle instal la tion, weighs 1066 kg (2351 ibs). The lightest t i l t nacelle inslal!ation engine, PD370-25E, weighs 1196 kg (2637 Ibs), The turboshaft center engine weighs 515 kg (1135 lks), The existing XT701-AD-700 engine requires modifications befora tleing fitted with tho variable pitch fan, Components requiring removal or mod1 ficotion include: inlet housing, compressor onti-icing, accessory gearbox, o i l system, torquemeter and engine eon trol5.

Optional water-alcohol iniection would add 4.5 kg (10 Ibs) to coch engine's weight and 10.3% total net thrust increase to a one lift/cruise engine inoperative system on a 305.3'K (90°F) day, Current engines have suffieiont turbine life to operate satisfactorily for 500 hours on rho RTA duty cycle and one hour a t a contingency power level, b v e l gear cross shaft drive locaiion study revealed thaf lightest gearbox arrangement for a fixed nacelle engine wclsn't the lightest For a t i i t nacelle engine. A cross shaft forword of the reduction gear i s used i n the Fixed nacelle engine and an aft crass shaft i s used in the t i l t nacelle.

T56 turboprop engine reducfion gearbox cornponen ts can be used in the lifti/cruise gearbox without a modification in the ratio.

A conceptual design O F the RTA propulsion system controls was undertaken with the XT701 control system being uti 1 ized. After meetings with the airframe monufucturers and analysis of the problem, i t was decided that the XT70T electronics had to be re- placed. A digital engine controller i s required to achieve the added nontrol tasks and f a i l operational reqvirernant. The XT701 hydromechanical c m trol i s retained with m i n o r modifications.

Propulsion system flight tested hardware con be delivered to the aircraft cnntractar 30 months aFter program go-ahead. The hardware would have cornplated 245 hours o f furbofan engine testing and 680 llours of fan testing.

1.0 Introduction Detroit Diesel Ailison (PIjA) Division of Genera! Motors Corporation provided engineering support to aircraft company studies of V/STOL aircraft during 1475, These studies established that the DDA XT701-AD-700 turboshaft engine was a logical choice for use in a research and techt7ology aircraft program and the need g*-r additiona 1 propulsion system definifion, The contract work defined the various components of the propulsion system shown in Figure 1, assessed the computibility of the existing XT701 engine in the V/STOL operational modes ond e x p n d e d propulsion system support to other aircraft contractors, The information obtained from this work can be used as the baseline in the design OF components for o reseorch and technology aircraft propulsion system as well as provide system performance to aircraft companies. DDA Model PD370-30 defines the system with separate jet turbofan engines, PD370-32 describes t h ~ system using confluent flow turbofon engines. ~ift/cruise turbofan engines are designated PD370-25 w i t h a suffix identifying aircraff installation, 2.0 Wind Tunnel Test Support An abbreviated large scale variable pltch lift/cruise fun inlet test was conducted t r ~ the Amos 40 x 80 wind tunnel in late July and early October 1976, The hardware used i n the test included a Boeing Company designed inlo t and a nacelle that incorporated a Hami I ton Standard 1.4 metre (4.6 feet) diamc ter variablo pitch Fan driven by a Lycoming T55-L-I I gas turbine core engine. The test was abbrcviatod due to hardware failures on the 26th of July and the 4th of October 1976. The fun cngine drive train was rebuilt aftor the first failure and only a limited number O F test conditions were run rr: the secand tcst before another failure..

DDA personnel reviewed the test plan and test instrumentation prior to the tcst and determined that sufficiant operating condi lions and tcst data would be avalfable to oval- uate the conditioas which would exist at the inlet of the care engine, This d a b was to

be used in estimating the effects of a variable pitch fan on cn XT701 core engine i n a

research ~ i r c r n f t by relating these da ta to prior XT701 engine test experience, Distortion l irnlts ore defined for the core engine inlet air pressure profiles in order to avoid three poton tia l detriments to acceptable ongine operation (1) compressor sta I I, (2) performance degradation and (3) excessive compressor blade vibration, Three prameters, circumferential distortion (Kg), radiai distortion (KR), and harmonie components of the inlet pressure distribution are used for evaluating distortion. The tesf date from the wind funnel was used to calculste the distortion prameters, The test w a s terminated with only seven of thirty-six test conditions completed. However, the limited &to was reviewed and typical points with the higher compressor face t o t ~ l pressure distortion (DISC) vaiues were selected for analysis, i t should be noted thclt the condition where inlet sepration occurred, c d = 600, Vo = 39 M/S (75 knots), was not selected for analysis because ~f the low compressor face distortion, Table 1 showsprametersoftheselectedpointsand theanalyticaIresultsofdistortion prameter calculatians. The negative KRE values colcula ted indicate a h u b oriented distortion. The XT701 engine i s sensitive to t i p oriented distortion and expected to be insensitive to the hub distortion. Accepbble limits of KGE and KRE are shown in Figure . The limits of the first four harmonic components are as foliofls: A l 5 1.6, ~~5 -20, and A 4 5 .20. Comprisonr show that the distortion levels obtained during t e abbreviated test were well within the limits established for *32f A \ 5 the XT701-AD-700 engine and were not affected by the angle-of-attack chongcs.

The distortion transfer coefficient i s described by the ratio of the circumferential distortion coefficient at the core engine (KQE) to the circumferential distortion coefficient a f the fan inlet hub ( K B F H ) . The Fan inlet hub distortion i s obtained from the readings of the h u h o s t

radial element when WEAI i s near 6%. When W E N 1 i s near 100/0, the readings of the two

radial elements closest to the hubare used fob calculating distortion. A few values of 0.

TABLE 1 . :,'JIND TUNNEL TEST RESULTS

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ENVELOPE KOE ACCEPTABLE .-.- " FIGURE 2. XT701-AD-700 COMPiJLSSOR INLEV DlSTORTlON LIMIT

KBE/KBFH are shown in Table 1 . These values show large variations in the transfer

coefficient. However, the very low values of both r'ne KQE and KQFH make this co- efficient highly susceptible fo experimental error and no conclusions regarding the distortion transfer characteristics of the fun can be drawn from this data.

3.0 Aircraft Contractor's Support Detroit Diesel A l !isall provided shaft driven propulsion system information to the b e i n g Company and McDonticll Aircraft Company for use in their research ond technology aircraft studies.

Itistallation druwings for the liFf/cruisc engine i n I fixed nacelle insfallation (PD 370-2541 and the two lift/cruise engines in a t i If nacelle instu llal.ion (PD370-253, These drawings show shaft forward and PD370-25E, shaft aft) have been provided.

envelope dimensions, accessory mounting puds and engirre weights. These drawings are shown in Figures 3 - 5 .

A large amount of inslulled propulsion sysbem performance has been provided Boeing und McDonnell during this work. Datu af the following conditions has been supplied: VTOL MODE 1 . 3 ~ - 3 ~ , SLS, 3 0 5 , 3 ° ~ ( 9 0 0 F ) D A Y , IntermedifitcPowerRafing 2. 3E-3F, SLS, 305.3'K (90°F) DAY, Intermediate Power, Attitude Control 3, 2E-3F, . , d , 305.3"K (90°F) DAY, Intern~ediate Power, One ~ift/Cruise Engine tnopero ti ve, DRY 4. 2E-3F, SLS, 305.3" K (90°F) DAY, In tel-mediate Power, O n e ~ i F t / ~ r u i s c Engine Inoperative, Dry, A ttifude Control 5.

2E-3F, 5 LS, 305.3Ok (90°F) DAY, In terrnediate Power, Onc ~ift/Cruise Engine Inopern tive, 3% W/A injection 6 , 2E-3F, SLS, 305.3OK (90°F) DAY, Intcrrncdirrtc Power, One l i F t / ~ r u i s e Engine Ir~operafivc, 3% w/A i n j e c t i m , attitude control 7. 3E-3F, SLS, 305.3OK (?O0F) DAY, Requested Thrust Levels below Int~rrnediafe Power 8. 2E-3F, SLS, 305.3"K (90°F) DAY, One t i t i / ~ r u i s e Engine Inoperative, Requested Thrust Levels CRUISE MODE 1 , 1 E-1 F, O.M, (0. FT) A LTlTUDi, O MACH PJUMBER, Standard Day I E - 1 F, O.M, (0. FT) ALTITUDE, . 2 MACH NUMBER, Standard Day 2.

1E-lF, 3048.M, (10,000 FT) ALTITUDE, .3 M A C H NUMBEL, Standard Day 3.

1E-IF, 6096.rv1, (20,000 FT) ALTITUDE, .5 MACH NUMBER, Standard Clay 4.

5 . 1 E- I F, d096.M, (20,000 FT) ALTITUDE, . 7 k%CH NUMBER, S tandurd Day 1 E - I F , 11000.M, (36,08?FT)ALTITUDE, .7MACHNUMBER, StandardDoy 6.

7. 1 E-IF, 11000,:4, (36,089 FT) ALTITUDE, - 9 MACH NUMBER, Sfandard Day The above Tables 2.- 5 show performance parameters of typical operational modes.

data was ca l culated using DDA inhouse performance t e c l ~ n i ~ u e s before the propulsion system card deck wa; complete.

Control margin i s defined as the thrust available for aircroff attitude control w l ~ i l c the engines are operating a t intermediate power. With horsepower transfer and Fan pitch angle adjustment, thrust From one turbohn can be increased as much as the control margin while tlie thrust from the other turbofan i s decreased the same amount.

The attitude control thrust values given in Tclbles 2-5 ore For either the high (+) or tow (-) turbofan.

FGURE 3. PD 370 - 2 5 A INSTALLATION FIGURE 4. PD 370-25D INSTALLATION L7M) / S / J (67.2) - , 7 0 7 - b \ d ENGINE CG .

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- h r - - i . ~--- - - - i 3 4.0 DIgital Computer Simulation Program The Detroit Diesel Allison Division of Gunoral Motors Corporation has prepared a program in the form of a deck of cards for use in calculating the steady state perform- ance of the DDA Modcl PD370-30 and PD370-32 propulsion systems, The program considers tho p r o p u l s i ~ ~ systarn as an arrangement of power producers and thrusting units coupled together by shafting through a ccntrolly mounted mixcr gear box as shown by the sketch of Figure 1 .

The Model Pi3 370-30 three engine-three Fan system has two turbohn engines which are cansidereu to be separate jet engincs (i. e : . , primury and secondary strean~s exit from separate nozzles), Model PD 370-32 system turbofan engines are considered to be confluent flow jet enginos (i.e., primary and secondary streams are combined before exiting from jet nozzle), M i x i n g i s accomplished by equal static pressures of the streams.

The remote fan and the two fans of the turbofun engines are identical Hamilton Standard variable pitch fans of 157.48 CM (62 inch) tip diameter.

The turboshaft engine and the two g a s generators of the turbofan engines are Detroit Diesel Allison XT701 turbo- shaft engines, With this program, the user may, for a given set of flight conditions, calculate steady state performance by entering the proper valt~es of the program input data.

Steady s b t e performance can be calculated far either vertical or conventional flight.

Vertica I flight performncc. i s ca lcula ted with the remote fan producing vertica l thrust and the turbofin engines configured to produce vertical thrust. Conventional flight performance i s ea lcula tcd with only the turbofans producing norms l forward thrust.

Vertical Flight performance ca fcvlations are perFormed under the following constraints: o eqvali ty of fan rotational speeds o equality of turbine inlet temperature, turbofan and turboshaft units o power &r lance in shafting o equality of l i f t thrust, turbofan and remote Fan cnits inon-altitude control ~ I Y ) Fan rotational speed i s Fixed a t a preset valuc? For. a l l vertical flight calculotions.

Equality clf l i f t thrust From the. remote fun and the two turbofan uniks i s achieved by adjustment of the i n d i v i d w l fan pitch angles, olang with power transfer within the shaFting for delivet'y of required power to each l i f t unit. Gcuring iosscs are con- sidered during the ca lcula tinn of power transfer, The thrust of a turbofun i s calculated us the total net thrust, the sum of the primary and socondory thrusts. Howcver, the residual thrust of Ihc turboshaft engine i s not considered duriirg thrust bo lancing, The syrtcm perfo~-rnancc can be colculnted duritlg vitrtir;ol modes of lcvcl flight and pitch and ro!l ottitude control. For uach of these modes, pc.rforrnonce can be calculated with a l l engines operative ui wikh anc cnginr! inoperative. Either ct turbofan or the turboshaft engine cun be inopsra l i v e . Performance con Lo co lculc~kuij either "dry1' or with u preset amount of water/oIcohol injection augmentation, Attitude control performance i s calculated mi the basis of a thrust increment, which is itiput by the user, Far r o l l control, this increment i s applied to one of the turbofan engines by adjustment of its Fon pitch angle, The fun pitch unglcs or the opposite t u r b f u n enginc and thc remote fan ore then adjustcd until a solution i s achievc?d which simultaneously balances power tra~isferred through the shaFting and causes the thrus! of the remote fan to Ix on average o f thc two turbohn thrusts. Thus, no pitching momenl exists.

For pitch control, the input thrust increment i s applied to the remote fan. The fan pitch angles of the turbofun engines arc then adjusted until a solution i s achieved which simultaneously balances power transferred through the :,hafting and mainbins equality of the turbofan engine thrusts, Thus, no roll moment occurs during pitch control , When a turbofan engine i s inoperative, power i s transferred tt11.0~lgh the shafting to sustain the Fan of the inoperative engine, Again, fan pitch angles of both turbobn engines and the remate fan are adiusted until thrust equality of the thi ee units or the ijropei- ro!l or pitch control thrust i s achieved.

