NASA Technical Memorandum 100929

NASA/Industry Advanced TurbopropTechnology Program

I_AS_-T_- I£C_9) NAS A/I NLLS_I_/ _EVANC££I[_EC_bC£ 21_(hI_(ICG_ £FCG_A_ II_A2A) 26 F

CSC£ 21_

G3,/07

Nd6-24t4 1

Joseph A. Ziemianski and John B. Whitlow, Jr.Lewis Research Center

Cleveland, Ohio

Prepared for the

16th Congress of the International Council ofAeronautical Sciences--ICAS '88

Jerusalem, Israel, August 28--September 2, 1988

https://ntrs.nasa.gov/search.jsp?R=19880015257 2020-03-20T07:12:47+00:00Z

NASA/INDUSTRY ADVANCED TURBOPROP TECHNOLOGY PROGRAM

Joseph A, Ziemianskl and John B. Whitlow, Jr.Natlonal Aeronautics and Space Admlnlstratlon

Lewis Research CenterCleveland, Ohlo 44135

Abstract

Experimental and analytlcal effort shows thatuse of advanced turboprop (propfan) propulsionInstead of conventlonal turbofans In the oldernarrow-body alrllne fleet could reduce fuel con-sumption for thls type of alrcraft by up to50 percent. The NASA Advanced Turboprop (ATP)program was formulated to address the key technol-ogles required for these new thin, swept-bladepropeller concepts. A NASA, Industry, and univer-sity team was assembled to develop and valldateapplicable new design codes and prove by groundand flight test the vlablllty of these new propel-ler concepts. Thls paper presents some of the

' history of the ATP project, an overvlew of someof the Issues and summarizes the technology deve-loped to make advanced propellers viable In thehlgh-subsonIc cruise speed application. The ATPprogram was awarded the prestlglous Robert O.Colller Trophy for the greatest achievement Inaeronautics and astronautlcs In America In 1987.

I. Introductlon

Until the advent of the ATP program in 1978,

propeller technology development had stopped Inthe mld-1950's. Thls 20-year gap in developmentoccurred because we did not know how to build prop

blades wlth the comblned structural and aero-

dynamic chaFacterlstlcs to operate reliably andeffIclently at the hlgh subsonic cruise speedsobtained by the turbojets then being IntroducedInto commerclal servlce.

Although propellers were more fuel efficlent

than turbojets and turbofans, propulsion develop-ment was largely concentrated on Improvements tothe turbojet durlng thls era of cheap fuel.

Perspectives changed beginning In 1973, as aresult of the Mlddle East oll embargo. Fuel costsescalated and by 1980 had gone from about one-

quarter to more than half of direct operatingcosts, as Illustrated In Fig. 1.

In January 1975, the U.S. Senate Commlttee onAeronautical and Space Sclence requested that NASAdevelop a program to address the fuel crlsls. (1,2)

In response, NASA formed an Inter-agency task forcethat consldered many potentlal fuel-savlng con-cepts. They proposed the Alrcraft Energy Effl-

clency (ACEE) program, which Included threepropulslon projects managed by NASA Lewis. Oneof these, strongly advocated by NASA Lewls andHamllton Standard Dlvlslon of United Technologles,was to develop advanced turboprops whlch could

overcome the hlgh-speed compresslblllty losses ofconventlona} propeller designs.

NASA conducted several system studies whichshowed that advanced propellers could have propul-sive efflclencles about 1.3 tlmes as hlgh as those

of equivalent turbofans at cruise speeds of Mach 0.8.These results are illustrated In Flg. 2. Also

shown Is an artlst's rendltlon of the new advanced

turboprop, or "propfan" concept and, for compari-son, the old four-bladed prop typical of thoseused on the Lockheed Electra. Although the old

turboprops were fuel efficient up to airspeeds

of slightly over Mach 0.6, they experience a rapidincrease In compressibility losses beyond thesespeeds due to thelr thick, unswept, large-diameterblades. Their propulsive efficiency is much higher

than that of hlgh bypass turbofans at speeds upto Mach 0.6+ because In generating thrust a prop

Imparts only a small Increase In axial velocity to

a large mass flow of alr, thereby reduclng Kineticenergy losses In the discharge flow. The advancedturboprop uses very thin, highly swept blades toreduce both compressibillty losses and propeller

noise during hl_h-speed cruise. High dlsk powerloadlngs (SHP/D Z) at least double that of theLockheed Electra are required for high-speed cruiseand are achieved by Increasing the number of

blades and lengthening blade chord. Counterrotat-

Ing blade designs can be used to provide stillhigher disk power loadlngs and eliminate some ofthe exit "swirl" losses assoclated with single-

rotation propellers. Installed propulsive effl-clencles roughly equivalent to those achieved with

the old Electra technology can be extended to theMath O.B regime with advanced propellers. Theadvanced turboprop produces fuel savings conserva-tively estimated at about 30 percent due to the

Improved propulslve efficiency of the propfan and50 percent better overall with core engineimprovements Included. (2) To illustrate theimpact of such a savings on today's U.S. alrllne

fleet, if advanced turboprops were substitutedfor the turbofans used in only the narrow-body727's, 737's, DC9's, and MD80's, about 2.5 billiongallons of fuel would be saved per year. (3)

II. ATP Programmatic ObJectlves and Plans

NASA formally began the Advanced Turboprop(ATP) project In 1978 with the overall objective

of valldatlng key technologies required for bothslngle- and counterrotatlng propfans in Mach 0.65to 0.85 appllcatlons. Project goals were to verlfy

projected propfan performance and fuel savingsbeneflts; to verify the structural integrity of

these radically different blade deslgns underactual operatlng conditions; to establlsh passengercomfort levels (i.e., cabin noise and vibratlon)approaching those in modern turbofan-powered alr-liners; and to verify that propfan-powered alr-craft could meet the airport and community noise

standards speclfled by U.S. Federal Alr Regula-tlons (FAR-36). The project plan projected system

technology readiness by the late 1980's.

