Advanced General Aviation Turbine Engine (GATE) concepts

NASA CR-159603 WRC Report No. 78-113-15 TABLE OF CONTENTS Section Title Paf_ Prellmtnary Design to Refine Concepts for Technological Devel- APPENDIX A KEY INFLUENCES ON A 1988 GENERAL AVIATION MARKET SCENARIO t57 APPEI_IX B vi NASA CR-15960) , • RL Report No.

78-I13-15

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LIST oi' !I.LUSTRAT[ONS

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Fi_._u re t975 and 1976 General Avtattan OEM Engine Sales Oy Type . . . . It

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1915 and 1976 General Aviation OEM Piston Engine Sates by 1975 and 1976 General Aviation OEH Piston Engine Sales by

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ProJected InfLuence of New Technology on OEH Engine Sales by

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Turboprop Parametric Design Point Study - SL/Hach O, STD Day, Turboprop Parametric Design Point Study - SL/Hach O, STD Day, Turboprop Parametric Design Polnt Study - SL/Mach O, STD Day, Turboprop Parametric Design Point Study - 7620 m (25,000 ft)/ Effect of Axial Pressure Ratio on Axial Efficiency for a Geometrically Constrained Compressor (Nominal Development) 39 Effect of Axial and Centrifugal Pressure Ratios on the Efficiency of a Low Specific Speed, Axial-Fed Centrlfugal 12 Effect of Overall Pressure Ratio and Axial Pressure Ratio on the Overall Efficiency of a Combined Axial Centrifugal Compressor 40 Turboprop Performance - Specific Fuel Consumption vs Compressor Pressure Ratio at 86% Turbine Stage Efficiency 25,000 ft 0.3 Compressor Pressure Ratio versus Number of Axial Stages for Various Work-Leve[ Stages of Compression ........... 4._ Estimated Performance, Low Cost Turboprop - SL/Standard I)av . 51 Estimated Performan_, , Low Cost Turboprop - 4572m (15,000 ft)/ NASA CR-159603 WRC Report No. 78-113-15 LIST OF ILLUSTRATIONS Title Pa_e Estimated Performance, Low Cost Turboprop - 7620m (25,000 ft) Comparison of Aerostar Piston E_gine Nacelle with Turboprop Comparison of Cougar Piston Engine Nacelle with Turboprop Turbofan Parametric Study Curves - SL/STD Day, gc = 0.76, Turbofan Parametric Study Curves - 9144 m t30,000 It)/Mach 0.6, Geared Fan - I000 Pound Thrust Class FI07 Derivative Turbofan 77 3O Low Cost Conventional Turbofan Design Concept Optimization Turbofan Engine P7808 Performance - 9144 m (30,000 ft)/STD Day. 85 Installation Drawlng - Low Cost Conventional Two-Spool Turbofan 86 Six-Place Study Airplane with Low Cost Two-Spool Turbofan GATE Family of Engines .................... [19 viii

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NASA CR-159603 WRC Report No. 78-113-15

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LISf OF ILLUSTRATIONS Title Pa_kge

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Core, Prop, Shatt, and fan Engine Preliminary DeslgrJ Schedule 139

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Avionics Required for Admittance to Different Types of Airspace 162 5O

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ix NASA CR-159603 WRC Report No. 78-113-15 LIST OF TABLES Table Title Pa_h _ II FAA Forecast of General Aviation Aircraft Production tn the U.S.A. (Baseline Scenario) ............. I0 III Detracting Features of Piston Engines and Piston-Powered IV V Oetracting Features of Contemporary Turbofans 17 VI Detracting Features of Contemporary U.S.A. Turboshaft Engines VII General Aviation Aircraft Production in the Un,ted States . 21 VIII Projected General Aviation Aircraft Engine Production 24 IX Potential Sales of Turboprop Engines in the 134 to 231 kW X Approximate List Prices for New and Remanufactured Lycoming XI XII Performance Capabilities of Piston Aircraft Before Retrofit 31 XIII Average Numher of Persons Traveling in General Aviation XIV k%' XVI Aerostar 601P Weight Breakdown Comparison (GASP Data) 62 XVII Aerostar 601P Drag Coefficient Buildup Comparison (GASP Data) 63 XVlII XIX XX XXI Mooney 201 Drag Coefficient lh_ildup Comparison (GASP Data). 69 L _- NASA CR-159603 _q<C Report No. 78-113-]5