Vditti the turboshaft engine illoperative, sufficient power i s transferred through the shafting to drive the remoie fan. A l l b n pitch un51es are adjusted fo achieve equal fhrust or control thrust.

Normally, vertical flight performance i s calculated a t "intermediate" power level.

However, the user can request a thrust level for any vertical flighf operational mode.

This thrust must be lower than that ~ b t a i n c d a t the intermediate power. The program w i l l calculate the fan pitch angles. A "contingency" power l e v e l greater than intermediate i s also available. Thc performance calculofsd a t this power level i s to be used for in forma tion purposes only. The intermediate power level i s the qualified rating o f the XT701-AD-700 engine. Any higher rating such as contingency would require development and testing beyond the scope of the baseline Research and Technology Aircraft development program.

The system performance during conventional flight i s calculoted with the t.urboshaft engine unpowered and disengaged, and the remote fan disengaged, The turbofan engines are configured so as to produce normel forward thrust and are considered equal i n a l l respects. The center mixing gear i s still coupled to the system. Thus, the calcu- lation becomes one oFa conventional turbofan with the addition of a sniall amount of power flowing to the center gear from the cross-shaft, T h i s power i s composed of the center gearing losses and any requested customer power ex traction from the con ter gearbox.

Predefined power levels of "infermedia tell and "rnaxiqum continuous, I' and a !so lower undefined power levels are available for the study of climb and cruise performance in the ccslventional flight mode.

Certain temperature and rotational speed limiters are built into the program. A t any flight condition and operational mode, system performance i s limited by whichever of these limiters i s i n effect. T h i s i s in accordance with standard procedure tor engine performance calculations, Additionally, for this propulsion system, the fan pitch angle i s limited to within a range bounded by o minimum (negative pitch) and a maximum (positive pitch' >.his effectively limits the torque moment available for r o l l and pitch 11s analytical performance model.

maneuvers in The card deck as delivered to users calculates the performance of an uninstalled propul- sion system. The user can input hi: qwn installa tion fuctors such as inlet losses, nozzle coefficients, and flow bleeds and power exrractions. The detai I s OF the method of entering these and other inputs i n t o the program are described i n a users manual report, ~ h i c h accompanies each card deck.

5.0 Design Information and Analyses of Shaft Driven ~i f t / ~ r u i s e Fun Propulsion Systom 5. 1 Genera I Description The shaft driven lift/cruise fan propulsioil system for the V/STOL Research and

Technology Aircraft shown in Figure 1 consisfs of two lift/cruise (UC) turbofun

engines, one turboshaft engine and one l i f t fan connected toqciher with shafting through a combiner gearbox, A disensging clutch permits the disconnection of the l i f t fan during cruise operations, Design work under this contract was limited to the definition of the lift/cruise turbofan engines and the fur boslia ft engine. The l ift/cruise engine, Detroit Diesel Allison Model PD 370-25, i s a modified DDA XT701-AD-700 turboshaft engine integrated with a tlarnilton Standard variable p i t c h F a n in a conventional front fan arrangement. A significant portion of the engine's pQwer con also be transf~rred through a radial drive gear set and cross shaft t o the other lift/cruise fan or the l i f t b n during V/TOL operations, attitude control operotions or emergency engine out operations, The center engine shown i n Figure 6 i s o standard XT701 except for o i l system and control modifications. This arrangement weighs 515 kg ( 1 135 Ibs).

Two basic concepkr of vectoring the thrust from the I i Fi/cruise engine were considered. 'The "fixed nacelle" engines (PD 370-254) have the thrust vectored by use of a deflector exhaust nozzle that i s nacelle mounted. The " t i l l nacelle" engines (PD 370-25E) J.~ave the thrust vectored by ti1 t i n g the nace l le including the engine. This neccssibtes a single point mount about which the engine rotaver. The rotation centerline is coinci- dent with the radial drive cen terline, Figure 7 shows the fixed nacelle l ift/cruise engine PD370-2%. general arrangement and Figure 8 shows PD 370-25E the lift/cruise engine used i n a t i l t nacelle instal la- tion. In both engines the power turbine drives the fun through a planetary reduction gear set to provide the proper sp-ed match. The L/c engine yearbox also incorporates an overrunn tng ct utch to allow the Fan of on inoperable engine fa be driven by the cross drive without the necessity of driving the inoperable power turbine. The ertgine and oirctafi accessories are driven by a core mounted accessory box a f t of the fan support assembly.

A n interbce agreement has been reached between Detroit Diesel Allison and liarnilton Standard on the interface o f the HS variable pitch fan rator and the DDA turboshaft engine, gearbox assembly and fan frame assembly which a l l together form the lift/ cruise turbofan engine, Appendix A i s the interface definition between DDA and HS w i t h the latest revisions included. This interfuce vrill be updofed a n d refined us the definition and design o f the V/STOL Research and Technology aircraft propulsion sysfern continues.

FIGURE 7 . PD 3 7 0 - 2 5 A LIFT/CRUISE ENGINE 5.2 Design Requirements The design requirements for the propulsion system are as fa1 laws: 5.2.1 Design Point VTOL mode of operation (Design for the more severe condition of 3 engines a n d 3 bns operating or 2 engines and 3 fans operating with one li Ft-cruise engine inoperative).

Sea level static, 305,6OK (90°F) day conditions, Intermediate Power (maximum thrust condition). Nominal maximum thrust horsepower distribution for all engines operating, one l i f t cruise engine inoperative without water alcohol, and one L/C engine inoperative with 3O/a water alcohol, are shown i n Figures 9 , 10 , and 1 1 respectively, 5.2.2 Operating Envelope The propulsion system shol l be designed for opera tion in a cruise mode up ta 1 1887.2 The temperature metres (39,000 feet) and a V T O t mode up to 609.6 metres (2000 feet).

range for operation sbl\ be 219.4' to 3 4 7 . 2 ' K (-65' to f165°F).

5.2.3 Dynamic Thrust Response The propulsion system shall have a thrust response characteristic (time constant) of 0.2 seconds or less. Time constant as used here i s defined as the time required to change thrust by 63,2% of the total thrust change after o step change in the power (or control) level.

5.2.4 Stall Margin The fan stall margin requirements for the l i f t fan and the lift/cruise fans shall not be \ess thun: 20% for takeoff and landing 20% for cruise 10% for Maximum Control = Totu 1 pressure a t surge 'SURGE

'SURGE X wA OP pt - 1

= Total pressure at operating point STALL MARGIN = P P

w~

O P ~t SURGE O P WA op pt = A i r Flow at operating point = A i r Flow a t surge W~suRGE 63 1 5kW (8469 shp)

T

(610shp) (610 shp) 6485kW+ (8697 shp)

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LRGJ

6541 kW / (8771 shp)

CONDITIONS: INTERMEDIATE POWER, 305,3'K (90°F) DAY, INSTALLED FIGURE 9 . NORMAL POWER DISTRIBUTION, ALL ENGINES OPERATING 3610 kw (4841 shp) 2557 kW

-- .-a - - -- --- -

6168 kW (8271 shp) (3429 shp)

t

' RG- J

5571kW (7472shp)

6221 k b (8342 rhp)

CONDITIONS: INTERMEDIATE POWER, 305.3' K (90°F) DAY, lt45TA LCED FIGURE 10. NORMAL POWER DISTRIBUTION, ONE ENGINE INOPERATIVE WITHOUT wATER/A LCOHOL 4186 kW (5614 shp) 3816 kW (5117 shp) 4289 kW

--+... - .--- - . - J

(575 I shp) CONDITIONS: INTERMEDIATE POWER, 305.3OK (90°F) DAY, INSTALLED NORMAL POWER DISTRIBUTION, ONE ENGINE INOPERATIVE, FIGURE 1 1 , 3% WATER/A LCOHOL 5.2.5 Specifications The pr~pulsion system shall be designed in accordonce with M i l i t a r y Specifications MIL- E-50070 and MI t-P-26366, Detroit Diesel Allison Prime Item Development Specification 8448 covers the rcqulre- ments of the XT701-AD-700 engine, 5 , Z . b Design Life The 1 i f t fan assembly and the l ift/cruise turbofun engine and a l l components thereof shcrll be designed for 500 hours of operation with 500 research and tschnology aircraft duty cycles. The duty cycle i s shown in Figure 12.

A l l propt~!sicn system components shall have a design life of 50 hours or more a t a VTOL moximum thrust condition, Maximum thrust i s defined CIS the thrust available from 3 engines anc! 3 fans operating ot the intermediate power rating or from the intermediate power operation of 2 engines and 3 fans with one lift/cruise engine inoperative, A l l propulsion system components shall have a design l i f e of 5 minutes or rnore during VTOL maximum attitude control operations.

5.2.7 ~ e c h c l n i c o l Limits R P M

-

Engine Gas Generator Turbine Speed 15450 Engine Power Turbine Speed 5.2.8 Strucfural Limits The propulsion system shalt be designed to operate with the acceleration loading envelopes shown in Figure 13 .

The engine, gearing, clutch, shafting, front frame, bearings and a l l components of both l i f t and lift/cruise fans shall be capable of withstanding a l l loadings imposed by the loss of a blade she1 I, leading edge sheath und fi 1 1 .

5.3 Design Goals 5.3.1 Design Approach Designs for engine modificatims and new components shall conform to a philosophy where weight i s minimized (within state-of-the-art. technology) while minimizing cost and 3 1 FJGURE 12. RTA DUTY CYCLE o MANEUVER LOADING CRITERIA UP

4 . -

) ,-3 --

e = ? 6 rad/sec2 .

4 I ) = 0

f---. ' - 7 $ = 0

1 2 - Y- \ t 4 i

I

'T

n f t e - I , . Fore

4 i 2 4

- -

A f t - - -

2 - 9=+2 rad,/sec

' -11J.-ore

S.L. = 4.0

4 - J

i

S * L . = 2.0 )

i 4 1 , i

B = O T- 6 - . ? : .

i = o ~i i

I & 4 , , 8 - I \ -1

8 =%14 rad/secz d = O i 6 -

! - , -

I.

S.L. d . 5 '

+ =+ 6 rad/sec2\

9 8 - i

u

u 1 0 .

lo i

Down 7 D O T J ~ Landing (0 to Flax. Thrust) F l i g h t (0 to &lax. Thrust1 Accelerations in g ' s AcceLerations in g's FIGURE i3. MAFdEUVER LOADIHG CRITERIA maintaining adequate safety margins. The engine modifications shall provide engine and drive capability which i s corisistent with operational requirements of the fans, lnterchangea bi lity O F lift fan and l ift/cruise Fan components shall be accomplished i n the design, In designing the propulsion system hardware, maintuinabili ty shall be considered.

Provisions shall be included for easy access to hardware requiring service, inspection and/or repair. "High risk" parts (those vulnerable for FOD, or fhosc where life esrpechncy may be reduced due to severe operating conditions) should have field joints to preclude the need For major teardown during a r e p i r cycle. Also, access ports should be located in sba tegic areas of the assembly where borescope or other routine nondestructive tests muy be performed. Self-conla ined I ube/hyd~+auI ic/ electrical systems shall be used.

5.3.3 Fan Distortion The fan shall have no greater than a 5% thrust loss due to the Following distortions, considered sepclra tely: a . Pressure Dislortion

Inlet total pressure distortions, * (152.4 MM (six inches) forward o f

( 1 ) fan face) of 15 percent

Exit stotic pressure distortions * (152.4 MM (six inches) a f t o f fan stator

(2) exit) of 15 percent b, Temperature Distortion Temperature over 50 percent of inlet. Face a t least 283OK (50°F) above ( 1 ) arnbien t in lef temperal-r~re Changes of average temperature of 283'K,h ec (50a~/sf ) for 1/2 (2) second "Distortion = P

MAX - 'MIN

5.4 XT701-AD-700 Engine Modifications When used as the center engine in a three engine aircraft, the existing turboshaft However, when the existing sngines are engine may be used without modification.

or 8 some additional modiFica- fitted with variable pitch fans as shown in Figures 7 tians ore required.

5.4.1 Inlet Housing and Compressor The inlet housing on the >(T701 engine i s removed beFore incorporating the variable pitch fan and k i n frame assembly. The removed inlet guide vanes, front compressor b a r i n g and related hardware from the inlet housing are incorporated into the Fan frame for use on the lift/cruise fun engine.

Tile PD 370-25 compressor discharge reaches a maximum temperature o f 71 3" K ( 8 2 3 " ~ ) a t sea level in terrnedia te on a 305" K (90°F) day. Compressor discharge tota I pressure i s 1406kPa (204 psia) for this poink. The compressor discharge total pressure of the PD 370-25 reaches a n~aximum of 1482 kpa (2 15 psia) a t . 4 Mn at sea level intermediate

with a compressor discharge temperature of b81° K (767OF) . This compares to design

mnximurns of 1431 kPa (207.5 psia) compressor discharge tatal pressure ond 757°K (903OF) compressor discharge temperature for the XT701 engine. Prel iniinary analysis indicates no compressor modification is required for the lift/cruise engines.

It i s felt Compressor anti-icing w i l l not be provided on the PD 370-25 UC engines, that the nature of the flying envisioned for the RTA does not warrant its inclusion, Modifications to the compressor inlet to accept the wate,~/olcohoI injection option are discussed i n Sections 5.5.

Preliminary studies indicate there i s no requirement for modification to the present XT701 di ffuser/combustor cornponen t 5 for use in the li ft/cruise engines.

5.4.3 Turbines There are n o present requirements for modifying the XT701 turbines for use in the PD 370-25 unless crddi tional cooling i s desired for the H . P. turbine for short term contingency opera- tion, Modified turbine cooling i s discussed in Section 5.6.