The NASA ATP project was organized in such a

way that technlcal Issues were first resolvedthrough wlnd tunnel testing of sma11-scale modelsbefore more costly large-scale ground and flight

testlng. The early years of ATP (prior to 1980)

providedtheenabllngtechnologyvla sma11-scaletestinganddeslgncodedevelopmentto establlshthe feaslb111tyof thepropfan. Afundamentaldatabaseof design,analysis,andtesting tech-niqueswasdeveloped.A large-scaletechnology}ntegratlonphasethendrewfromthls knowledgeto deslgn,fabricate,andgroundtest large-scalepropfansystems.Thelarge-scaleeffort wasneededto e11mlnateuncertalntlesconcernlngthescale-upof structural andacousticdataobtalnedwlthwindtunnelmodels.(1) F11ghtresearchtestlngof large-scalepropfanpropulslonsystemswasInitiated in 1986to verlfy bladestructuralIntegrity, to determlnecabincomfortandgroundnoiselevels, to provldescalingcomparlsonswlthmodeltunneldata,andto validatecomputeranal-yses. Asthesetests werecompleted,anexten-slveanalyslseffort wasImplementedto assessthedataacquired.

Fromthebeglnnlngof the ATP project, a sys-tems approach which consldered the entire aircraft

was used In deslgnlng the propulslon system, asshown in Fig. 3. This Included elements such asthe propeller and the nacelle, the drive system,Installatlon aerodynamics, and the aircraft Inte-

rior and communlty environments and the effect ofthese elements on meeting the goals of reduced

fuel consumptlon, low operating costs, and pas-senger acceptance. Thls approach followed the1oglc path illustrated In Fig. 4. The sequencestarted with analyses and systems studles and

proceeded to design code development based onscale-model wlnd tunnel tests or component tests.

Finally, large-scale systems were designed,built, and tested both on the ground and infllght as proof of the concept. This approachwas used In the three propfan conflguratlonalareas shown: slngle rotation, gearless counter-

rotation, and geared counterrotatlon.

The technical expertise of all three NASA

aeronautical research centers (Lewis, Langley,and Ames), more than 40 contracts distrlbuted

over the maJorlty of the U.S. aircraft Industry,and over 15 universlty grants were required to

complete the project. Major contracts wereawarded to General Electric on the Unducted Fan

(UDF), Hamllton Standard on the Large-ScaleAdvanced Propfan (LAP), and Lockheed-Georgla onthe Propfan Test Assessment (PTA). In addition

to NASA-sponsored research, a significant inde-pendent Industrlal research and developmenteffort was applled to develop these new concepts.Beglnnlng in August 1986, the advanced turboproppropulsion concept was proven by three flight

programs uslng large-scale hardware (Fig. 5).The NASA-General Electrlc-Boelng fllght test andthe General Electric-McDonnell Douglas flight

test used the Unducted Fan as a proof-of-conceptdemonstrator for the gearless counterrotatlngconcept. The NASA/LocKheed-Georgla Propfan Test

Assessment verlfled the structural Integrlty andacoustic characteristics of the slngle-rotatingLAP propfan bullt by Hamilton Standard. On thebasis of the success of these tests and previous

scale-model work, Pratt & Whttney-Alllson builta geared propulslon system with Hamilton Standard

counter-rotatlng propellers that they plan toFly on the MD-80 in late 1988.

The three top pictures of Fig. 6 illustrate

the post-fllght test NASA generic propellerresearch program of analysis and scale model wlnd

tunnel tests leading to design valldation andverification of aerodynamlc, acoustic, and struc-tural codes. This on-golng research program willmake extensive use of the previously acquired ATP

database. Advanced concepts of a slngle-rotatlonpropfan with stator vane swirl recovery and ahlgh-bypass-ratlo ducted-fan configuration areIllustrated in the two bottom drawings. Hhlle

future ducted props will not have the efficiencyof unducted propfans they may be more sultable

for "packaging" on large alrcraft such as theBoeing 747,

Ill. Slngle-Rotation Systems

The first 2-ft-dlameter propfan single-

rotation model, deslgnated SR-I and Incorporat-ing eight thln, swept blades was developed byNASA and Hamllton Standard i_ 1976.(2)

When tested In a wind tunnel, the SR-I

achieved an efficiency of 77 percent atMach 0.8. The model blades were stable even

when an attempt was made to force flutter.

Encouraged by thls but st111 needing to fullyunderstand the efflclency and noise potential ofthe propfan, several more models were designedand tested.

One was an Improved verslon, the SR-IM, wlth

a modlfled spanwlse twist to better dlstrlbutethe blade loading, which resulted in a l-polntgain In overall efficiency, Another model, the

stralght-bladed SR-2, was designed to provide abasellne for comparlson. Its efficiency wassllghtly less than 76 percent at Mach 0.8. Thesubsequent SR-3 model, whlch incorporated 45° of

sweep for both aerodynamic and acoustic purposes,achleved an efflclency of nearly 79 percent.Some of the early blade models that were wind-tunnel tested are shown in Fig. 7.

The performance gains (i.e., Incrementalgains in propeller efficiency) and noise reduc-tions due to increased blade sweep for the SR-2,

SR-IM and SR-3 configurations are summarized inthe plot of measured data In Fig. 8.(2) All

three blades were very thin but varied In amountof tip sweep from 0° to 45°. All three of these

configurations are compared at their design polnt

condition _f Math 0.8, disk power loading of37.5 SHP/D L, and tip speed of 800 ft/sec. It Isclear that both performance and acoustics bene-fit as blade sweep is Increased. These model

tests eventually led to the wind tunnel test ofthe SR-7A model, which Is an aeroelasticallyscaled 2-ft model of the 9-ft proDfan flighttested later In the PTA program. (4)

Net efflclencles of the propeller models areshown as a function of Math number In Fig. 9. (5)

At Math 0.80 the SR-7A propfan has the hlghestmeasured propeller efficiency - 79.3 percent.

The performance of the SR-2 propeller Is lowerthan that of the others because of its unswept

blade design. The number of blades and amount of

tlp sweep are tabulated below for each of thesemodels.

Design Numberofblades

SR-7A 8SR-6 I0SR-3 8SR-IM 8SR-2 8

Sweepangle,deg

414045300

Near-fleldacousticresultswererecentlyobtalnedwith theSR-TAmodelat hlgh-speedcruiseconditionsin theNASALewis8x6ft tunnel,asshownin Flg. IO. Peakfundamentaltonelevelsareplotted againsthelical tlp Machnumber(I.e.,thebladerelative Mathnumber)for threeloadinglevels.(6) Thedatashowthat fundamentaltonelevelsmaypeak,level off, OFdecreasebeyondahellca] tlp speedof Machl.l, dependlngonoadlng.