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L_q] OF IABLES Table ] _t fe P aBe

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h XXtl 81 Iurbofan Engltle P?80_. Component Performance Sun_nary

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turbotan/"101rhop_c}p,'P_-_,,n A:,-[,lane Portc, rman,.-e Cotnp:ari._,.n (GASP) .... g?

KXrV Turbc, fan/T_lrbop,,.ptt-_._c,,n Weight Breakdown C,)mpartson

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uASP Da_) . . .

XX_d u_,ootan/TurboproplPJst0n Drag Buildup Compar t_on ,,_AbP

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_,,XVI q_ amm_i Froductio_ qu_t,c_ E_t_ates for Eaglne Pr_ _a$

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]_'Vi i 9-, Prodattion Eng£ne Pricing X_VI II Hethodolog_ for Det ermi a.,g Pet ro] eum/Oxl,'Lubrl cant s (POL)

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_6 Use. Twxn Turbofa, btud_/ Airplane P7814 ....

XXIX _¢tno,iology for Oeterminxag 20-Year Turboprop _leet POL XXX Englne-Retated lnspeltlon. Halntenance and Overhaul Cogts Pistor, Hooney 20t q9

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XXXI Enwlne-Reiated ;(h.. t'.llglI',f } l_i' [.,e¢ Lion, "_a|nterlallce a_l(l t c,verlaaul Costs - Pist,,n her,_star n0lP tOO XX.ki i t_ rD 1 n_-F rig _ ,,e ScheO, t e,/ I n.sp_ct _ on and Ma ) nt,_,,,,,i_ _ t.o_, t. I01 XXXI I] Tarbine-Engine Inspection _n,l :la]ntenance Cost 102

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XXXIV _ominal Twenty-Year Yurboprop/Ptst.,n LCC Comparison t, omponents, Nom:ual TIT ....

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f XX_V lwenty-Year Tu, bcprop/P_st,_n LCC Compac_s,)n - Improved Components, Nominal [I f ........

J XXXVI Twenty-Year Turb,,prop/Plsron I.CC Comparison - homtnal _XV I 1 Twenty-Year Turboprop/P_sr_). ICC Comparzson - _mproved tit (,_mponents, tmprovod i11" XXXV I t i Twenty-Year Fuvho:_ha/t i _fe _.v(le , ,_.t: XXX IX _HIII tlA J Ive,It', "_',t_ ],_rt,,,i II),['i>i,,t_ [,t'(" C,q]lDdl IN(_II , ,, r_ t.(,flll'uqh'l)[ S , NOIII] ,iA _ I I ¢ NASA CR-159603 WRC Report No. 78-113-15 LIST OF TABLES Table : i t 1e Pa_e XL Twenty-Year Turbofan/Piston LCC Comparison - Improved XLI Twenty-Year T::rbofan/Piston LCC Comparison - Nominal Components, Improved TI'r .......... I IO XLI I 2%¢entg-Year Turbofan/Piston LCC Comparison - Improved XLIII Low Cost Turboprop Performance Sun_aary at Haximum Rating gLIV Lo_-Cost Turbofan Performance Suamary at l_ximum Rating XLV Low Cost Turbofan Performance Comparison - 9164m (30,000 Ft)/ XL_I Frost and Salllwm Estimte of Aircraft Added to Active US i XLVII Impact of t.he 1974 Fuel Crisis and Emergency Petroleum XLVI[1 Fuel Conservation Actions InLtiated During 197a Fuel Crisis 160 XLIX Additional Act ions to be Expected with the Increasing Pre- L Expected 1988 Noise Maximums for Newly Certificated Small Business Jets (Thrust Per Fngine > 6612 N (1,5OO lbf) 167 LI Expected 1988 Noise Regulations Applicable to Newly Certificated, Turboprop-Powered General Aviation Airplanes 11_7 LII Average Nu.nber of Persons Traveling in General Aviathm xii