5.4.4 Shafting and Rotor Dynamics - -.

To provide for growth, the XT701 mainshafting was sized for 15,000 LCF cycles from 0 to 7457 N m (5500 ft-lb). This would provide 9,000 KW (1 2,000 tip) capability without modi fico tion.

The rotor dynamics are dependent upon rotor mass, rctor construction, ma inshaft spun a n d stiffeners and bcaring suppo~.t rates. As these parameters remain unchanged, the rotor dynamics for the L/C engines should be acceptable.

5.4.5 Mainshaft Bearings and T h r ~ ~ s t Balance The mainshaft bearings remain unchanged. Preliminu~*y vent. system and thrust balance analysis indicates that thrust lwds remain a t acceptable levels.

5.4.6 Accessory Gearbox

The new fan frame used on the 1/C engines does not provide clearance far the starter

when mounted on the present accessory gearbox. This, and the probable requirement for providing power frwn the high pressure rotor For driving aircraft hydraulic and eiectricol power systems, necessitate the design of a new accessary gearbox for the PD 370-25, A tenta tive accessory arrangement can be seen from the instal lo tion

drawings, Figures 3 and 5 . The arrangement shown can be modified t o suit

the individual airframer's requirements when they are better defined.

5 . 4 . 7 Oil System The present XT701 Oil System i s driven by the high pressure rotor. In the PD 370 i t i s possible to have a power section failure but to huve the lift/cruise gearbox driven via the cross shafting from the remaining operable engines, and in this case, the present system would n o t provide gearbox lubrication. Therefore the XT701 o i l system must be modified to provide a lube pump driven by the L/c gearbox.

The present o i l system for the power section must also be retu ined to assure proper lubrication when the interconnected low pressure rotors are dropped o f f in speed.

The fwo pump system can hove some components in common, however, s u ~ i ~ as f h e supply tank, the Filters, and the o i l coolers.

The labyrinth s e a lr for the L/C gearbox-compressor in l e t assembly must be replaced by carbon face seals because in the case where the L,/c gearbox i s operable and its power s e c t i m i s not operating, there may not be sufficient a i r pressure available to pressurize the gearbox labyrinth seals.

Also in the case of the ti1 t nacelle engine sump, modification must be accomplished to allow scavenging and to prevent leakage when the engine i s operated or shut down i n the vertical position.

The fuel consumption of the lift/cruise engines i s insufficient to provide use of the fuel as a s o l i b r y cooling medium with the additional cooling load of the l i ft/cruise gearbox.

Therefore supplemental cooling must be supplied by either an airframe mom ted airfoil cooler or a i r f o i l surface coolers mounted in the fan flowpctth in the nacelle.

5.4.8 Torqueme ter The present PD 370-25 engines have the XT701 torquemeter deleted. It i s thought that the lives of the engines can be enhanced and a lighter installation obtained i f the engines are matched on turbine temperature i n lieu of torque. The a d d i t i m o l gear and cross shaft loads incurred as a result should not noticeably affect gear/bea~~ing/sl~aft system life.

5.4.9 Con trsls XT701-AD-700 engine contFols require modification and these are described in Section 5.10.

5.5 Wa ter-Alcohol Injection 5.5,1 In froduction The RTA 3 engine-3 fan propulsion system does not include a water-alcohol iniection system. However, the a d d i t i m of the system to the turbofan and turboshaft engines can be accomplished. The system would be added for the purpose of providing additional power and thrust to the system when one engine becomes inoperable in V/TOL flight modes, The augmen to tion would add to the output of two engines and three fans.

5.5.2 Effects on propulsion System Performance The effect of wa ter-a I cohol injection on the horsepower and thrust augmenta t i m of the RTA lift/cruise engines and center turboshclft engine is obtained using an empirical method, This method i s based on c i v i l T56 (501-013) turboprop engine tests with wcrter- alcohol injection. Briefly, the method i s based an the experiment01 observation that the 501-013 augmented shaft horsepower, at a given compressor inlet temperature, can be calculate!' by running the engine nonaugrnented but a t a reduced compressor inlet tempera- ture. This tempera ture i s defined as the equivalent in let tempem ture.

The wa ter-a1 cohol mixture imp1 ied in the equivu fen1 tempera fure correlation consists of 2/3 water, 1/3 methanol by volume and i s injected immediately upstream of the com- pressor i n the case of the lift/cruise engines, and at. the compressor inlet for the center (turboshaft) engine. I t was assumed that the 501-Dl3 equivalent temperuture vulues could be directly applied to the RTA engines based on a general similarity of compressor charac- teristics. To simula te water-a lcohol iniection on the customer card deck cycle, the inlet temperature into the compressor i s automaticol l y reduced to the equivalent temperature value from the actual Compressor inlet temperature according t o the expetimenta l correlation. In the ca;e of the lift/cruise engines, these values correspond to the Fan discharge temperature and pressure. The resulting cycle calculation produces an increase in the horsepower delivered to the fans with a corresponding increase in engine thrust.

For the Typical results of the simulated water-alcohol operation are shown in Figure 14 l ift/cruise engines and furbosha Ft engine for a range of engine .in let ternperotures; a ISO shown for comparison i s khe 501-Dl3 engine test data.

The RTA propulsion system computer card deck can be used to calculate OEI performance with water-alcohol augmentation. O n a standard day, the total net thrust of a dry, un- installed system i s calculated fo be 131325N (29523 Ibs) and when water-a lcohol i s injected into the remaining turbofan and turboshoft engines tht: total net thrust increases to 138197N (31068 I b5) on increase of 5.2%. On a 305.3'K (9O0F) day, the total net thrust of a dry, uninstal led system i s 120556bl (27102 Ibs). W i t h wu ter-alcohol iniecfion, the total net thrust i s 132953N (29889 Ibs), an increase of 10.3%.

5.5.3 Addition of Water-Alcohol System to the Propulsion System The addition of a water-alcohol system to the I ift/cruise turbofan and the XT701-AD-700 ~urboshaft engine requires minor modifications to the fan frame a t the compressor illlet and inlet housing, plus the addition of the required plumbing, valves, pumps, tank, etc. as described below.

Water-A Icohol Mixture A 33% n.rethanot, 67% distilled water mixture by volume i s used.

Genera I Arrannemen t of the 5 ystem shows the water-alcohol system general arrangement For a 3 engine V/STOL Figure 15 configuration. Figure 16 shows the iniection components installed in the turboshaft shows the nozzles instal led in the fan kame a f the inlet of the engine. Figure 17 turbofan compressor.

Tables 6 acd 7 show fhe water-a lcohol Flow capacity and weight breakdown, respec- tively.

I 3Eg3F SYSTEM

1s - ENGINE INLET TOTAL TEMPERATURE FIGURE 14. EFFECTS OF WATER/ALCOHOL INJECTION ON RTA ENGINES ..- -.-.-

---

FIGURE 16. COMPRESSOR INLET WITH W~LTER/ALCOHOL MODIFICATIONS

FIGURE 17. TURBOFAN IN LET w ITH WATER/ALCOHOL

MODIFICATIONS TABLE 6.

RTA ~ a t e r / ~ l c o h o l Flow C o p c i t y RTA WATER ALCOHOL MIXTURE RATIO Methanol Water 67% 67% E N G I N E REQUIREMENT N o . Engines 4 3 Required Engine Flaw, litre,/sec (GPM)

- 7 17 0 . 6 )

- 5 (8) Toto l Required Flow, litre/sec IGPM) 2 1321 7 (31.8) M O T O R DRIVEN PUMP No. Used 2 2

Single Pump Capocity 7 1034 kPa, Iitre,'sec (150 PSI, GPM)

2 (32) 2 (32) FLOW REGULATOR AND SHUTOFF (DDA SUPPLIED) Required Flow, l itre/sec (GPM) .5 (8) . 7 (10.6*) * NOYE: T56 Flow Regulator W i l l be Modified Internally To Provide Required Regulated Flow Output TABLE 7.

Estimated Weigh1 Breakdown of Water A lcohol System for kTA 2 W/A pumps 2 Pressure 5wi tchas .9 12) 2 Check Va l ves e 9 (2) 3 Regulators and Shutoff 9.5 (21) 30 Nozzles 4 (9) Tank, Wiring, Piping and Mounts 44 (96) An interface definition w i l l bo established between airframei- NOTE: and DDA on supply of W/A system components, Therefore, the above components, except for regulators and nozzles, hove not been included in engine weights, Components The engine furnished system consists of an aircraft mounted flow regulator and ten in lector nozzles. The regulator control valve solenoid i s energized to open the regulator valve, The regulator controls the flow a t ,669 litre/second (10.6 gal I o n s per minute) per engine, The ten nozzles are mounted on the i n l e t housing, in the front of the HP cornDressor inlet.

An i n l e t Each nozzle has a plug in the center which produces a widely diffused spray screen incorpora tcd in ouch nozzle protects the nozzle from contamination, The fuselage-mounted tank has a useable capacity o f 121 to 148 litres (32 to 39 gallons).

This results in u 'low duration of over one minute, sufFicicnt for one hot day b k e o f f or an emergency landing, The centrifugal pumps are motor-driven and each has a c a p c i t y o f 2.0 litre/second (32 One pump i s capable of s u ~ p i y i n g the required volume o f flow. Two gallons per minute).

pumps mounted in para1 l e l are desirable for r e l i a b i l i t y . The pump outlets connect i n t o a distribution manifold which delivers water-alcohol to each engine.

5 . 6 XT701 -AD-700 Engine Short-Term Con tir~gency Operations 5.6.1 Introduction The XT7Ol-AD-700 engine i s qualified a t an intermediate power rating and this i s the maximum rating used for the engines in the RTA propulsion system. A n y highar rating such as contingency would require development and testing beyond the scope of the h s e l ine RTA development program. However, since aircraft total weight and thrust to weight r a t i o requirements a t ViOL modes of operation might exceed the thrust available in a one lift/cruise engine inoperative condition, o feasibility of operating the XT701-AD-700 engine a t short term contingency levels above intermediate powers has been studied.

5.6.2 Component Life A prelimir~ary study ha. .>. t~ that the XT701-AD-700 engine has sufficient turbine life to operate so tisfactor! l y FOI ~d hours to the duty cycle shown in Figure 12 , w i t h the inter- mediate power rating of 1561" K (2350°F) burner out temperature, plus one hour of c m t i n - gency opera tion a t a butner out temperature of 161 7" K (2450°F), T h i s study considered two types OF engine environment, o lift/cruise engine with a 1.218 pressure ratio fun Forward o f the compressor, and a center engine without the ram effect of the fan.

The H P 1 vane showed the highest surface temperatures, w i t l i maximum loca I temperatures of 1397O K (2054°F) and 1439') K (2 130°F), a t burner out temperatures of 1561 O K (2350°F) and 1617O K (2450°F) respecfively . W i t h current cooling levels, some vane deterioration would be expected a t the higher temperahre.

The analysis indicated that fhe HP2 blade has the lowest sfress rupture life, but would meet fhe above requirements w i t h current cooling levels.

5.6.3 Modifications Required ----- ..

The preliminary study has shown that the existing XT701-AD-700 engines have sutfieicrt t Mowcver, slight increases in 1 s t \:une and turbine l i f e to meet the goals defined obove.

2nd blade cooling flows, which could be accarnplishcd by increasing metering hole sizes, would provide rsdt~ced nlctal tempera lures for additional operctting margin, 5.6.4 Contingency Performance Augmerrfo tion - - - --.

The RTA propulsion system computer card deck can be used to calculate one level of contingency. This level i s a 57" K (100°F) increase in burner outlet tcrnpe~ature (BOT), Using Boeing ins@ 1 lation hctors, one ! ift/cruise wlginc inoperu t.ive performance on u 305.3"K (90°F) day ha5 been calculated with the new card deck, The total net t t ~ r u . t available at intermediate power i s 1 16Y08N 126292 Ibs) while a t the contingertcy l o v t I the total net thrust is 120tt?2N (27018 Ibs). The increased DOT provides a 2.8% incrt:osr~ i n thrust.

To ew lua te the effects of water-alcohol injection and contingency level ~urbine ternpera- lures on t o b l system fhrust, ca!culations were made usillg Boeing installation faclor~ on u 305.3OK (90°F) day. With one lift/cruise engine inoperative, the total net thrust wus 1337434 (30067 Ibs), an increase c r 14.4?/0 over the intermediate power setting tokil thrust.

5.7 Bevel GearCrmsShaft location Study I t became obv:nus early in thc V/STOL studies tho t the general arrangement tlw t a f f o r ~ k d tlie lightest wr ~ g h t for the fixed nacelle engine (Figure 7 ) with the cros5 shaft dt ivc: located i n front of the reduction gear was probably not the optimum errcli~gt:mr?nt far he ti;: ,SAcel le engine instul lations.

I t i s necessary i n the t i l t nacelle engine for the eny ine to rotate ubout the center line of the cross shaft. placing the cross shaft Forward an the engine leads to three distinct dis- advantages: o The distance from the pitch axis ko thc cxhausk n n r z l e i s greater, which would result in longer landing gear for an eqmI nozr!e ground c l c ~ i ~ ~ n c e . While not possible For h e engine manufacturer to evaluate the impact on systenl weight i f i s believcd to be consideroble.

The distance from the engine pitch oxis fo the crrgine C S i s greater, resulting in a o larger CG shift when the engine i s rotated and also resulting in larger engine mount.

loads.

o The engine must be lengthened over the Fixed nacel le engine with the same gear arrangen~ent to provide a load path for the engine mount loads to be distributed into the fan housing, This increases engine weight.