Althoughtheeffect of increaslngbladesweeps positive in termsof Improvementsto propulsive

efficiencyandacousticsat highflight speeds,the structural andaeroelastlcdeslgnbecomesmoredifficult. Onestructuralconcernrelatestosteadystate stresslevelsdueto centrlfugalandsteadyaerodynamicloads. Anotherrelatesto anaeroelastlcinstablllty phenomenoncalled flutterwhich can occur In response to forced excitations

caused by unsteady, unsymmetrical airflows pro-duced by gusts, upwash from the wing, and airframe-induced flow fleld distortions. The presentlylll-deflned boundary for hlgh-speed classical

flutter will occur at increasingly reduced flightspeeds as blade tlp sweep is increased, all otherthings (e.g., materials, constructlon) remalnlng

equal. Avoidance of this flutter boundary is onereason why the sweep of the SR-7 was llmited to41 ° at the tlp.

An experlmental and analytlcal research pro-gram Is belng conducted to better understand theflutter and forced response characteristics ofadvanced hlgh-speed propellers. A comparlson ofmeasured and calculated flutter boundarles for a

propfan model deslgned to flutter is shown inFlg. 11. (7,8) The theoretlcal results, from theNASA Lewls-developed ASTROP3 analysls, _nclude the

effects of centrlfugal loads and steady-state,three-dlmenslonal alr loads. The analysis does

reasonably well in predicting the flutter speedsand slopes of the boundaries. However, the dif-ference between the calculated and measured flut-ter Mach numbers Is greater for four blades than

for eight blades. This Implles that the theoryIs overcorrecting for the decrease In the aerody-namic cascade effect wlth four blades.

EuleF code so]utlons have recently been deve-

loped to descrlbe the unsteady, three-dlmenslonalflow fleld generated with advanced propellerdeslgns. (9) A graphlca] dlsplay of the blade pres-

sure contours from this analysls is shown InFlg. 12 for an SR-3 propfan in regions where theflow is supersonic when the axis of rotation Is

tilted upward 4° from the Math 0.8 free-streamflow. The downward moving blades (on the right_n the flgure) experience the hlghest ]oadlngs,

whereas the upward movlng blades experience someunloading due to the alternating alignment of the

upward component of the free-stream veloclty with

the upward and downward blade rotational velocltyvectors. Pressure contours for the blades at the

top and bottom of the rotation are relativelyunaffected by the tilt of the rotatlona] axlsbecause at these positions the rotatlonal velocity

component Is perpendicular to the upward free-stream component.

Cabin noise and vlbratlon levels with past

turboprops has been less favorable than wlth turbo-fans. To achieve a cabin environment with propfans

that Is comparable to current turbofan transports,a reduction of 25 to 30 dB beyond the capacity of

a bare-wall untreated cabln is likely to be

Fequlred for a wing-mount installation. NASA

Langley Research Center has been involved in sev-eral ATP noise reduction activities, includlng the

evaluation of advanced cabin sldewa]] concepts. A

Langley/Lockheed-Callfornla effort has led to thedevelopment of an advanced cabin wall acoustictreatment utlllzlng Helmholtz resonators tuned tothe fundamental blade passing frequency. Thls con-

cept was flight tested in the PTA program and is dls-cussed later In con_ectlon with those fllght tests.

Under NASA sponsorship, Hamilton Standard

initiated the design of a Large-Scale Advanced

Propeller (LAP) In 1982. The resultlng 9-ftdlameter SR-7 propfan deslgn Incorporates the

spar-shell type of blade construction Illustratedin Fig, 13. (1) All new Hamilton Standard straight-bladed commuter aircraft propellers use a similar

spar-shell type of construction which has provento be very safe, reilable, and lightweight. TheFOD problems inherent in earlier solid aluminumblades are avoided by protecting the single Ioad-

bearing spar with an aerodynamlcally-shaped flber-

glass shell. This construction technique, however,was unproven for the thin, swept, LAP blade deslgnwhich is subjected to complex nonlinear deflec-tions under load. The possibility of hlgh-speed

classical flutter and the need to verify the designcodes that were used reinforced the need for

large-scale fabrication and fllght testlng.

The 9-ft dlameter size for the LAP rotor

assembly was selected as the minimum size thatwould permlt a reallstlcally-scaled blade crosssectlon wlth the mlnlmum allowable shell gauge

thlckness. Exlsting drive system capabIllty waslimited to about 3000 SHP at the Mach 0.8/35 000 ft

design polnt and also dlctated a diameter of about9 ft If disk power load_ngs (SHP/D 2) in the desireddeslred 30 to 40 SHP/ft L range were to be obtained.

A LAP static rotor test was completed In late

1985 at a Wrlght-Patterson Air Force Base facll-

Ity, as shown In Flg 14, using a facility elec-trlc drive motor. (IUi Forty-slx strain gages,

some of which can be seen In thls photo, wereinstalled on the LAP durlng manufacture. Of this

total, 30 strain gages were connected through sllp

rlngs to a data system and continuously recordedduring propfan operation while the remainder were

spares which could be used in the event of a mal-functlon of one of the primary gages. This %nstru-

mentatlon and data system were retained throughoutthe LAP and follow-on PTA testing. In early 1986,the LAP was Installed In France's Modane wlnd tun-

nel to verify blade structural integrity at speeds

up to Mach 0.83. A second Modane tunnel entryoccurred In early 1987 to acquire blade steady and

unsteady pressure data for verifying and improvingaerodynamlc predlctlon codes. Figure 15 depicts

the second LAP Modane entry, blade pressure Instru-mentatlon locatlon schematics for two blades spe-

cially instrumented for this particular test, andalso one of the two completed propfans delivered

to the PTA project.

The two-bladed verslon of the elght-blade

propfan shown In F_g. 15 was used In some of theModane testing because of the llmlted fac111typower available to drive the propeller. In th_s

way the propeller could be operated at a reason-able power 1oadlng per blade. The large size ofthis propeller allowed much more detailed blade

pressure measurements than could be obtalned onthe 2-ft diameter models tested prevlously.

Prior to f11ght test, several scale-modeltests of the PTA alrplane were conducted In NASA

wlnd tunnels In the Interest of fllght safety andto obtain data for valldatlng aerodynamic predIc-tlon codes. Both a 11g-scale airplane aeroelas-

tic model 11 and a !_-scale aerodynamlclstab111tyand control model _m_j were built and tested. The

results From both models showed excellent agree-

ment with analytlcal predlctlons. A rake surveyof the flowfleld at the propfan plane was alsoconducted wlth a modified version of the PTA aero-

dynamic model to verlfy the flowfleld predIctloncode. (13) The predicted ?lowfleld at the varlousf11ght test conditions was used as a correlatlon

parameter for measured propfan stress.