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NASA CR-159603 ;,_RC Report No. 78-113-t5

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SECTION 1 SUHNARY

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The most significant improvements in small [under 2-122 kg ib,O00 lbm) Rro>._

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weight] general aviation airplanes which have led to gains 112 providing essentl,_i services in the last two decades have been in the avionics realm. Advances t_, ground and airborne electronic devices, a by-product of massive federal support

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of military and space-relatea research, have made way for needed but as yet unrealized airplane performame, ut 1lily, and safety improvements._ The key to the needed airplane improvements ts believed by many to lie with the small turbine engine, but smaI1 turbine engine technoIogy is moving forward at a laggardly

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pace. The problem involves both financial and technical risks which the private sector is unwiIl.ing to take.

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A Govermaent-sepported General Aviation Turbine Engine (GATE) program has bee_ suggested as a means for stimulating the pace of stall turbine engine development.

but the mechanism for doing this is not clear. Initial GATE study activity has

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been direcued "toward delineating the proper content of a Government°sponsor.d program to develop and demonstrate advanced technologies for small general aviatlow turbine engines. The part of the study which was accomplished by NRC t s covered Ln this docent.

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The present work develops a circa 1988 general aviation market scenarao directea toward the postulation of advanced technology turboprop (T/P), turboshaft (T/_.

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and turbofan (T/F) engines and considers market needs, energy tnfluences, and the regulatory environment. The postulated engines were then defxned in terms of configuration, weight, size, performance, and cost througb trade studies of the practicality of using a single gas generator as the core for each engine type

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The study culminates in identification of tong |end, key technology elements requiring attention in an advanced engine components research program, and a plan for such a program.

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The market analysis projects" a modest growth in general aviation annual unit sales through 1988 with pronounced fleet grqwth over the period betause fleet additions can be expected to exceed retirements by a ratio of about five to one.

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This fleet growth will stimulate sales of airplanes capable of operating over a broader range of altitudes for traffic and weather avoidance. These a_rplanes el necessity will be pressurized, deiced, and faster than today's airplanes in ,_rder

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to deal with high altitude conditions. The turbine engine, and more specitlcal,_ the turboprop, will probab!y be the popular engine of this era if cost and fuel economy constraints can be surmounted.

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A family of low-cost, flat-rated T/P engines in the 134 to 261 kW (180 to 350 npl class is foreseen as meeting the needs of the era. These engines, designed for long life and warts commonality, can play a prominent role in the new airplane ","General aviation fatalities per million vehicle kilometers (or mites} is ton times that of tt,e automobile.

NASA CR-159603 WRC Report No. 78-1] 3-15 and retrofit markets. Low cost would be realized through the exploitation of advantages gained from the relaxation o{ design stress and temperature levels (e.g., by the selection of alt.ernate materials) antl through a t:e_' anti unique approach to component manufacturing. This ._pproach requtres that engine aero- dynamic components be geometrically constrained to enhance producibitity. The trade-off involves cost versus component efficiency and engine weight, where the weight influence on marketability can be shown to be mlnimat The req_lreme_tt for geometric constraint affects axial compressor and turbine bJadir_g and involves constant camber, constant cross section, chord-taoer, and twist consJ, derat_on> A Government-sponsored research activi ty to maximize component aerodynamic etf t- ciency through the optimization of the constrained geometry is suggested.