With these Factors in mind, two lift/cruise gearboxes were designed; Figure 18 shows the optimum cross shaft forward arrangement modified to a t i l t n ~ c a l le mountii~g. Figure 19 shows the optimized cross s h a f t a f t design. The designs were coupled with an engine to farm the PD370-25D shaft-forward and PD370-25E shoFt-aft turbofun engines. Installation envelopes are shown in Figures 4 and 5 , Pertinent data i s summarized below: Shaft Fwd Shaft A f t Engine M d d f D370-25D P D370- 2 5E Cross Shaft to exhoust flange length mm (in) 1971(77.6) 1819(71.6) Overall length rnm (in) 3183(125.3) 3155(124.2) Cross Shaft to C/G mrn (in) 406(16.0) 315(12.4) Weight Kg (pounds) 1 262(2783) 1 196(2637) Mount Moment N- m (Ft. Lb.) 5030(3 71 0) 36?5(2725) 1 t can be concluded therefore, that for a t i l t nacelle installotion, the aft shaft engine i s lighter, shorter, and should also result i n substantial savings i n airframe weight.

5.8 Preliminary Design Studies of lift/cruise Engine Unique Components Since the purpose o f the Research Technology Aircraft i s to prove a flight concept and not hardware technology, the criterion for the design of the components unique to the lift/cruise engines i s to design within the present state-of-the-art using present design concepts, ma teria Is, ma teria f a 1 Iowa bles, and ma king the maximum use o f existing hardware where possible.

5.8.1 Reduction Gear P r o b b l y the best example of the use of existing hardwore i s the reduction gear. f h e T56 turboprop engine reduction gearbox shown i n Fis!~re 20 has a planetary output stage with a ratio of 4.33 to 1 . The input i s to the sun gear a t a speed of 4424 R P M and the output i s from the planet carrier a t 1021 APM. This gearbox has accumulated over 65 million flight hours and i s capable of continuous operation up to 4101 kw (5500 horsepower). The opera ting parameters for the T56 planetary i e t are shown in Table 8 .

FIGURE 18. PD370-25D LIFT/CRUISE GEARBOX GENERAL ARRANGEMENT ,- -- I . .

,-.

, . , . a { , r-* - -

4 .- . .

. -

8 . , cia I FIGURE 19. PD 370-25E LIFT/CRUIS€ GEARBOX GENERAL ARRANGEMENT

FIGURE 20. T56-A- 7 REDUCTllDN GEAR TRAIN SCHEMATlC

TABLE 8 . T56 PLANETARY PARAMETERS SUN PLANETS ( 5 ) RING

- -

NUMBER OF TEETH 3 5 100 DLPNETRAL P XTCH 6 PRESSURE ANGLE 2 2 5 O F I T C B DtAMETER- 148.2 427.6 15.8333) (in) (16.8333) FACE WIDTH-rnm 56.9 53.2 ( in) (2.24) ( 2 . 0 9 5 ) RP M 2916 0 The design speed of the 1575 mm (62 inch) diameter Hamilton Sta~ldard Fun i s 3543 rpm and the design speed of the XT701 power turbine i s 12,000 RPM, therefore the required reduction r a t i o i s 12,000/3543 or 3.38/1.

When the existing T56 planetary gear set i s used as a "star" or fixed planet set the resul tant ratio i s 100/30 or 3.33 to 1. This provider a match a t o turbine speed (NT) of 1 1,810 R P M and a fan speed o f 3543 R P M according l o prcsen t component and cycle analysis and the reduction in performance parameters was negligible, Therefore, when the gear set i s used in this manner, operating parameters shown in Table 9 arc applicable, TABLE 9 T% PLANETARY PARAMETERS WHEN USED IN PD370-25 P L A N E T S (5) C( 1 N G SUN --*-

-

30 3 5 1 00 NUMBER OF T E E T H DiAMETRAL PITCH 6 6 6 PRESSURE ANGLE 2 5 ' 2 5" 2 5"

PITCH DlAMETER - mm 131.2 148.6 427+6

FACE WIDTH - mm 66.7 56.9 53.2

BENDING STRESS ' - MPa

(5 B) (f'S 1)

CRUSHING STRESS * - MPa

( S C ) (PSI) MATERIALS AM5 6265 AM5 6265 AMS 6265

"

ALL ENGINES OPERATING - 6540 k w (8771 HP)

-12 I t i s also easy to attain other reduction ratios by changing the number of teeth in the sun and ring gears while keeping the planet bearings unchanged; in this manner existing tooling can be used and the T56 experience i s still applicable. Ratios attain- able are shown i n Table 10.

TABLE 10 RTA V/STOL POTENTIAL RATIOS USING EXISTING T56 P L A N E T ASSEMBLIES R A T I 0 3.182!1 3.258:l 3.333:l 3.414:l 3.59211 RING GEAR TEETH 105 101 100 99 97 SUN GEAR TEETH 33 3 1 30 29 27 PLANET GEAR TEETH 35 35 3 5 35 35 NUMBER OF P L A N E T S 5 4 5 4 4 MAX. POWER KI LOWATTS 10482 7878 9530 7332 686 1 (HORSEPOWER) (14,057) ( 1 0,564) ( 1 2, 780) (9, 883) (9,201) NT = 11,810 RPM = 1103 MPa = 193 MPa 'B (28,OOOPSI) 5.8.2 planet Gear Bearings The only bearings seeing an appreciable gear load in fhe planetory set are the planet bearings. The planets in the T56 reduction gear box, and thus the PD 370-25, are supported by a pair of spherical bearings.

A t 4i01 k w (5,500 horsepower) in the T% gearbox these bearings have a radiat load of 2844 rg (6270 pounds) w c h a n d , for a C130 aircraft long range mission, a cubic mean load of 1529 kg (3371 pounds). In each case the r o b f i o n a l speed i s 2916 RPM.

In the PD370-25 application the bearing twd i s 1698 kg (3744 pounds) at 6540 kw (8771 horsepower) (all engines operating) and 1 1 93 kg (2630 pounds) For the cubic mean load for a typical V/STOL mission. In the V/STOL uppl ico tion the bearing rotates a t 10,123 RPM.

The only reduction gear member having o major change from the T56 i s the planet carrier, The planet carrier has been changed from a rotufing steel member in the T56 to a stationary aluminum costing in the V/STOL application. This change in materials i s permitted since the carrier no longer carries propeller moment loads as in the T56.

The use of an aluminum cast carrier also allows the use o f cored passages for variable p i t c h fan services.

5.8.4 Overrunning Clutch Test data from the XT701 indicates that 1 t requires a bout 4.62 kw (620 horsepower) to rotate the L. P . rotor with the gas generator not operating. W i t h a n inoperable engine on the V/STOL aircraft this would represent an unacceptable loss t o the propulsion system, therefore a means of uncoupling the power turbine on an inoper- oblg engine i s required. This i s accomplished w i t h or, overrunning c l u t c h in the PD370 cnglnes, The T56 incorporates a "safety coupling,

Figure 31 , tho t uncouples the reduction

gearbox from the power secfion in the event power section failure to a l l o w the propel- ler to windmil I .

The T56 safety coupling consists of a set of h e l i c a l drive splines held in engagement by the axia I force generated by the spline h e l i x angle a n d suplernen ted by a pprelooded set of kl l e v i lie springs. When fhe h e l i c a l splines generate rufficien t a x i a l force from a n t t s t i v e torque fo balance the b l l e v i l l c p e l a d and spline fr0iction the intermediate member moves a x i a l l y as shown in Figure 21 and disengages. The ~ e l l e v i l l e springs continue fo exert an a x i a l load on the disengaged splines which llrachei I' u n t i l khe speed of the intermediate member reaches that of the inner member when engagement isagain accomplished. The spline ratcheting wears the spline ends where this type cf coupling i s overrun for long periods of time.

A centrifugal anti-ratcheting device was incorporated into the T56-A-18 gearbox safety coupl ing.

The w fef-y coupl ing, shown in Figure 22 i s comprised of these rnaior parts: o H e l i c a l SF ned pinion drive shaft, B o Outer h e l i c a l splined coupling, C o A f t h e l i c a l splined coupling, D o Belleville springs, F o Centrifugal hydraulic darns, A and E COUPLED DEGOUQLED

FIGURE 22. T56 - A18 SAFg'T'.' COUPLING

o Ball lock, G Positive torque i s transmitted through helica I spl inos on members B, C, and D. Tho o f t motion of coupling C i s limited during positive torclue trunsmission by tliu uff retaining ring aguinst member D. Negative torque, acting through the helicul splines, produccs u forward motion an coupling C. nucoupling occt~rs when the B e l l e v i lle springs F force i s exceeded by this a x i a l compcrncnk and coupling C is moved forward out of engagement with splincd member D. Hydruulic centrifugol action then holds the splines of C und D sepura tcd until speed i s almost ocluuli~ed, a i which time the parts ratchet briefly clnd rrcuuple, The hydraulic anti-ratcheking features of this safoty coupling are unique to the T56-A- 18. Thc Front and reur hydraulic durn5 ior cups) are uttcrchcd to the autcr splitled member to retain the rotation o i l . A x i a l force generated by this rotating oil i s used in the following manncL to rcduce ratcheting load a n d differential speed during decoupling operut ion.

During normal running, an annul us of o i l i s retilined by cups A and E which fill the entire coupling assembly o u t w r d f r m the inner diameters o f the cups, When decaupled, fhc aFt member D, which i s coupled to the power section, rotates a t a lower speed than outer coupling C which i s coupled to the gearbox. The viscous or "spoi ling1$ effect5 resulting from the reduced speed of the aft member D decreases the hydraulic pressure in the area of cup E. Hydraulic pressure against cup A i s maintained a t a higher level due to the higher speed of the gearbox. During flight, the speed of cup A remains unchanged. The unbalanced centrifugal head i s sufficient to keep the helical spline teeth separated on members C and D, thus eliminating ratcheting. As speeds of members C and D equalize, the hydraulic unbalance is decreased al towitig re tcheting to occur only briefly at. reduced differential speed prior to recoupling.

Thc \//STOL overrunning clutch i s identical in operation to the T56-A-18 c l u t c l ~ except in i h e source OF the o i l for anti-ratcheting. In the V / S T O ~ clutch the o i l in the annular relained by cup E w i l l be provided by the power section and the o i l retained by cup A w i l l be provided by the L/C gearbox supply. When the power section oil supply i s removed (as in the case o f a power sectian fu i lure) and the clutch uncouples, khe oil droins From cup E and h ~ e pressure generoted i n cup A drives the outer splined coupling C to the left and prevents rotcheting until ftte power section becomes operable and the o i l i s restored to cup E . When the power sectian o i l i s restored the Bellevi l le springs can move the outer coupling C to the r-ighf and re-engage the clutch.

5.8.5 Right Angle Drive in bo1.h the fixed nacel le and t i l t nacelle engines tljere i s a requirement to extract and nbsorb power from the propulsion system via a right angle gear set. While the nominal power transfcrrcd fur normrrl operation i s slight, h e powcr muy rrtach LI peuk transient condition of 2587 k i l o w t t ( 3 4 8 horsepower) with a l l engines opcrating and 5170 kilowatt (6931 hor.~upowcrj during a trunsicnt with s L/c: ct~gino inoperable, 5.8.5.1 Bevel Gcars The bevel gears for Loth the t i l t nucellc and the fixed nacelle engines are designed to provide infinite l i f e during norma 1 v/TOL operation incl dding i l s control transients, In tile proposed gear material AM5 6265 (9310 CEVM), khe design conditions for infinite l i f e ore considared to be: Crushing Stress MPa (PSI) 1724 1250,000) Bonding Stress Mpa ( P S I ) 24 1 (35,000) Pitch Line Velocity M/S ( ~ t / ~ i n ) 127 125,000) A l l o w b l e Scoring Index * O K ( O F ) 422 (30) * Assuming a maximum o i l i n fernperature of 42Z01< (300°F) with MI L-I.-23699 oil.

The choice of 9310 CEVM ma ter.ials was rnude because the slight improvement in scoritig and/or bending strength exhibited in other materials i s not f e l t to be of sufficient magnitude to overcome the extensive experience in the monuhcture of gears made of 93 10, 5.8.5.2 Lift/Cruiso and Lift Fan Bevel Gcur Commonality I t would be desirable, from a logistics point of view, to use the same bevel gear set for

the L/C engine that i s used in the l i f t fun. I f this could bc accomplished i t might

become possible to devise a Fan/gear module fha t i s common to o I I three loca tims.

In the l i l t nocel le engine the pinion turns u t power turbine speed, or 11,810 rpm, and to obtain the needed gear capacity i t i s necessary to run the pitch line velocity a t or near the 127 m/s (25,000 FPM) limit. The available envelope allowed by the engine flow poth dictates that the ratio i s near unity. This combination does not offord suFfi*+ cient capacity when used at the lower speed of the l i f t fun (3543 RPM), therefore a common set of bevel gears for the t i I t nacelle L/T, engine and the l i f t fan are not proctica I.

As the gear member rotates a t the same speed (3543 PPM) i n both the l i f t fan and the fixed nacelle lift/cruise fan gearboxes and the power requirements of both sets are similar, i f i s possible to use the same gear s ~ t i t 1 both locations. There i s a poblern with shaft angles. Most present airframe designs have a shaft angle apprwching 102' with the l i f t fan gearbox and 90a with the lift/cruise g e a r b x , I t was concluded that tho gearsets can be comrnsn .vhcn u forward cross-shuft design i s used but the gains therefrom mdr? probably be slight when compared to forces csmpro- mised elsewhere in tllc system. As this trade i s sensitive to the design of the uirfrumu, any serious study should be delayed u n t i l an uirframc luyout i s sclcctrd.