A ground static test of the entire PTA pro-

pulsion system - Including the straln-gage-Instrumented propfan, englne/gearbox, and forwardnacelle - was conducted In the sprlnq of 1986 atan outdoor thrust stand (Fig. 16).(14) Over 50 hr

of extenslve testing was accomplished, with essen-tially flawless system operation and with bladestresses at a level somewhat below those seen at

the previous static rotor test. In neither ofthese static tests was there any evidence of blade

flutter and stresses were low except at very hlghblade angles where buffeting characteristic ofseparated flow was sometimes Indlcated.

After G-II alrcraft modlflcatlons to Install

the propfan propulsion system on the left wing and

the subsequent ground checkout testlog_)PTA_I flighttesting began In the spring of 1987. Somephotos of the PTA fllght testing are shown InFig. 17. One unique feature of the PTA design wasthe varlable tilt nacelle whlch allowed the for-

ward portlon of the nacelle to be tilted up ordown so that blade stresses could be assessed as a

functlon of Inflow angle. The PTA flight test

program was performed to verlfy LAP structuralIntegrlty and characterize LAP acoustics both out-side and Inside the cabin as well as on the

ground. Data were obtained over a Flight envelopeextendlng from Just above low-speed stall toMach 0.89 and at altltudes from 8OO to 40 OOO ft.

A total of up to 613 acoustlc, g-loadlng,pressure, strain, temperature, and miscellaneous

operatlng parameters were recorded onboard theaircraft at each of almost 900 completed test runs

durlng the PTA research fllght tests. The PTAflight test program Involved more than 133 hr of

flight tlme over a total of 73 flights.

Inltlal fllght nolse test data agree favor-

ably wlth both scaled-up NASA wlnd tunnel data

and predictions obtalned with an analytical code. (15)

Comparatlve maximum noise data along the fuselageexterlor are presented at axial locatlons fore and

aft of the plane of rotation In Flg. 18. A maxi-mum sound pressure level of 147 dB was measured at

the fundamental blade passlng tone of 225 Hz. Themeasured local noise reduction at an adjacentlocation Inside the bare-wall cabin was 25 dB.

Subsequent to the basic bare-wall cabin fllghttest effort, a lO-ft sectlon of the PTA cabin was

cleared for acquiring data wlth an advanced cabln

acoustic treatment In early IgBB. The treatedenclosure, located fore and aft of the propfanplane of rotation, consisted of tuned Helmholtzresonator wall panels attached to a Framework

mounted to the cabin floor through vibration iso-lators. Preliminary results obtained from the 31

cabin Interior microphones Indlcate that noiselevels 25 to 30 dB below that of the bare-wall

cabin were obtalned. This is the approximate

level required for comparability with existingturbofan-powered airliners.

The PTA flight test effort was concluded in

March 1988. Although some preliminary results areavailable, because of the masslve quantity of datato be analyzed the final results will not be avall-able until October 1988. It is clear, however, that

these fllghts, as intended, verified propfan struc-tural Integrlty. There was no evidence of flutteranywhere In the fllght regime and blade stressing

was In good agreement with predictions. Measuredblade stresses were wlthln limits established byHamilton Standard for Inflnlte llfe. Prellmlnaryacoustic data analysis Indicates that magnltudes

and trends are generally as predicted wlth prop-fan nolse sllghtly lower than predicted. Itappears that advanced cabin acoustic treatmentscan reduce interior noise to acceptable levels andthat FAR36 (stage 3) airport communlty standards

can be met when data are extrapolated to a productdesign.

IV. Gearless CounterrotatIon Systems

CounterrotatIon propeller systems are ofInterest because of their potential to further

enhance propulsive efflclency by reducing or ellm-Inatlng the swirl component of the discharge

velocity. In order to generate propulsive thrust,a propeller must take essentially axial flow andturn it to do work, in much the same way that anairplane wing must turn the flow slightly downwardIn order to generate upward llft. The result with

a single-rotatlon propeller is that the dischargeflow must have a nonax1al rotational component ofperhaps several degrees, depending on disk loadingand tlp speed. A decrease In net thrust (and,

hence, propulsive efficiency) results from thlsnonaxlal, or "swlrl," velocity component slnce thetotal change In momentum is not In the axial dlrec-tlon, as It Ideally should be for the production of

thrust. The swlrl losses for an Isolated single-rotatlon propfan at Math 0.8 design point operat-Ing condltlons are typically equivalent to about

elght points In efficiency. These losses can bereduced or ellmlnated by usln9 counter-rotatlon

as a swirl recovery technique. Wlth counter-rotation, the second, or aft stage, propeller

rotating in the opposite dlrectlon returns theflow to the axial direction as It performs Itswork.

By 1983 General Electric became convincedthat a gearless counterrotattng Unducted Fan (UDF)engine would be a vlable fuel-saving alternativeto the turbofan. Rather than venture into theuncertain area of gearbox design for a 20 000 hp_class engtne, GE chose to ellmlnate the gearbox byusing a counterrotatlng power turbine to dlrectlydrive the props. (2,16) A cutaway of thls conceptls shown in Fig. 19. The gas generator ahead ofthe power turbine tn th_s concept demonstrator isnot mechanically linked to the power turblne,which is drtven solely by hot exhaust gas.

The UDF prop blades were designed for an over-

all disk power loading almost twice as great asthat of the slngle-rotatlon designs. Hub-to-tlpradlus ratio of these blades _s about 75'percent

hlgher than that typical of geared deslgns Inorder to accommodate the large-dlameter power tur-

bine. The large turbine dlameter Is required forpower generation and compensates for the low rpmrestraint imposed by the prop tlp speed IImlt.

Except for a set of Inlet and outlet guide vanes,this 12--stage power turbine Is unlque in that Ithas no stator vanes between the alternating

opposlte-rotatlon blade rows.

The UDF blade structural deslgn chosen by GE

Is somewhat dlfferent from that used by HamiltonStandard for the LAP blades. The LAP blade con-

sisted of a full-length structural spar and aF_berglass she11; the shell of the UDF blade, Incontrast, Is the structural element and the spariS used for attachment. (2) The blades have a

half-span tltanium spar covered wlth an aerody-

namlc shell of epoxy-bonded carbon-flber/flberglass plies, as shown In Fig. 20. The pllesare orlented In such a way as to tune the dlrec-tlonal stlffness for blade shape control,

strength, and aeromechan_cal stablllty. The bladedeslgn also Includes a nickel leadlng-edge sheathand polyurethane film bonded to the outer shell toprovide addltlonal protectlon. Blade design andconstructlon Is baslcally the same for both for-ward and aft rows.