The performance of proposed TIP and T/F engines using a common core was defined to make possible airplane application and cost studies for evaluation of the merit of the engine concepts. These studies were made using the NASA-developed General Avintion Synthesis Program (GASP) and, for the T/P work, involved retro- fitted contemporary single- and twin-engine airplanes. The studies showed that both the turboprop- and turbofan-powered airpl:mes exhibit significant life cycle cost (LCC) economies when compared wtth piston-powered counterparts, while demon- strating new dimensions in performance capability. Also, the T/P airplanes, It flown at high altitude, are more productive in terms of seat-km/l (seat-nm/gal) The major contributing factor to the LCC benefits is the lower turbine engine and turbine-powered airplane maintenance _ost.

IJntil the present time there has been little govermnent support for upgrading general aviation powerplant technol'ogy, and this can be justified in light of the evolving stale of the national airspace system. This system as now quite sophis- ticated and getting more so with each passing day. The work here shows that general aviation, already ., vital tool oi commerce, is ready, owing to airspace system advances, to beneftt from a new generation of powerplants. This technology needs to be nurtured through government supDor't, however, because the risk is far more than mdustry is w-ll_ng to accept.

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NASA CR-159603 WRC Report N_. 78-1!3-15 SECTiON 2 IN'[RODUCTION A 1975 Federal Aviat ton Administrat ion (FAA) census revealed tt, :t tt:_re _., ,_ i7_. [56 active aircraft _._ tile" [;.S. c_vil fleet: and only q, pe,,ent ot tt, e._,.

_r,c;uding _9 percent of the helicopters, were turbine-powered !_iere were very few turbine-powered, fixed-wing airplanes in the under 2722 kg te_.O00 ibm) weight class, a (lass dominated by tirplanes with six or fewer seats and making up qo percent of the entire general aviation fleet.

Desplte such statistics, the marketing flies of major turbine englne manulacturerb bulge _ith airframe company inquiries about progress being made toward the devel- opment of potentially viable small liP and T/F engines for general aviation Pilots, too, are interested in progress in small turbine engines, because they are acutely aware of the benefits turbine power offers from a safety-of-flight/weather- avoidance standpoint.

The lack of small turbine-powered airplanes, despite designers" and pilots intense interest in them, provides clear e-._idence that obstacles exist that ate i impeding their emergence. The acknowledged foremost obstacle ts the lack of _matt turbine engines at a marketable cost. The high cost is influenced by development, tooling, materials, certification, and accessories costs. Wh_ie operationally acceptable engines can be devei,_ped using traditional techniques, little compotttl,_n _s otfered the less costly pistun engine alternative. Clearly. a unique appr, _rh _s needed to develop a small turbine engine that wiIi make a significant marmet penetration.

Fuel efficiency is a second major obstacle _nh_blting the proliferation ,,_ small turbine engines and is becoming _ncreasingly significant wath t_e steady ad, aace in fuel prices. Yet, bet-ause of the superiority of the turbine engine fr-m an operational standpoint, fuel consumption parity with the p_aton engine does ,_o', appear to be an absolute necessity. Airplane utility and productivity must be considered, in addition to the often mistlsed seat-km/l (seat-nm/gal) parameter General aviation has made great progress over the last 30 years. The major part of this has to be attributed to advances in avionics and a modernization ot th_ airport/airways system. Here dramatic advances, abetted by the ongoing av_,m_(" revolution and a nearly 4 biIlion dollar accumulation in the Airport/Airways _rust Fund,"r are promised for the future.

The under-2722 kg (6,000 lbm) airplane, which has undergone l lttle change tn three decades, needs new technology powerplants before it can realize a potent,al commensurate with that promised through avionics advances and airport/a_r_avs development. The airport/a_rways system can now handle modern, high-speed _,r- planes saI-ly, and these _rplanes _an surely be built in small, fuel-elt_ ,ent sizes if, ao.d when, small turbine engines become available. A govermnent sti,r, ulus through support of an advanced technology program is clearly needed t,_ lessen .:A, rport/Airways Trust Fund status as of 31 December 1978.