5.8.5.3 Bevel Gear Doarings The bearings incorporated i n t o the design of the lift/cruise gearbox w i l l takc advantage of the Detroit Diesel Allison high speed bearing experience and as a result, conventional b a l l und roller bearings w i l l be p r ~ d o m i l ~ o n t . t-lowcver, tapered r o l lcr bearings have been used to an u d v o n ta ye in the PD370-25E l i ft/cruisc gcclrbox where space i s limited, a The gearbox bearings in !he PD 370-25 are designed to provicle an L-I i F e in oxcel;s of the requirements of paragraph 5.2.6.

The makerials for the races and r o l l i n g elements For the ball and cylindrical roller elements w i l l be CEVM 52100 or M50 steel with the scprators for fhose bearings being silver plated one piece steel fabricated From AlSl 4340.

The materials used for the tapered roller bearings w i l l be o proprietary carburizing sfeel used by the Timken Company, 5.8.5.4 Shafting The shafting For the liff/cruise gearbox w i l l be made from Allison Specification EMS64500 nifriding steel w i t h the material properties derated by 3 sigma. The shaffing v t i l l be sized for 10,000 LCF cycles a t the maximum normal transient load. The splines on the shofting w i l l be designed to a maximum of 138 MPa (20,000 psi) crushing stress i f the spline i s fixed and maximum O F 103 MPa (15,000 psi) i f the spline are subject to relafive motion.

Splines subje-t to relative motion w i l l be hardened, ground, and lubricated, The end treatment of the cross shaft shall be such that either u dry flexible cobpling or an o i l lubricated spline can be used by the airframer for the cross drive, 5 . 9 Fun Frame and Engine Mounts 5,9,1 T i l t N a c e l l e

--

The The design of u t i l tubl{.! nocr:llc p r e s c ~ i k 5um[? unique structural design problems, t i l t fcatu1.c dicfa tes u single point m o m t nttachmenl to thc airc:raft. The axis ; I F rotation must be located u t the center lirre of the radiul drivc output shaft. W i t h thcsc requiremen t 5 in mind i t was doemod necessary thot a number sf Fan frame and engine A prc1irninu1.y weight versus stre, ~ g t h inlet housing configurations should be studied, trade-off study of t l ~ c various configu~.ations was made. Several design faccl.s were

considered as i terns to be studied, These included the type of construction -

casting versus welded fabrication, location of the radial drive output shaFt relative to the plrinctary gear set and various mcthodr of transferring the loads from the engine carcass across the fan flowpclth lo the single point mount.

A Frame consisting of a cast mount attachment waldad into a shcct metal ring, formed frwn a box section, was considered. Ihc loads fran~ tho engine carcass are conducted across the fan Plowpath through ten aercdynamic shaped struts i n t o fhc sheet metal ring and into the cast mount which i s attached to khe aircraft, Preliminary type stress and eight analysis indicated that this was n o i an e f f i c i e n t structure for the l w d s encountet-ed i n this upplica tion, Calculated lift/cruise engine weight w i t h this fan frome was 1310 kg (2889 lbs). A second opprnach u t i l i z e d a single structural pylon extending from the aircraft mount pad diroctl:, to the engine core (engine i n l e t housing).

This configurokion proved to give a lower weight to strength ratios than the previous scheme. T t ~ e single pylon presen tcd a sizeable local uerodynornic blockage immedio tely u f t of the fan rotor, This undesirable feuture dictated the eliminution of this configura- tion as at1 optimum design, C u l c v l n t c d lift/cruisc. erigine weight w i t h this design #as 1284 kg (2830 Ibs). A third configuro!;:>n utilizes u full r i n g t y p e cmstrucfisn w i t h ten aerodynamic struts which conduct the !cads ucross :I,rz fan f l o w p t h , similur to the first kame described. The r i n g d e t c i l construction i s quite different, four stiff rings form the corners of the box section, The rings are tied together by thin sheet meto I cylinders a t the inner und outer circumfsrcnce orid flat plotcs nt thc fare und a f t ends, In addition, hollow stiffening pact; span between the rings i n a longitudinal and rc~dial direction to prevent the rings frorn rolling. Thc thiclcness of the rings und the shear webs are increased us the l a d s ore gathered i n approaching the mounting pad, thus providing a more favorable weigh!. to strength rtitio. T h i s sclieme 0 1 5 0 r ~ t i l i r e s a welded engine inlet housing, which o ffords a17 additi onnl weight- s~ ving. Ca lcula bod lift,'cruisc engine weight with this design i s 1196 kg (2637 pounds). T h i s fan frome and engine i n l e t housing was iudged to be the aptinivrn cunfigural.ion after ~ r e l i r n i n a r y weight and stress analyses and i s used an the PD370-25E engine.

In order to perform the prel iminory fens:bil ; t i and trude-off studies, the fa1 lalving loading imposed by b e i n g a irfromt. tequiremrnts was used instead of the maneuver l a d i n g criterio shown in Figurr 13. Boeinglr, recllJll~nicr~ts were considc>ted to be more relllistic and helpful in rninirnizitlg the structure's weight.

~ ~ l t i ~ ~ ! e I m d i n g criter iil were used to r ~ r t out .r~aximurn t e r t i c c J l c l r l r l qyroscopir corrditiitn5 since thrust can eithcr increase or d r c t t * ~ ~ n . ~ O ~ L I lmds depending on li.c i,ucellc othchment,.

Positive values ore for up loading and n r g ~ t i , ~ ~ .<oIucs rot dr3wrl Iwding.

bC Ennines in Convent~onal Flight Position - - -. - - -- - . . .

14.09 (accc!era tion of gravity) -+ 1 .5 T Vertical: MAX (Maximum ~ a c e l le Thrust) 4 , Og -3.33 + 1 . 5 TMX - 3 . k Lu teru I: i3.0g Longitudinal : 9 g (crash )

Thrust: 1 . 5TMx i 1 -59 Vertical

Equivalent to 5toppillg rotating mass in 0.6 Engine Seizure: secwlds 6022 h-rn (53,300 in- Ib) 12.5 ~ a d / r e c yaw Gyroscopic: 1.5 TMAX - + 1-59 ver tica I

*". 5 ~ad/sec. pitch + 1.5 T MAX S 4*59

v e r t i c a l Ennines in Vertical Flight Position Vertical: 5.Og -1- 1 . 5 T MAX 5. Og -3.09 1 . . 5 T M A X -3,Og Lcl fera I : Longitudinal : Thrust: 1 STwX -t 1 - 5 9 vertical Engine Seizure: . Equivalent f o slopping rotating mass in 0.6 seconds 6022N-m(53,300 in-l b)

Gyroscopic : $ 2 . 5 ~ad/sec yaw + 1 . 5Twx t 1.5g vertical

t 1.5g v e r t i c a l 12.5 ~od/sec pitch -1 1 .5T MAX 6 1 The limit load design conditions are 2/3 of these valucs, The fun case and engine inlet l ~ o u s i n ~ are T i 6-4 welded Fabl-icotions. The Following material properties and criteria were used:

Mat1 - T i 6-4 (Annealed)

Ul Firnote - 3 F 1 1 422" K (300°F) 758 Mpa ( 1 10,000 psi)

Fan Frame -

Web - a l l o w elastic buck1 ing up to y i e l d

a)

Rings - to y i e l d w i t h no buckling

b)

Stiffners - to y i e l d w i t h no buckling

c)

Struts - no buckling or y i e l d

d)

Inlet Housing - no y i e l d

5.9,2 Fixed Nacelle The fixed nacelle fan frame provides the support F o r the Fan shroud and the Fun 'rl l e t cowl. The inner bypass duct is also mounfed on the Fan frame. Aerodynamic and maneuver loads on the outer bypass flowpath members ore conducted actoss the bypass annulus by ten aerodynamic shaped strut5 t o an inner structural r i n g which i s moun t a l on the inlet housing a t Fore und aft flanges. The radial drive shaft i s conducted across the flawpath w i t h i n one of the struts as are khe necessary engine services. The fan tip treatrnenf ring i s an aluminum honeycomb fubrication. The outer and inner b y p s s Flowpths and struts are fabricated as a single aluminum weldmenf, A saddle type s t r u c t ~ ~ r e mounted between the outer r i n g Flanges provides an outbwr-d bearing support for the radial drive shaft.

The fan kame i s mounted on the engine i n l e t houjing. This i n l e t housing i s an aluminum casting which incorporates the flowpukh entry to the high pressure compressor, compressor rotor front bearing support, compressor i n l c t guide vanes, and front engine mount pads, Provision i s also mode to mount the fan Beta regulators. The i n l e t housing i s designed such that i t i s capable of being mounted as either a right hand or left hand installafion.

Preliminary stress analysis has been performed i n order to determine the feasibility of the scheme and to size components for engine weight calculations. The flight maneuver l&ds deFined in Figure 13 were used in the analysis. The ~ 0 3 7 0 - 2 5 4 lift/cruise engine with this fan frame weifjhs 1066 kg (2351 pounds), 5. I0 Propulsion System Control Studies Introduction The RTA controls effort was directed at the conceptual clesign of the RTA propulsion control system. The program apptoach wos to analyze tnie control problem and meet with Hamilton Standard, NASA LeRC, Bceing and McDonnell Douglas fo establish operational requirements. This i n forma tion was used f o formulate a problem stu temen t ' upon which the conceptual design was based, A basic constraint in the design was the maximum use o f X1'701 control system components t o minimize development cost. However, the relationship O F the propulsion and Flight control systems in vertical and transition modes necessitates a similar degree of reliability. In order to achieve high operational reliobil ity, the propul5ion control incorporates triple redundant, digi to l logic. The XT701 hydro- mechanical control i s retained w i t h minor modifications.

Control Problem Two lift/crulse angines with integral, vcriable pitch fans, a Fuselage mounted turbo- shaft engine, a clutched l i f f fan and appropriate power transmission system comprise

the RTA propulsion system as depicted i n Figure 23 . Controlled variables and

con trolling parameters ore shown in Table 1 1 .

Table 11 Control led Variables Control led Variable Symbol Contro11ing Parameter

--

Fngine fuel flow Power level, fan pitch, engine limits, compressor W~ discharge pressure Compressor variable C V G High pressure robor speed, inlei temperature Geometry

Fan Pifch 6 Flight con trot input

Lift fan clutch - -- Transition input

~iFt/cruise nozzle *N Mach number area Water alcohol Engine out Pl/c fan .c C lift/creise fan

a I

flight controller

outputs 7 digital eearbax

- -

I I 1 contrnller 1 . - I eneine I I 17 fan 1

a t e n g i ~ e feedback signals I

I Uy digital HPC 3 4 right engine & I

controlier 1--------*&A lift/cruise fan

- 1 I

The RTA propulsion control system performs the fol lowing functions: Engine starting

Engine Iimi ting - pressure, speed, temperature

Stubilire system fan pitch control Clutch control Load sharing Airframe interface Condition monitoring Engine out operation Engine starting i s initiated upon request from the flight deck. The control rnmi tors engine speed and temperatures during the start sequence and aborts the start i f ignition and normal run characteristics are not achieved. This logic i s used on the XT701 and w i l l be continued.

The control system limits engine pressure, speed and tempera ture a t maximum power conditions t o prevent the exceedance of engine operating limits. T h i s i s required to eliminate engine overstress condi tims with attendant rapid loss of life.

Past experience w i t h shaft and rotor systems on helicopters has shown that i t w i l l be necessary to compensate the system in order to achieve proper transient response and eliminn te engine supported osci lla fions a t the propulsion system no tura l frequency.

This i s achieved through phase and gain compensation in the control.

The Hamil ton Standard Division of United Technologies (HS) fan pitch actuator i s a

dual redundant system, ref. Figure 24 , that incorporates t.wo electrohydraulic

servovalves, EHV, with position f e e d k c k on the spools, o solenoid controlled bypass valve, a pitch changing mechanism and a three winding linear variable differential transformer, LVDT, tho t senses blade pifch.

Triple redundancy on the servova lve i s achieved by o mode! in the electronics. 7'0 illustrate operation assume a Fan pitch, @ , command from the flight control system.

The propulsion control w i l l send a current to the trh.:, EHV'5 and the n~odel. The two EHV LVDT's and the modeled EHV and LVDT ~vill be compared. I F fhe three do not agrec, the bypass valve w i l l switch the malfunctioning EHV o f f line and transmit a

b

TO REDUNLAWT SYSTEM

t7 I AIRCRAFT

CONTROL SYSTEM BLADE

I

ANGLE POSITION LVDT (3 WINDINGS) TO Pt-rCH

-

CHNC. ACT, 1 ! BYPASS

7 VALVE

BYPASS IND SWITCH A

I SOLENOID V A L V E (N.O.) I

FIGURE 24. PITCH CHANGF CONTROL SCHEMATIC warning discrete. The flow through the EHV w i l l cause the pitch mechanism to drive which w i l l be sensed by the blade angle position LVDT's, The propulsion control w i l l sense this movement and compare i t to the flight control commanded @ .

The propulsion control system provides compensation and condition monitoring for this position conirol loop.

Power f l o ~ ~ to the l i f t fan IS controlled by a clutch. Although definition of l i f t fan on/off operation control i s not complete, it. could be a Function of transition condition, aircraft Mach number, etc. Some modulation of clutch actuation pressure w i l l be necessary to reduce shock l a d s on the drive system, This clutch actuation pressure mowla tion and lockup logic w i l l be incorporated in the propulsion control system, The propulsion con fro1 system w i l l operate the three engines a t approx;mately the same power level to bolance engine loads. The XT701 contlol system balanced engine l a d s on output torque; however, the XT701 torquemeter was eliminctted to reduce engine jength, The engines w i l l be matched on turbine tern? 're. For an operational system, the torquemeter must be reconsidered as i t malces .

s s i ble to operate the engines to a measured rother than an implied output Faran .,, and i t i s a valuable condition monitoring tool. I t permits a very close match on power, but tho t i s not on RTA requirement.