The NASA Lewls counterrotatlon pusher propel-

]eF test rlg shown wlth a UDF blade conflguratlonin the 8- by 6-Ft Wlnd Tunnel In Flg. 21 was oneof three bullt by GE for model blade testlng. (17)

The other two were used at Boelng and GE. Per-formance, flowfleld, and acoustic measurementswere made during thls testlng. The UDF modelblade configuratlons tested at NASA Lewis are

shown In Fig. 22. The designs dlffered in tipsweep, planform shape, alrfoll camber, andIncluded one case with a s_gnlf_cantly shortenedaft rotor. The planform shapes for most forwardand aft rotors were very slmilar. These bladeswere deslgned and bullt by General Electrlc. Netefflclences for the F7-A7 blade configuration are

shown In Flg. 23 as a functlon of crulse speed forthree power 1oadlngs. (5) At Mach O.72 design con-

ditions, efflclency Is strongly affected by 1oad-

Ing, but as Math number Increases, compresslbilltylosses domlnate and efflclencles fall off essen-

tlally Independent of loading.

General Electrlc used NASA design codes for

propeller ply deslgn, flutter analysls, aerody-namic design, and nolse predlctlon. Model rlgdata was also used by GE to modify, Improve, and

verify their own Inhouse codes.

ORIGINAL PAGE IS

OF pOOR QUALIT_

Both steady and unsteady flow predictioncodes have been developed for counterrotation pro-pellers. Flgure 24 shows the three-dimensional

Image of the UDF pressure distrlbution generatedwith a NASA Lewis-developed Euler predIctlon

code. (18) The coupling between rows is done in aclrcumferentially-averaged sense which does notconslder blade-wake interactlon effects. The pres-sure dlstributlon is shown on the nacelle, bladesurfaces, and on a flow cross section downstreamof the aft rotor. An unsteady Euler code solutlonhas also been applled to the FT-A7 UDF configura-tion to obtain the full unsteady three-dimensionalsolutlon for the flow field. _> Pressure contoursat a particular instant in time are shown inF_g. 25 at a plane just downstream of the aftblade row. The low pressure islands shown are theresult of tlp vortex shedding.

The UDF demonstrator engine is shown duringground static testlng at GE's Peeb[es, Ohlo out-door test faclllty In the top photo of Fig. 26.After completlng thls ground testing, the UDFdemonstrator was installed on a Boelng 727 air-plane for flight testlng In August 1986. Theengine was later Installed on an MD-80 In May 1987for a Douglas f11ght test program. These flighttest conflguratlons are shown in the two bottomphotos of Flg. 26. The UDF was substltuted forthe rlght-hand JTSD powerplant on the 727 and forthe left-hand JT8D on the MD-80 alrplane.

Fundamental tone dlrectivities for the F7-A7blade comblnatlon, the proof-of-concept UDF con-

flguratlon, are shown in Fig. 27 for scaled windtunnel model data and fu11-scale fllght data

obtalned by the Instrumented NASA Lewls Leafierin formation flights with the UDF-powered 727. (5)There Is excellent agreement among the model wind-tunnel measurements, fu11-scale f11ght data, and

prediction at most sldeline angles, although thewind tunnel data appear to be somewhat high atforward angles.

V. Geared Counterrotatlon Systems

NASA has also sponsored research leadlng tothe development of a more conventlonal gearedcounterrotating propfan system uslng blade tech-nology which _s basically an extenslon of thatpioneered in the LAP slngle-rotation propfan. As

part of this effort, A11ison Gas Turblne Dlvlsionof General Motors designed, fabricated, and rig-tested a 13 OOO shp-class advanced In-llne differ-entlal planetary counterrotatlon gearbox. (15)

Ease of ma_ntalnabillty, hlgh (over 99 percent)mechanical efficiency, and a high durability were

of paramount Importance In thls gearbox design.A111son used the results of these tests in deve-

1oplng the technlcally slmllar flightwelght gear-box for the Pratt & Whltney-A111son/Douglas578DX/MD-80 flight tes_ program.

The Hamilton Standard CRP-XI propeller model

(Flg. 28) was evaluated in wlnd tunnels for per-formance and flow fleld data, blade stresses, andacoustics.(19, 20) At the Math O.8 deslgn power

1oadlng, a net effIclency improvement of ~6 percentwas obtalned over the analogous SR-7A single-rotatlon model. The counterrotatlon propfan

deslgn used In the MD-80/578DX flight test isbased on the technology of the CRP-XI model andthe earller SR-7.

Flowcharacteristics of these types of coun-terrotatlon propellers have also been predictedwith analytical codes. One such example, shown in

Fig. 29, Is the analytically slmulated flow vlsu-a11zatlon of an off-deslgn vortex shedding phe-nomenon with the CRP-XI nx)del design. If notaccounted for In analytical models, errors In pre-dicted performance and noise wlll occur. Theresults shown are from an Euler code developed atNASA Lewis and ruff to slmulate takeoff with a

high blade Incidence angle. (18)

Levels of the first flve harmonics of slngle-rotatlon and counterrotatlon propeller nolse,based on Hamilton Standard model test data, are

shown in FIq. 30 at three axial locations in thefar field. (20) The single rotatlon tone levels

have been adjusted upward 3 dB to compare theequivalent of two independent propellers wlth theCRP-XI counterrotation conflguration. Single andcounterrotatlon fundamental tones are then roughly

equal, but the counterrotation higher harmonicsare dramatlcally higher at all locations due to

the unsteady aerodynamlc Interactions betweenblade rows. The hlgh fore and aft harmonlc levelsmust be dealt wlth to achieve acceptable counter-

rotatlon community nolse levels.

VI. Future Research

In future work, NASA w111 attempt to achlevesome of the swirl recovery benefit of counterrota-

tlon without the addltlonal complexity and noise,by uslng swirl recovery vanes wlth a single-rotatlon propfan model (Flg. 6, bottom left).

Future NASA research effort will also be

directed toward the ducted propeller, whlch wasdlscussed brlefly in connection wlth Fig. 6.Ducted props are more easily integrated Into thedesign of large long-range aircraft than unducted

props because they are more compact than unductedpropfans. Underwlng %nstallatlons of propfans onlarge heavy aircraft are unllkely because thelarge tlp diameter requirements will cause groundclearance problems. There are several technicalIssues which must be addressed with regard toducted props. (5) At crulse, the drag of thelarge-diameter thln cowl must be Kept low whilemaintaining acceptable near-field noise levels.Tradeoffs between propeller and fan aerodynamicdeslgn methods are required to arrive at the opti-mum combination of ducted prop design parameters.At low-speed condltlons, far-field noise tn thecommunity; tlp flow separation and blade stresseswlth a short, thtn cowl at high angles of attack;and reverse thrust operation are technlcal lssuesrequiring further Investigation.

VII. Concludln 9 Remarks

Propfan technology In llttle more than a dec-

ade has evolved from a fuel savlng idea, throughthe test of small-scale models, to the current

fllght test of large-scale complete propulslon

systems. In 1987 the advanced turboprop propul-slon concept was proven by three fllght programs

uslng large-scale hardware. The NASA/GeneralElectrIc/Boelng fllght test and the General Elec-trlc/McDonnell Douglas flight test used theUnducted Fan demonstrator to prove the gearless

counter-rotatlng concept. The NASA/Lockheed-Georgia Propfan Test Assessment fllght test used

a propfan built by Hamilton Standard to prove thegeared slngle-rotation concept and to comparelarge-scale fllght data with an extensive windtunnel model data base. On the basls of the suc-

cess of these tests and previous scale-model work,Pratt & Whitney - Allison built a geared counter-rotatlng propulslon system that they plan to fly

thls year on the MD-80 aircraft. It is clear Fromthese efforts that much of the predicted fuel sav-

Ings potential can be realized with designs whichmalntaln their structural Integrlty In flight.Prellmlnary Indicatlons are that propfan noise

trends are also generally as predlcted and thatacoustic treatments can be deslgned to satlsfacto-

rlly attenuate cabin noise.

Because of the fuel savings potential for

thls concept, whlch can be implemented without anysacrlflce In the speed and comfort we have grown

to expect, it Is anticipated that this technologywill lead to a whole new generation of propfan-powered alrcraft - both clvll and mllltary.

Although more work is yet to be done in ATP

data acqulsltlon and analysls, the rapidly expand-Ing database already In existence w111 permittechnIcally sound marketing decisions to be madeby industry in the deployment of prlvate invest-ment capltal. Early candldate production aircraft

are now on the drawing boards. The Lewls ResearchCenter and the entire NASA/Industry Advanced Tur-boprop Team were cited In the Collier Trophy award

(Fig. 31) presented in May 1988, by the NationalAeronautic Assoclation for this work which theyconslder as the greatest achievement in aeronau-tics or astronautics In America demonstrated by

actual use in the previous year.

5.

6,

Vlll. References

Whltlow, J.B., Jr., and Slevers, G.K., "Fuel

Savlngs Potential of the NASA AdvancedTurboprop Program," NASA TM-B3736, 1984.

Hager, R.D., and Vrabel, D., "AdvancedTurboprop Project," NASA SP-495, 1988.

Whitlow, J.B., Jr., and Slevers, G.K., "NASA

Advanced Turboprop Research and ConceptValldatlon Program," NASA TM-lO0891, 1988.

Stefko, G.L., Rose, G.E., and Podboy, G.G.,"Wind Tunnel Performance Results of an

Aeroelastlcally Scaled 219 Model of the PTA

Flight Test Prop-Fan," AIAA Paper 87-1893,June 1987. (NASA TM-89917).

Groeneweg, J.F., and Bober, L.J., "AdvancedPropeller Research," Aeropropulslon '87,

Sesslon 5, Subsonic Propulsion Technology,NASA CP-IOOO3-SESS-5, 1987, pp. 5-123 to5-152.

Dlttmar, J.H., and Stang, D.B., "Cruise Nolse

of the 2/9 Scale Model of the Large-ScaleAdvanced Propfan (LAP) Propeller, SR-TA,"AIAA Paper B7-2717, Oct. 1987. (NASATM-I00175).

7. Kaza,K.R.V.,Mehmed,0., Narayanan,G.V.,andMurthy,D.V., "AnalyticalFlutterInvestigatlonof a CompositePropfanModel,"28thStructures,StructuralDynamicsandMateFlalsConference,Part2A,AIAA,NewYork,1987,pp.84-97. (NASATM-88944).

8. Ernst,M.A.,andK1raly,L.J., "DetermlnlngStructuralPerformance,"Aeropropu]sion'87,Session2, AeropropuIslonStructuresResearchNASACP-IOOO3-SESS-2,1987,pp.2-11to 2-24.

9. Nh_tfleld,D.L., Swafford,T.N.,Janus,J.M.,Mulac,R.A.,andBelk,D.M.,"Three-DimensionalUnsteadyEulerSolutlonsfor PropfansandCounter-RotatingPropfansIn TransonicFlow,"AIAAPaper87-1197,June1987.

10.DeGeorge,C.L., "Large-ScaleAdvancedProp-Fan(LAP),"NASACR-182112,1988.

ll. Jenness, C.M.J., "Propfan Test AssessmentTestbed Alrcraft Flutter Model Test Report,"LG85ER0187, LocKheed-Georgla Co.,

Marletta, GA, June 1986, NASA CR-179458.

12. A1Jabrl, A.S., and Little, B.H., Jr., "Hlgh

Speed Wind Tunnel Tests of the PTA Aircraft,"SAE Paper 861744, Oct. 1986.

13. A1Jabrl, A.S., "Measurement and Predlctlon ofPropeller Flow Field on the PTA Alrcraft atSpeeds of up to Mach 0.85," AIAA Paper88-0667, Jan. 1988.

14. Withers, C.C., Bartel, H.W., Turnberg, i.E.,and Graber, E.J., "Static Tests of the

Propulsion System - Propfan Test Assessment

Program," AIAA Paper 87-1728, June 1987.

15. Graber, E.J., "Overvlew of NASA PTA PropfanFlight Test Program," Aeropropulsion '87,Session 5, Subsonic Propulsion Technology,

NASA CP-IOOO3-SESS-5, 1987, pp. 5-97 to 5-122.

16. Adamson, A., and Stuart, A.R., "Propulsionfor Advanced Commercial Transports," AIAA

Paper 85-3061, Oct. 1985.

17. Delaney, B.R., Balan, C., Nest, H., Humenik,F.M., and Craig, G., "A Model Propulsion

Slmulator for Evaluating Counter RotatingBlade Characteristics," SAE Paper 861715,Oct. 1986.

18. Celestlna, M.L., Mulac, R.A., and Adamczyk,3.J., "A Numerlcal SImulatlon of the Invlscid

Flow Through a Counterrotatlng Propeller,"Journal of Turbomachlnery, VoI. I08, No. 4,Oct. 1986, pp. 187-193.