The propulsion control i s the airframe/propulsion system interface unit, Boeing and McDonnell Douglas plan triple redundant, digital, fly-by-wire flight control systems.

Therefore, a l l propulsion control inputs - fan, pitch, power level and transition -

w i l l be triplicated. The propulsion control w i l l incorporate majority voting on these inputs.

Although each engine incorporates an alternator for control electronic power, aircraft critical bus power w i l l be supplied to the control for backup. The airframe hydraulic system w i l l supply power to the fan pitch change mechanism.

Tho details of the inkerface have not been worked out as cwrceptutrl designs have not progressed that far. Analog and digital interfaces transmitted electrically or optically have been considered .

Condition monitoring i s incorporated t.0 reduce the possibility of sudden power losses, The performance of the three engines i s monifored and compared to skandard set of performance conditions, In an operational rvstem, this would be expanded i n diagnostic scope and depreciation compensation would be implemented. The condition monitoring system i s used to detect an engine out condition and trigger a contingency opera fion when required.

Engine out operation requires increased powor from the remaining engines, The propul- sion control w i l l increase fuel flow u n t i l thc required power level i s obtained consistent with engine limits. Where necessary, a water alcohol injection system could be used to achieve an increased power level For engine out vcrkical operation. Initiation o f water alcohol would be automutic and under propulsion system control. A ra:aylf.

indication woutd be transmitted to the oirc;roft.

Control M o d e Logic

-

Figure 25 depicts the control activities necessary for each ol' the three engines on the V/STOL RTA application, Interfacing wikh the a i r c r o f t Flight controller i s a d i g i t a l controller on each engine, W i t h outputs From the flight controller arid engine feed- k c k signals, ecch digitzll contrcrlter i s rest>onsiblc for its own engine control inputs, those being fuel flow and compressor variable geometry.

In addition the two engines w i f h l i f t / c r ~ ~ i s e h n s r e l y on their respective digital con- trollers to posifion the fan variable geometry position commanded by the flighf con- kroller and to schodule the corresponding duct nozzle areas. A data bus between the fhree d i g i t a l controllers permits p:~&itioning of the lit7 fan variable geometry as w e l l 0 5 determining load sharing and wo ter/alcohol injection activities. Condition mon itoriny of engine signo Is provides additional fa ilsafe opera tion.

Fuel Flow Control

--

The fuel flaw required tc run each of the three engines i s provided by a hydromechanical u n i t on each engine interfbciny w i t h the d i g i t a l controller. The hydromechanical unit i s the orre used on the XT701, T h i s d i g i fa1 controller relics on the following engine feed- back signa15: Gasifier rotor speed (hi ) G Far1 rotor speed (N F ) Fan i n l e t temperature (FIT) Compressor inlet temperahre (C IT) Fan inlet pressure (FIP) Compressor in let pressure (C IP) Compressor discharge pressure (CQP) '!iith these a n d flight conl~roller output signals of power lever angle f P L A ) , aircraft M a c h number (MNAiC), and ccndition lever IC Lj, the fuel f l a w required for ongine operation i s obtuined.

FIGURE 25. R T A C O N 1 RQL ACTIVITIES 69 Fan rotor speed i s controlled on a closed loop compensated proportimul plus integral basis using inputs of P L A and ambient conditions tMN FIP, and FIT). I t i s A desired to operafe the Fan at a constant speed in the vettlcal mode. In a cruise condition i f is necessary to coordinate the fan speed schedule with fan variable geometry position to yield optimum specific fuel cmsumption.

The error it1 Fan speed i s used to adjust the gasifier set. speed (NG ) for each 5 ET engine. i s the sum o f this sdiustrnent and t:vo scheduled values, one NG valve N G ~ ~ ~

scheduled on condition lever and the other scheduled P L A . Another delta factor to

i s made to assure equal sharing of the load between the three engines. The

YG2v oa s wring option w i l l be accornplishsd by establishing the highest of the three

gasifier speeds as a reference value und increasing the other two gasifiers to approximo tely match this reference speed. The resulting NG from these previous actions, after being compared to the maximum and minimum gasifier speeds, i s pssed on to the fuel flow modulation loop.

The governor fuel Flow ( W F ) i s controlled on a proportional p!u5 integral b a s i s using NG A N D NG ET. ~%%%dule of acceleration fuel flow ar o function of NG a n d CIT arrurner the 3ransient 1 irniting responsibilities. After 7 -crier of comparisons X , and WFMIN, the appropriate fuel flow a m m ~ WFGOV, WF%cCt WFpECr WF voluc ir chosen. Fur! er adiur rnents t o w e fuel flaw signal aid engine starting characteristics.

Compreesor Variable Geometry Con fro1 f ositioning of the compressor variable geometry ICVG) i s hondled by tha hydromechanical unit. The desired CVG setting i s scheduled on gasifier speed biasei by CIT, The scheduled C V G setting i s compred to an LVDT feedback to assure positioning accuracy.

Li ft/Cruise Fan variable Geometry Control The determination of the b n variable geometry setting ( 8 ) on each of the two lift/cruise fans i s the responsibility of the flight controller in t.he vertical mode. By cornporing the reference signa l to the feedback assures correct pasi tioning.

In aerodynamic flight the condition lever Functions as a normal flight deck input to the sonfro1 system. The system w i I ! establish the proper fan speed and @ rela tionships for best propulsim afficiency, D u c t Nozzle Control The duct nozzle area on each O F the two li ft/cruise engines is to be scheduled versus oircraft Mach number. Positioning to achieve this proper nozzle area is the respon- sibility o f the digifal controller by comparison o f fhe scheduled urea with a feedback signo I.

Lift Fan Variable Geometry C m trol The l i f t fan variable geometry setting i s determined by the Flight controller and positioned in the same manner as were the 6's for the two lift/cruisc engines, The possibility exists for automating water/alcohol in icction in the case of on engine out condition, If only two engines ure operable and a particvtar high thrust paint cannot be achieved, injection of a water/alcohol mixture would ullow incwased power for o short duration.

Condition Monitoring A l l signals fed back From the engine and a l l inputs from the Flight controller are condition insure reasonableness, logic to detect monitored. This includes range and rate chocks an engine out condition, and siurt sequencing of the engine.

Clutch Command When the transition i s mode from vertical flight to horizontal flight, i t i s necessary to disengage the front l i f t fan from the gearbox drive, This clutch command i s assumed by the digital controller. 70 prevent sudden disengagement, and therefore possible shock Iwding, the clutch command i s rote lirnited.

Mechanization In the course of the RTA propulsion control system study the mechanization evolved considerably from the XT701. The XT701 system incorporates a hydromechan ica l control on HP rotor speed and compressor variable geometry. This unit has n c direct interface w i f h the aircraft and receives its inputs From h e Engine Electronic Control, EEC, a single channel electronic assembly.

An electronic power management control i s incorporated to match the three engines' power output and provide isochronous rotor speed govc-ning.

After meetings wit11 the airframe manufacturers and unalysis of the problem, i t was decided that khe XT701 electronics must be replaced to accomplish the added fan pitch control task and to achieve fail safe operation consistent with that of the flight control sys tern.

In order fo achieve this, a triply redundani, digital engine controller i s required. A Foil operate requirement in the vertical mode i s necessary because of potential damage to khe RTA. I t would not be acceptable for a single electronic failure to cause on engine shutdown or loss of fun pitch control.

The triply redundant system oehicves tlio fail operate r e q u i r e m ~ t ~ t . A d i g i b l system i s

used because O F i t s more powerful logic and computational Funcr'ons as wall as con-

tinucrlly decreasing cost, The XT701 hydro unrts arc modified to incorporulc new CVG and Fuel schsdu:os and to increase the authority o f fhc fwin input. This provides satisfactory tolerance la servo- valve failure, The hydro utii t itself has demonstrated very high r e l i a b i l i t y so that : l o backup Fuel metering system i s planried, 6,O Program Plclrr 6.1 Introduction Detroit Diesel Allison @CIA) has prepared a program plan far the development of variable pitch, shaft driven, l ift/cruise propulsion units and l i f t fun units for a V/STOL research and technology aircraft. The propulsion system consists o f tlhe integration of a Mumilton Standurd vur.iahle pitch fan and u Detroit Dicsel Allison XT701-AD-700 turboshoft engine.

6.2 Assumptions The Following i s the list of assumptions made while developing the program plan.

A .

The five new XT701-AD-700 engines owned by the U.S. t y rmy would Le available for use in the program, B, DDA w i l l subcontract to Hamilton Standard the definition, design and hardware of the variable p i t c h l i f t b n assembly a n d he lift/cruise variable pitch Fan rotor asscmbty.

C. The program plan i s based on the program starting a t thc. conclusion o f the currently defined Phase I preliminary design.

D, A m e aircrafb program w i l l be the baseline.

E. The a i r c i a f t w i l l h a ~ e 3 engines and 3 funs.

The baseline engine used i n the plan i s the standard XT701-AD-700 engine w i t h intermediate power as the highest rating, As direcfcd by NASA, u two a i r c r o i t program and three a I terna te upra ted engine configuration programs were evu I w f e d . These are preseii t r d as "dc I tos " to the bas el in^.

The engine development plan i s divided in to four specific tasks: Design, Fabrico tion, Testing and Support. Each tusk i s described below, Design The design task is divided into two subtasks.

First i s the propulsion system design through detni!ed manubcturing drawings. Subtask two i s engine design and control system design followup. T h i s subbsk would provide the manpower required during the period O F hardwarc fobricotion, romponenf testing and engine testing.

f:ivc futlclions ale ir1r:luduci i t \ tile fubriciitiorl r3f t t ~ c l i f t 'GI t ~ i w , variable pitch tu~boftrn plopulsion ;ystern, i h e tci:li i s t h r . dr:iyn r ~ n d fubl ic;~tiurl L)! (1 Iift,'cruisc cnginc mockup, Mtrtrubcturirry toeling h h r icution urid/'or plocutcrncnt i; i r r ~ l u d c d . Fob1 ication c l i o require5 manutactur ~ n g Iioi~orr onrl mufu~ io I control suppal t . Lastly marlpower and rnotcrial requirccl to fubrica te conipont~nt~ und l o rnadify existing XT7U1 otlgines clre included, N c w compononb rcquirc:d i ~ r the ~ift,/cruise turboliln engine include: o Engine U L C E ? ' J S L ) ~ ~ drivc gectrs o O i l system for vcrticol rrttitudc o Fun drivc lubricofion system o Mount ;tru(;turfa o ~ a d i f ; c d hydrornc~hunicul control r3 M d i C i e d eIc.:trr;nic control N e w components I equil.cd for the turbwshnft centcr crtgir,e to makc i t inrerchangr.at.le w i t h Iift/cruisc engine include: o Modified hydromechanical control o Modified electronic contrr,I Testing Included in h i s task i s the design und ptocurcmer,t of ccrnpor,cnt te:t equiprnenf and the following component tests: Components Test

-

Static Deftecrion lrllct Housing and Engine Muin Drive Gears Vibration and DeFIcztion Control Compo~en t s FIG.# Pcncll and Electronics Checkout 011 Pump Calibration Engine test equipment must ulso be designed and procurcd in order to run the follow! .

tesbs: Est, Hours E s t . Builds Special Conditions Type

-

- - - -

Inlet Distortion 15 2 Rammed & bleated, BIadc !nstrurnenla tion Verlicol Lube Sysiern Testing 20 4 Ambient, Power Extraction Conirol System Evaluation 10 2 Ambient, Power Extraction Gas Ge:..t:ratar In let Survey 10 1 Ambient, Power Extracfion Star ring Tests 10 2 Ambienl Endirrance Testing 120 4 Ambient, Power Extraction Flight Clearance Test 60 1 Ambient, Powst. Extraction Two engines would be used in the developmenr testing. E r i 3 i n ~ numbel, one would have 1 1 5 hours O F festing including the 60 hour flight clearance. Engine number two would hove 130 hours of tesfing including 100 hours of endurance fo be completed beFore the flight clearance test.

Also included in ihe tusk is the buildup, acceptance tc:t, partial teardown and inspection, reassembly and 'inul run, and shipmerrr of lift/'cruisc t . u ~ , ~ f a n engines and rurboshaft engines for use In the Research 2 n d Technology AircraFt.

This bsk iccludes the ovel*alI management of the program by the Chief Project Engineer, scl~cdule and budge1 control and .other administrative effort5 by various service groups.

A, lso included is the ~mnpower grid rnaferial seq red fo design and procure engine ground support equipment and to provide training mu terials and user classes. Project engineering, design engineering, test engineering and other teclinicui s z r * ~ i c e support begins after the deiivery of flight hardware and w i l l continue for a projected 12 months while l:,e airfrumer completes ground and flighk festr o f the RTA .

Upra ted Engine Configurations Considered - - - - - Level I . S tatidard XT701-,AD-700 w i t h water,/aIcohol iniection.

XT701 wibh a short term cantingcncy rating and woter/alcahol injection.

Levc! 2 .

Contirigency achieved by throftle bending and rework of existing engine hardware.

XT701 with a contingmcy rating and wc!e:,/alcohol injection. Major L e v e l 3.

in terna I modifications and new hurdwc:re for angin65 are ~Aequired, The !an dovelor~ment plan i s also divided i ~ t o Four specific tasks: Design, Manufacturing, Developmc;nt and Field Suppork. The uc.tivi tics and participating disciplines of tasks are out1 ined below: Design

--

o Design Engineering o Drafting o Project Engirieering o A s dynamic Design e Dy mic Analysis o Struct,~ml Analysis o Program !q\anagerneni.