19. Natnauskl, H.S., and Vaczy, C.M., "Aero-

dynamic Performance of a Counter RotatingProp-Fan," AIAA Paper 86-1550, June 1986.

20. Magllozzl, B., "Noise Characterlstics of ModelCounter-Rotatlng Prop-Fans," AIAA Paper8?-2656, Oct. 1987.

DIRECTOPERATING

COST,%0

100

80--

60 m

40 m

20 m

CREW

b"///,r ,FUEL & OILJ,/ (24%) /

Y/Y/,

MAINTENANCE

DEPRECIATION,INSURANCE& OTHER

\\

N

YEAR YEAR1973 1980

FIGURE 1. - FUEL COST ESCALATION AS PERCENTAGE OF DIRECT OPERATING

COST. U.S. DOMESTIC TRUNK OPERATIONS. (CAB DATA)

ORIGINAL P_._,.,., IS

OF POOR QUALITy

E0

oO.

0r-

_E0

_.o

r-

90 --

80 --

70

60

5O.3

/-Advanced high-/ speed turboprop/

Electra -/11 '__ "_ _

I.4

I I I I I.5 .6 .7 .8 .9

Cruise Mach number

FIGURE 2. - EFFICIENCY TRENDS FOR TURBOPROP AND TURBOFAN ENGINES.

PROPELLER/NACELLE

• AERODYNAMICS• ACOUSTICS• STRUCTURES

AIRCRAFTTRADEOFFS

DRIVE SYSTEM"PROPPITCHCHANGE,,GEARBOXES• INLETS• PUSHERENGINES

GOALS• LOWFUELCONSUMPTION

INSTALLATION •LOW0PERATM6COST NOISE ANDAERODYNAMICS • PASSEN6ERACCEPTANCE VIBRATION

• DRAG • PASSENGERNOISE• STABILITYCONTROL • COMMUNITYNOISE

FIGURE 3, - ELEMENTS NEEDED TO DEVELOP ADVANCED TURBOPROP AIR-

CRAFT.

ORIGINAL PAGE I3

OF POOR QUALITY

FIGUREq. - NASA/INDUSTRYADVANCEDTURBOPROP(ATP) PROGRAR.

PTA/GULFSTREAMGII

UDFIBOEING727 UDFIMD-80 AND578DX/MD-80

FIGURE5. - FLIGHTTESTINGOF ADVANCEDTURBOPROPS.

10

FIGURE 6. - POST-FLIGHT-TEST AREASOF ON-GOINGPROPELLERRESEARCH

AT NASA LEWIS RESEARCHCENTER.

FIGURE 7. - ADVANCED PROPELLER BLADE WIND TUNNEL MODELS.

II

u_

..Q

"13

,c-

O

0-.j

0e'-

E

-@(b

U_

O(

-2

-4 m

-6

0

1SR-2 Oil

SR-1M

S

SR-3I I

1 2

Efficiency gain, percent

FIGURE 8. - NOISE AND PERFORMANCE CHARACTERISTICS OF PROPELLER

MODELS AT MACH 0.8, 37.5 SHP/FT 2 DISK POWER LOADING, AND 800

FT/SEC TIP SPEED.

85

80

NETEFFICIENCY,

qNET75

70.60

F"---------"......... _ /- $R-7A

"_',.\ _ SR-1M-,,%-SR-6

\_- SR-2

J I J I i I.65 .70 .75 .80 .85 .90

FREE-STREAMMACH NUMBER,MO

FIGURE 9. - SINGLE-ROTATION PROPELLER PERFORMANCE.

12

MAXIMUMSOUND

PRESSURELEVEL,

dB

,,oF

160 --

150 --

140.8

O_IG!NA!. PAGE _'1,_

OF POOi_ (_L;AI.[TY

/

d I [ I J ,I•9 1.0 1.1 1.2 1.3

HELICALTIP MACH NUMBER

BLADESETTINGANGLE,

deg

O 60.1 (DESIGN)[] 57.7A 63.3

FIGURE10. - SR-7PEAKBLADEPASSINGTONEVARIATIONWITHHELICAL

TIP I'tACHNUMBER.

4 I_ADES B BLA_S

k _ I" THEORY .%\

ROIAT IO_A(S_REP[MD" [:COO '_ UNSIAI_L [ U_SIAPa [

,,> .& , ,! . 5 6 .7

FRL_ SIR_ FIA(H HUMOR

FIGURE 11. - COMPARISON OF MEASURED AND CALCULATED FLUTTER BOUND-

ARIES WITH SR3C-X2 PROPFAN MODEL.

FIGURE 12. - UNSTEADY THREE-DIMENSIONAL EULER CODE SOLUTION AT

MACH 0.8 WITH SR-3 PROPFAN AXIS AT 40 ANGLE OF ATTACK.

13

Fill--_ /-- Spar /--- Shell

/ /

Ice protection--_\

/-- Erosion

/11// protection

( /

FIGURE 13. - SCHERATIE OF THE SPAR-SHELL BLADE CONSTRUCTION

CONCEPT USED FOR THE LAP.

]4

ORiC_NAL I'I'_GE I$

OF P()O_ (_U,,\[_ITY

FIGURE 14. - LAP STATIC ROTOR TEST AT WRIGHT-PATTERSON AIR FORCE

BASE FACILITY.

TWO-BLADESR-71N MODANE,

FRANCE,WINDTUNNEL

TRANSOUCER LOCATIONS

/ ,

/ Y

(", )

STEADYPRESSURE UNSTEADYPRESSURE

INSTRUMENTATIONFOR MODANETESTING

TWO SR-7ASSEMBLIESDELIVERED

TO PTA

FIGURE 15. - NASA LARGE-SCALE ADVANCED PROPELLER (LAP) PROJECT

HIGHLIGHTS.

15

ORIGINAL PAGE I$

OF POOR QUALITY,

___J

FIGURE 16. - ROHR PTA PROPULSION SYSTEM STATIC TEST.

16

ORIGINAL PAGE IS

OF POOR (_UALITY

FIGURE 17. - PROPFANTEST ASSESSMENT(PTA) FLIGHT TESTING.

BLADEPASSING

TONE,dB

150

140

130

//

/¢

120_

1.0

--- PREDICTED WIND TUNNEL(LEWIS8x6)O PTAFLIGHT(MACH0.8; 35 000 ft)

PROPFANPLANE \UPSTREAM= I _ DOWNSTREAM \

I \I I I Y

.5 0 -.5 -1.00

FORE AND AFT LOCATIONS IN PROPELLER DIAMETERS

FIGURE 18. - COMPARISON OF PTA FUSELAGE SURFACE NOISE WITH PRE-

DICTION AND WIND TUNNEL DATA.

17

FIGURE19. - UNDUCTED FAN ENGINE (UDF).

POLYURETHANEFILMSURFACEPROTECTION

%

NICKEL LEADINGEDGE PROTECTION-,,

CARBON-FIBER/2Oe/oFIBERGLASSAIRFOIL

_- TITANIUM SPAR

FIGURE 20. - GE UDF BLADE CONSTRUCTION,

18

017, POOR QUALITY.

FIGURE 21. - UDF COUNTERROTATION PROPELLER MODEL IN NASA LEWIS

8- BY 6-FOOT WIND TUNNEL.

if

FIGURE 22. - WIND TUNNEL MODELS OF UDF COUNTERROTATION BLADE CON-

FIGURATIONS.

19

90--

NETEFFICIENCY,

PERCENT

85--

80--

75--

PERCENT OFDESIGN LOADING

80

120 -_"

"N\

?o I I I I I I.60 .65 .70 .75 .80 .85 .90

MACH NUMBER

FIGURE 23. - FT-A7 UDF RODEL BLADE PERFORHANCE RESULTS. 8 + 8

BLADE CONFIGURATION, 780 FT/SEC TIP SPEED.

FIGURE 2q. - THREE-DIRENSIONAL EULER ANALYSIS OF UOF PRESSURE

DISTRIBUT[ON.

2O

Oh. .... '_',AL PACE IS

OF POOR QUALITY

FIGURE 25. - UNSTEADY THREE-DIMENSIONALEULER SOLUTION OF PRES-

SURE FIELD JUST DOWNSTREAM OF UDF AT ONE INSTANT IN TIME.

GE STATIC TEST AT PEEBLES, OHIO

BOEING 727 FLIGHTTEST

DOUGLAS MD-80FLIGHT TEST

FIGURE 26. - NASA/GENERAL ELECTRIC UNDUCTED FAN ENGINE GROUND AND

FLIGHT TESTING.

21

BLADEPASSAGE

TONESPLAT

155-FTSIDELINE,

dB

130 --

120--

110 --

100 --

I9020 40

URCE0 /" deg PERCENT

" MXCL/

// 0 58.5155.7 100

,_ 59.0153.3 76

{" 59.2/52.8 77O 59.3/52.9 87------ 58.7157.7 1pOI I

60 80 100

SIDELINE ANGLE, deg

8×6 WIND TUNNEL

LEARJET

PREDICTEDI I

120 140

FIGURE 27. - UDF FUNDAMENTAL TONE DIRECTIVITY MEASURED IN FLIGHT,

COMPARED WITH SCALED MODEL DATA AND PREDICTION. MACH 0.72 DE-

SIGN SPEED.

FIGURE 28. - HAMILTON STANDARD CRP-XI COUNTERROTATION PROPFAN

MODEL IN UTRC WIND TUNNEL.

22

ORIGINAL PACE IS

OF POOR QUALITY

FIGURE 29. - NASA LEWIS PREDICTION CODE SIMULAIES LEADING EDGE

AND TIP VORTEX SHEDDING AT OFF-DESIGN MACH 0.2 CONDITIONS FOR

CRP-XI PROPFAN.

120 --FORWARD

SOUND 110 i

PI1ESSURE 100LEVEL,

dB

9o

8o0 1 2 3 4 5

--_ CRP

---C)--- SRP+3 dB

INTERACTION NOISE

--r'-I PLANE OF AFT

-

I I I I I1 2 3 4 5 0 1 2 3 4 5

HARMONIC OF BLADE-PASSAGE FREQUENCY

FIGURE 30. - CRP-XI MODEL NOISE TEST DATA COMPARED WITH ADJUSTED

SINGLE-ROTATION PROPFAN DATA.

23

OF POOR QUALITY

FIGURE 31. - 1987 COLLIER TROPHY AWARDED TO LEWIS RESEARCH CENTER

AND THE NASMINDUSTRY ADVANCED TURBOPROP TEAM.

24

Report Documentation PageNallorbal Ae=_)f_auh( _ and

Space Adm.,_shahun

1. Report No.

NASA TM-100929

2. Government Accession No,

4. Title and Subtitle

NASA/Industry Advanced Turboprop TechnologyProgram

7. Author(s)

Joseph A. Ziemlanskl and John B. Whltlow, Jr.

9. Performing Organization Name and Address

National Aeronautics and Space Admlnistr&tlonLewis Research Center

Cleveland, Ohlo 44135-3191

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washlngton, D.C. 20546-000]

3, Recipient's Catalog No.

5. Report Date

6. Performing Organization Code

8. Performing Organization Report No.

E-4198

10. Work Unit No.

535-03-01

11, Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplemenla_ Notes

Prepared for the 16th Congress of the International Council of Aeronautlcal

Sciences, Jerusalem, Israel, August 28 - September 2, 1988.

16. Abstract

Experlmental and analytlcai effort shows that use of advanced turboprop (propfan)

propulslon instead of conventional turbofans in the older narrow-body airline

fleet could reduce fuel consumption for thls type of aircraft by up to 50 percent.

The NASA Advanced Turboprop (ATP) program was formulated to address the key tech-

nologies requ|red for these new thin, swept-blade propeller concepts. A NASA,

industry, and unlverslty team was assembled to develop and valldate applicable

new design codes and prove by ground and f11ght test the viability of these new

propeller concepts. This paper presents some of the history of the ATP project,

an overview of some of the issues and summarizes the technology developed to

make advanced propellers viable In the hlgh-subsonic cruise speed application.

The ATP program was awarded the presttglous Robert J. Collier Trophy for thegreatest achievement in aeronautics and astronautics In America in 1987.

17 Key Words (Suggested by Author(s))

TurbopropPropellerCommercial aircraftFuel conservation

19. Security Classif (of this report)

Unclassified

NASA FORM 1626 OCT 86

18. Distribution Statement

Unclasslfled - UnllmttedSubject Category 07

20. Security Cla_if. (of this page)

Unclasslfled21. No of pages

26

"For sale by the National Technical Inlormation Service, Springfield, Virginia 22161

122. Price"

A03

National Aeronautics and

Space Administration

Lewis Research Center

Cleveland. Ohio 44135

om:_ seeing8Pqmlly kx I_va_ _ 11300

SECOND CLASS MAIL

ADDRESS CORRECTION REQUESTED

IIIIII

PoMIIge Ind Fees PlddNalk3nalA_ronautt_Sio_e .,_rninistrm_N.'_.Sk461