Mur~!~!bctur ing - -----.

o P i nducticm Control a P r o ~ r l ? m Mar !gemen t o Tcoiing Dcrign 8, Mr~r:ufocture o Tesl Hardware Comporien t Test Hardware Equivalent uf I Fan Deve lopmen t and Testing Test Equipment Project Engineering Design Support Program Management Component Tests ~ t r u c t u r a l Tests lade Rotor Disk Mount Vibration

Subsystem Tests - Total of 250 hours of testing

Whirl Tiltlube

Lift Fan System Test - Totu I of 200 hours of te-ting

Field Support DDA ~ift/craise Engine Testing Aircraft Ground Test.; Aircraft Flight Test 6.4 Engirie and Hardware Requirements as For a one aircraft program, the f i v e clvoilable XT701-AD-700 engines would be follows:

Engine I - A/c Engine

Engine II - A/c Engine

Engine I II - A/c Engine

Engine IV - Development Engine

Engine V - Development Enginc

The Follovring hardware must be obtained For a m e aircroft program: 2 Sefs of L/c engine development hardware 2 Sets of L/C engine hardware F o r A/c useage 2 Sets of L/C engine hardware f o r refurbishing development engines to spre status 1 Set of hardware for modification of XT701 center engine 1 Lift Fan for ~evctopmorr t Testing L i f t Fan Hardware For Refurbishing Tesf Fan 2 ~ i f t / ~ , r u i s e Fan:, for ~evelopment. Testing ~ i f t / ~ r u i s e Fan Hardwore for Refurbishing Test Fans 1 Lift Fan Assembly for A/C Useage 2 ~ i f t / ~ r u i s e Fan Rotor Assemblies for A/C Useage A two aircraft program would require the fol towing engines:

Engine I - A/C ft I Engine

Engine II - A/c # I Engine Engine I il - A/c # I Engine

Engine I V - Developman t Engine and A/c # Z Engine

7 8

Engine V - Development Engine and S p r e

Engine VI - A/C H2 Engine (to be manufactured)

Engine VII - A/c H2 Engine (to be manufactured)

In odditior; to m n u f a c t u r i n g t w o additional engines (one I , & , engine and one center engine), the r'ollowing ha~.dwal.e must be obtained for u two aircraft program:

2 Sets o f L/c Engine Development Hardware

2 Sets of L/C Engine Hardware for A/'C Useage 1 Set o f L/C Engine Hardware t o ~ e f u r b i s h One Devetoprnenf Engine for A/C l J s e a g ~ i Set of L/C Engine Hardware for Refurbistring One O e v ~ l o ~ r n e n t Engine to cr Spare 2 Sets o f Hardware for Modificafion of XT701 Center Engines 1 L i f t Far1 for Deveiopment Tesfing L i f t Fcn Hardwars For Refurbishing Test Fan 2 l . i f t / ~ r u i s e Fans for Development Testing ~ i f t / ~ r u i s e Fan Hardware for Refurbishing Test Fans 2 L i f t Fan Assemblies for A/C Useage 4 ~ i f t / ~ r u i s e Fan Rotor Assemblies for A/c Useage 6.5 Schedule The schedule of the development program i s sliown i n Figure 26.The schedule i s for the "ba5elinet1 .qe aircra F t program. Two liFk/cruise turbofan engines, one turboshaft engine and ,,ne lift. fan would be delivered to the oirfromer after 30 months and 245 hours of engine testing and 680 hours o f fan testing. Similar camponen ts For the second aircraft would be available after 33 rnonfhs arrd the spares avails ble after 35 months.

Total length of the baseline program i s 42 rnat~ths. Twelve months of technical support w i l l be provided to the RTA builder during the ground and flight test prc:rams.

I .I DESIGN 1 .Z DESIGN FOLLOWUP 2.1 MCCKUP DES & F A 0 2.2 TOOL DES. L FA0 2 . 3 COMFDN & ENG. f A B 2.4 MFG LIAISON 3.1 COMP. TEST EQUIP. FA3 3.2 CDMO. TESTING '3.3 ENG. TEST EQUIP. DE5IG 3.4 X T 7 0 1 ENG. TEST 3-5 t l F T FAN TEST f1 *c:

DEVELOPMENT 7 a & & Y 2

3.6 L./C F A N OELiVERY HAEDI~ARE I START .P T C C t A P L E T E FLIGHT CLEAPANCE TEST 3.7 L ' C ENG. TESTING

-

3 . 8 ACCEPT. TEST 8 DEL.

HA3Dt1JARE DELIVERY 4,'C '1 A:C '2 SPARES A . I P R 3 J . MAkAGEMENT 4.2 GSE HWO

I

4 . 3 TRAiNltiG 4.4 FLIGHT TEST SUPP.

I

FIGURE&.RTA PR0PULSIOt-I SYSTEM DEVELOPMENT PLAN SCHEDIJLE W i t h uprated engine level 1, the delivery limes would also be 30, 33 and 35 months.

A l l engine testing would include water/alcohol iniection and ten additional hours of engine testing would be required.

F i f t y additions I hours of engine testing from the buseline would be required on upra tcd engine levef 2 and the delivery dates would be increased to 32, 35 and 37 months for one and two aircraft programs.

A total of 355 engine test hours would be necessary before delivery of uprated engine level 3. Delivery dates would be 36, 39 and 41 months from go-ahead.

6.6 Program Options Detroit Diesel A l l ison believes that the power management control of fhe RTA should be supplied with the propulsion system but the design, fabrication and test elements were not included in the progrum plan due to the current definition of propulsion sysfern component suppliers (i. e ., power management control to be airframer supplied).

Component testing of new hardwo-e can add r e l i a b i l i b and shorten development On4 has identified the l i ft/cruise gearbox as a component where addirionot schedules.

testing would be beneficial. 200 hours of back to back testing with the equipment shown in Figure 27 i s suggested.

b e i n g Aerospace Company and McDonnell Aircraft Company have both defined iron- bird tests or ground tie down tests for RTA propulsion system hardware before the first flight of the research and technology aircraft. The currently defined program does not have hardware identified for those tests. As a minimum 3 engines and 3 Fons would require overhauling back to zero time for flight test use. A maximum addition to the program would be a requirement to rnonubcfure three additional engines and fans for the ground testing.

An alternative would be to use the Army's 501M62B engines from the Boeing Vertol dynamic system test rig. Engines would likely require overhaul and would require modification to accept a liff/cruise fan. However, the turbofan engine wauld not be.

flight qualified hardware.

The aircraft con:ractor's testing schedules are incompatible wifh the develol~rl~ent.

schedule shown i n figure 2 6 and would h v e to be worked out during the prcr.,)sol phase of tilo program.

TEST EQ'JI?kENT GEAR B3X (CSES ENGIKE GEARS)

-

-

-

- . -

\ - I

-

- -. TORQUE a / NOTOR =PLIER

\ \\

lf /

\ ! t . // , 1

-- A - TEST ZQUIPbENT GEAX BOX I TOXCOE \\ i m s a i// FIGURE 27. EACK f O M C K TEST RIG The work under this contract hos defined the V/STOL Research and Technology A i r c r a f t turboshoft propulsion system. Two lift/cruise turbofan engines, one turboshaft engine and a l i f t fan interconnected w i t h shafting and a combiner gearbox form the syslem.

Conclusior~s derived from this program are: XT701-AD-700 turboshaft engines a n d ~ a r n i l t o n Standard 1 . 5 7 metre 1.

(62 inch) diameter fctnj provide fhe required VTOL and cruise thrust.

Modifications required to integrate a fan and engine a n d produce a 2, VTOL engine are straightforward applications of exisking technology.

3 . Production hardware w i l l be used in the L,/C reduclior~ gears. This w i l l reduce developrtent cosfs and add reliability to the program.

4.

'The propulsion system card deck allows airframe cont-racfors and other users t o generate performance formally provided in fabular form.

5, Propulsion systern development w i l l be paced by the l i f t and l ift/cruise fans.

8,O APPENDICES

APPENDIX A INTERFACE DEFINITION FOR NAsA/NAvY V/STOL RESEARCH A N D TECHNOLOGY AIRCRAFT PROPULSION SYSTEM HAMILTON STANDARD DIVISION U N I T E D TECHNOLOGIES CORPCRATIOI.4 DETROIT DIESEL ALLISOI-l DIVISI(SPJ GENERAL. MOTORS CORPOFtATION JULY 1976 REVISION " 8 " JANUARY 1977 APPROVED: HAhSlLTOt,.4 STANDARD I. INTRODUclTi('l1.I GENERAL 1SET:INll'r DNS A N D R ~ S P U I ~ S I Rt l.JTIES 1 1 .

1 1 1 . REFTRENCt' DRA'NINCS, DOCLIMENTS AI'ID SPECIFICATIONS IV INTERFACE DEFlP4ITICf. I5 I . INTRODlJf. TION (REVISION B ) The Hamilton Standard (HS) Division of Unitad T c c l ~ n a l o ~ i e s Carporotion and the Detroit Diesel Allison (DDA) Division o F Gencral M o b s C o r p o r ~ t ~ o n arc engaged in concept definition studies of lift/crui:c propulsion syskcrns F o r n NASA/Novy V/STOL research aircraft under NASA contracts NAS3- 19414 a n d MS3-2flO33 with HS, and NAS3-20034 with DUA. Thuse contracts require that t l ~ c intcrf-aces bctwccn the HS fan componcnfs and the DDA cnginc carnponent.5 LC defined. This docunlertt defines the interface details which have h e n iden tiFied to date arid the responsibility for components resull ing from these interface dotoi Is. I The refinement of details of thc inkerfaccs between tho HS and UDA components defined herein w i l l be recordecl i n revisions 1.0 this intcrfocc dcfirlibion document as the program proqresses. A n y majar interfcrce changes From this docur,~rsnt shall be identified in writing to tIie NASA Project Manager.

fan interbcc dc(iniiion rjglecil to Lctsfeen t h t ? oirbsrnc canhactorr and T\,e Hamilton Standard as port of HS's work under Contract NAS 3-20033 i r incor- pomtctl into this document as Addendurr. - .

Thc V/STO~. propulsion systt?rri consists of Iwv 111rIjnfori rtrrgin~~s, a ~,ernote l i f t f a n and thc as:ocicltt>d t]curing and r1-i litinr, rr!ql,irr5d 1.o cnul?ic t h ~ l q r ~ c : ~ ~ l p r ) l ) f ! n t ~ , Iden tica 1 voriohft. pi tcl.1 fc~n rciturs me: 1 ) $:.-I in rtir turLcjF.rn t'rtcjinc?r, r ~ t l i l kl~r.: l i f t kin. This dacumtvt deals w i t h the i t l k i ~ r krl:c bcf\;:.i!n klorr~iIlur~ S ta~~dard v r ; iablc pitch k . 1 rofor and ftlc Alliscu t u r l ~ t ~ : t ~ o i t r?riqivc, grorlu:: a!m rnb,ly urld Fon Frr.~me rind case, which togerher. f'r.>rrti !.Ire ~,.rlloSr~ri cnginc t 1 1 1 1.41 ~ c . cu~ripcs~it?t~ t s otrtl file rcsul tiny complete turbobn DDA i s rr?spcsnsihlc: for tllc 1 l u n ~ i Itnr, Strlnrk .ri i: !v:l!otl.:itlII.~ for tlii: : i ~ , ! . i i i i st.1~11' v~riuhic, pitch fsln and cr,gine, the actuufors u r ~ t i corltr.r51s ( 1 , . ;. i ~ f e r l v ~ i t l i ! . : ~ ~ I ~ . ~ c rtro.;c.! c - r \ t . This fat) req~onsibility I . . : r ; 4 ~ ~ ~ i - i c ~ 'n id upt:ra tirig envelope, and

includes defining tt~r:: uvcrk, 1 , . I ~ I r 8 i [ l ! j r 3

providing fhe a r , i ~ ~ d ~ t ~ t ~ r r i i t - l i . ~ ' i r i i l : : . , r : fc~r I . ~ J : ~ . I : ~ I ~ < I clrlrt ~~~~~~~~~~~~~y canpot~enh wi!l.riri tllu stcgc:, Wanliltr,~,~ Storldll~ .J i: crl )a r ~~',~,url:iL*It: i'or J I I ri;cchariic,al :o~nponcnfs and functions of the iun i ~ f 1 - 1 1 . . . i i l t r : ~ I . t l ; , ,.lr~l-j V , i t I f ! l e r ~ t v i: r~>r>rdi~\c:i.e the i 3 2 i t l - r n ~ ' ~ h a n i c a l desiyr,. DDA w i I l or! r e!;pc;lr,si kv4-: !'w file ~oc~cl;anico! dcasign of the stirlioriory fan carnpol>cnts since fjlc:rtb %:#it l i:Ls ;~lttlgrutt;d hi t o ifit: t urkruf~n cndi~.le Fnrwurd Frame sfruclurc, Signals for t i l ~ p o s i t i n i ~ i r ~ g i r f t ! l . ~ iT:rl b l r ~ r ; c . r,i;~y tonl, ft.c,rn ti.1~. cnginc fue! contra] and the aircraft f j i y ! ~ t r t l r i t t l 1 . I { i ! ~ ~ l l t ~ , i \ ~ t r l r i I l r r f i 5 r~\1.11~7,15ible Far the componenls required to condition t l l r :c8 r i g * I ; . ( : ( orl ;eft t l ~ c l r ~ , i r ~ t ~ L i r ~ d o t3:1~lcsetkings or1 the variable L j l r . 4 + ) i l l ! r l i > ; ( I t s f l l r l pi~,vl'r ; r ? ttrr. f o r m .-' l t y d r ~ l r i i t pressure and Flow pitch fat, rotrlr.

for mi t t l t . [it, ~ Y C ~ C C . I ~ ~ > T ~ . l\tt.' gcnrin.1, lut~ricr~til-kt,, acct.scnry drives arid aircraff i t r - I - - 5 r - PI^ I y f A . (7v1.1.al l liFt/cruisc turbofan engine +rrfoit,iarlcc i s lhtt rist>nr~r it.1 l i r a , a i @PA.

Ill. REFERENCE DWWINGS, - DOCUMENTS A N D SPEClFlCATlONS (REVIS I O N B )

The Following drawings define I he components and the as5ociaied

DRAWINGS -

interfaces which arc the suliact of t h i s document: SK 92249 Beta Regulator Envelope l.ift/~ruise Fan Insla llutian S K 92250 L-13081-8 Control Schernatie

Preliminary Aero Lines - Dl3 4/14/75

GDA DRAWINGS * 5 K 20163 PD370-25 A RTA Fun Engine installation I 5 K 20148 RTA Engine-Fan Inferfuco Definition PD370-25A RTA Genera 1 Arrangement SK 20219 PD370-25E RTA General Arrange~nen t SK 20249 FD370-25E RTA Fan Engine [nstal(ation 5 K 20276 The fo~ruwing documents provide definitior~ of the subiact interfaces:

DOCUMENTS -

Statement of Work for NASA Contracts NAS3-19414, NAS3-20033 and NAS3-20034 w i t h Hamilton Standard arid DDA respectively, A coordination memo system exists between HS and DDA w h i c l ~ w i l l be used to define interfaces for this proyra,z. as fhe Fon and engine compcnent designs progress. Data such as rotor speeds, pressure profile, and Flow rates w i l l be coordinoted using this system, These interface coordination memos will be inciuded i n this Interface defini t iun Document as an addendum.

SPECIFlCATi ONS -

The following specificafior~s apply or may be s c d by reference to define thc subiect interface: MIL-E-5007D - General engine requiremenfs.

AS3694, 3 1 May '1973, 'Transmission Systems, VTOL-ST01 Genera!

Requirements for. " A DDA engine speciFication w i l l be issued to mver tk? selected lift/ cruise turbofan engine which w i l l covcr bolt1 tic Hamil -on Stondord and DDA components a5 a u n i t . This specific&ian w i l l be issued after the engine design characteristics are estdbi ished.

IV. --. INTERFACE --- DEFINITIONS (REVISION B )

The following toblc defines the rcsponsi blo con tractor for thc vorious components of the lift/cruisc turbofnn engine and i n turn Ihc interfaces between rnating ~ o m i l t o n Standard and DDA cornpononts: RESPONSIBLE REFERENCE CONTRACTOR DRAWING

-

1.0 MECHANICAL INTERFACE

1 . 1 Fan - Engine Insful Intion

1 . 1 , l Fan Instal lation Drawing HS S K92250 1 . 1 . 2 Engine-Fan Interface Drawing DDA SK20148 1.1.3 Fan Engine lnsta liafion Crowing D DA S K20163 & S ~20276 1.1.4 Fan Engine ~ e n e r a f Arr.angemen t D DA SK 20219 & Drawings 5 K 20249 1.2 Fan-Engine External E n v e ! ~ p e 1.2.1 Fan Ex tcrna l En vc l ape HS SK 92250 1.2.2 Engine Extcrno 1 En vclope D DA SK 20163 1.2.3 Fan-Engine Envelopc DDA 5i( 201 63 1.3 Fan Drive 1.3.1 Fan Drive Shaft Flange D DA SK 20148 1.3.2 Fan Wheel Rear Flange H S S K 92250 1.3.3 Fan Drive Shaft 0 DA S K20148 1.3.4 Fan Drive S l l c l f t Bearings and Support DDA 5E20148 Actvalor 1.4 \ 1 -4.1 Acfuutor Envelope HS S KT2250 1.4.2 Transfer Bearing E n v e I ~ ~ r 3 klS 5 K92250 1 . 4 . 3 inner LVD'i' Envelope HS 5 K92250 1.4.4 Beta Regulo tor Envelope HS S ~ 9 2 2 4 9 1.5 fan Purarne tors 1.5.1 Fan Dcsigr~ Spced 1 . 5 . 2 Fan Blade Tip C1eara:ice 1 . 5 . 3 Fan Speed Pickup 1.6 L/C Rotor Assembly HS S Kg2250 1.6.1 L/C Rotor Component Weight and C(> H5 NA 1.6.2 L/C Rotor Carnpcnent Polar Mornen t H5 NA 1.6.3 L/C Power Recjuiremcnts HS NA

1.7 L/C Gearbox Assembly DDA S K202 19 & S ~ 2 0 2 4 9

i . 7, I

Reduction & Ekvel Gears & Cross S K202 19 & 5 ~ 2 0 2 4 9 D DA S ha ft.

1.8 5 tafionory Cornponenfs 1.8.1 Fan r h c t S to tor Dofini tion D DA S K20118 1.8.2 Engine Inlet Sfotor Definition D DA S K20 148 1.8.3 Fan-Engine Transition Definition DDA SK20148 1.8.4 prirnarv. jecondury Flow Cjpli tter D DA S K20 148 RE5 PONSI BLE REFERENCE CONTMCTOR DRAWING 1.0 MECHANICAL INTERFACE (Continued) 1.9 Forward Frame 1.9. 1 Forward Frame Materials DDA 1.9.2 Forword Fromc Temperatures DDA 1.9.3 Fan Blade Tip Seal Materia! DDA/HS 1.10 L/C Modules ', 10.1 L/c Fan Module Definition H5 i. 10.2 L/C Fan Turboshaft Engine Module Def, DDA 2.0 AERODYNAMIC INTERFACE 2. 1 Component Dcsign Responsibility 2.2 Fan Stage Maps 2 . 3 Engine Inlet Vane Aero Paramefers 2 . 4 Fan Duct Stator Aero Parametor: 2.5 Primary-Secondary Flow Splitter 3.0 ELECTRICAL INTERFACE Pitch Control Schematic Electrical Connection Definition Redundance Requirements wiring Definition Wiring Diagram HS Amperage in Wires MS Voltage in Wires HS L/c Fan Control Modes HS/D DA

L/c Fan lnstrvmentation Requirements

H S

L/c Fan Turboshaft Engine Control System

HS/D DA/AC Fan Speed Pickup DDA 4.0 HYDRAULIC INTERFACE Hydraulic Connections HS 1' ype o f 0 i l DDA Oil System Type & Size of Oil Supply Lines DDA Oi ! Pressures HS O i l low Rates HS L/C Fan Oil Filter Requirements HS Redundan ce Requirements HS Fan Rotor iubr ica tion Requirements HS Pump Drive Locations D DP.

Leakage A1 lowables HS T BD - To be determined as the pragram progresses,

-

N o t applicabte, this n o b t i o n applies in this table t o the farm of N A tronsmittingdata, Thc majority of the data sgnoted w i l l be supplied in tlic form of interface coordination memo which w i l l become a port of this interface document.

ADDENDUM A LIFT FAN INTERFACE DEFINITION ADDENDUM A LIFT FAN INTERFACE DEF1N ITION The l i f t fan interface whiclr has been coordinated with the airframe study contractors, Boeing and McDonnell, i s pro.ided by the installation drawings noted below.

These drawings w i l l be updated dwing the fan detail design to dcFine a l l mechanical inter- faces.

Airframe Confracfor Dro-..v.ving No.

-

Boeing McDonnel l The beta regulator envelope as defined by drawing SK 92249 i s commm to both airfravc contractors and DDA far the lift/r,ruise fan.

Addiiional data pertaining to the lift hn interface which w i l l be established during the )hn detail design i s as follows: 1.0 Fan Onera tiona 1 Parameters 1 , l Fan Design Horsepower 1.2 Fan D ~ s i g n Speed \ 2.0 Fan Characteristics 2.1 Weight 2 . 2 Polor Moment of Inertio 2.3 Center of Gravity 2.4 Vibration Limits 2.5 Gear Ratio 3.0 Aerodvnaniic Cliorocteristics 4.0 Electric01 Interface 4.1 Wiring Diagram 4;2 Voltage Requirements 4.3 Power Requirements 4.4 lnstrurncntafion Requirements 5.0 Hydraulic Interface 5,1 Pressure Requirements 5.2 Flow Requircrnents 5 . 3 ~ i l t r a t i o n Requirements 5.4 Pitch Control Schematic 5.5 Leakage A I Iowablcs 5 . 6 Heat Load 5.7 Type of Fluid

APPENDIX 0 . . - LIST OF SYMBOLS

- -

SYMBOL - DEFINITION - UNITS

A Centrifugal MydrsuIic Dam A 1 I Engines Operating AEO 7 7 rn (in ) A~ ~ i f f / ~ r u i s e NoxrIc Area A 1 First Harmonic Componenf Third Harmonic Component *3 Fourth I-lormonic Component *4 I3 I-lelicai Splined Pinion Drive Shaft BPR Bypss Ratio C Outer H e I i c a I SI?lined Cauplir~g CDP Compressor Discharge Pressurc kPo (psio) CE Center Engine Consumable Elecfrode Vacuum Me1 t

: EVM

rfG Center of G r a v i t y kPa (psia) CIP Compressor In lee Pressure Compressor ItiIe t Temperature C IT C L Condition Lever Compress:., Variab te Geomc try CVG 1.1 A Ft H e l i c a l Splined Coupling LIST OF SYMBOLS (Cont'd) SYMBOL DEFINITION UNITS

-

D DA Detroit Diesel A I1 ison Centrifugal H y d r a u l i ~ Dam E Elccfrohydraulic Servova Ive EHV k l l e v i l l e Springs F Fan Inlet Pressure kPa (psio) .Fly

Fan Inlet Temperature " K("R)

FIT

-

Thrust Control M a r g i n FNCM FNP Primary N o z z l e Thrusf N (Ib) Secondary N o z z l e Thrust FNS N Ilb) System Total N e t Thrust FNT N (Ib)

-

Foreign Object Damage FOD Fan Pressure Ratio FPR Ball Lock G Accelerotiorl of G r a v i t y G ear box High Pressure Hamilton jfnndord tn termedia te Power L e v e l 'TR MAX '- 'TR MIN Radial Distortion Parameter,

'T A V G - 'S A V G

LIST OF SYMDOLS (Cotlt'd) SYMBOL UNITS -8

'TR MAX - ' T R MIN

-

Engine Radial Distortion Parameter, KRE

' T AVG - '5 A V G

P -

TO MAX 'TQ M I 14

CircumFercnfial Distartion Parameter, - -

'T A V G - 'S A V G

P ~ ~ ~ ~ M P ~ e ~ ~ ~ -

Engine C ircumferentia I Distortion Parameter, ' T A V G '. ' S AVG

pi^^^^- 'TQMIN -

Fan Hub Circumferential Dilltortion Parameter, 'T AVG " AVG LC F Low Cycle Fa tigve ~ift/cruise L/C L/C F tift/cruise Fan Lift Fan L F LVDT Linear Variable Differential Transformer

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A i r c r a f t Mach Number MN~,k Fun Rotor Speed K P M N~ Gasifier Rotor Speed R P M

No

Gasifier Set Speed N~ CIE l One Engine Inoperative Power Lever Angle Degrees P 1 A Primary Nozzle Pressure kPo (psia) P P

Power Rating -

Secondary N o z z l e Pressure kPa fps:o) LIST OF SYMBOLS (Cont'd)

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DEFINITION UNITS SYMBOL

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kPa (psia) Arithmetic overage of six equally spacod w a l l static pressure values measured in plane of the total pressures kPo (psio) In let total pressure a r i thmetica I l y averaged aver tho compressor in let annulus Arithmetica I l y averaged total pressure contained in an area kPa (psia) bounded by the compressor inlet inner diameter and an outer diarne ter dcscri bing 60 percent of the compressor in l e i annulus orea Arithmetically averaged tata l pressure contoined in an area bounded kPa (psia) by the compressor inlet outor diameter and an inner diameter describing 40 percent of the compressor inlet annulus area ~ r i t h m e t i c u I ly averaged tota l pressure in any contiguous 240 degree kPa (psia) sector of high-pressure measured o t the compressor inlet annulus Arithmetically averaged toto I pressure in any contiguous 120 degree kPa (psia) sector of low-pressure measured a t the cornpleessor inlet annulus

Reduction Gear -

R G

Research and Technology Aircraft -

RTA Tota I Specific Fuel Consumption m g / ~ - 5 (1 b/hr/l b) S FCT Sea Level Static Conditions + S LS

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TF: Turbofan Engine

-

T i Titanium

Maximum Nacelle Thrust N (Ibs)

Primary Nozzle Temperature " K !OR)

Secondary Nozzle Tempera ture " K("R)

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Turboshaft Engine 9 9 SYMBOL I>EF lNITlf;N UNITS

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Wind r ~ ~ n t r c l Vtilucify rn/s Il;riotsl %I Vcrtir:t~l l'oktuft find 1 a t ~ d i n f j V'TQL ~erticot/Short Ts.?.~?ot.'f urld landing Core Engine Cornpr.cssor Airflow Fuel F l o ~ Accelcrotion Fuel F!ow Deceleration Fuel Flow Governs,. Fuel Flow Maximum Fuel Flow Minimum Fuel F l o ~ Inlet Corrected Airflow Primary Nozzle Flow Secondary Nozzle Flu# Total Flovl In Inlet Flow XT701-AD.-7013 Engine One Engine and Ot~c Fan Operating Two Engines and Three Fans Operating Three Engines a n d Three Fans Operating N a c e l l e Angle o f Attnck Fan Pitch Anglc