NASA CR-159603 WRC Report No. 78-113-15 current business risks and revitalize a nearly stagnant sinai1 engine/small azrplane technology. The payoff will be safer, more useful, superior performing general aviation airplanes that benefit from the advantages gained from net,, technology turbine powerpIants. These airpianes will assure continued U.S. leadership of the world's general aviation market and place general aviation in a valued posit lt n among the various modes of transportation.

In an eflort to explore ways to accelerate the emergence of small turbine engines, NASA has asket GATE contractors for assessments of the option to use a single gus generator as the core for T/P, T/S, and T/F engines. It has also asked for concep- tual engines of the three types and the identification of critical components, high risk items, and long lead key technology elements. Recommendations are also sought with regard to an advanced engine components research program, including a schedule and projected costs for component design, fabrication and test, and an engine test program. The Williams Research Corporation answer to NASA's request for ieformtion is developed in the succeeding sections.

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NASA CR-159603 WRC Report No. 78-113-15 SECTION 3 STUDY PROCEDtJRES The study postutat,,d the, tutt_e general aviation market to enablv id_'nt _fic:;t ,_m ot the most appropriate turbine engine sizes and configurations for subsequent parametric work and conceptual tzation. It included eng,ne and ai rplane con_ et,- tualizatlon, performance predictions, and the evaluation ot _omeptual engines that could be adapted to _ _ ommon core concept intended to attack the t _. e,, ne engine cost problem. Terhn,,l,,gles appropriate for Government sponsorship _ ,_: would reduce engtne development risks to levels acceptahl, for continuation _,_,k ill _ by private tndustry were identified, and a technology program pJ.,n was dew loped In specific terms, the study was broken down into four major tasks identified as follows" Task 1.0 Harket Analysis Task 2.0 Broad Scope Trade-off Studies task 3.0 Evaluation of a Common Core Concept Task 4.0 Technology Program Plan "D_RKETANALYSIS A circa 1988 market scenario for general aviation powerplants was forecast ,l, accordance with a six-step approach that included: ° Acquisition of availabl.e projections of the .,_t ,re f teet _n _t_88.

including fleet composition, annual addit ion. and the compostt _on ,,l added aircraft.

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An assessment of the projections in the light of pos>tble market iull,- ences involving energy aval labi ] it.y_ user charges, a, rport/a_ r_,ays development, the regulatory environmeat (including that associated with no_se and emss_ons), economic regu!ations, new technology, and product liability.

3. Engine distribution projections by power level and type.

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ldent,fication of qualities lacking in present airtraft and aircral_ powerplants that could influence the course of powerplant technology.

The postulation of advanced technology engines including dest, _' ,. _ turbine engine sizes and configurations which take into account m_,ket needs, the potential retrofit market, energy influences, and the regulatory environment.

The detecmin;_t 1.011 Ol .tl I'! ri_'lTit'r illt,:'l't?:_t lit I hc t,o.';tu] ,it od eng_n_:. _;, an _.,,.sessmont of the pote_t:ial imp_,.'t c,t the p_,:_tulated p_werplan,._ _.

the projetted d_strtDtit_oh of engines by p(,_,e_ _cw'l dlld type.

i i, NASA CR-159603 WRC Report No. 78-113-15 Table I lists some of the information sources used for the market analysis and highlights specific data base material obtained from each.

At the conclusion of Task 1.0, candidate engines, airplanes, and mission profiles were identified for use as major inputs for performing the Broad Scope Trade-Off Studies.

BROAD SCOPE TRADE-OFF STUDIES Parametric studies for projected 1988 state-of-the-art general aviation turbine engines were conducted, using massion profile and aircraft characteristics data developed during Task 1.0, to aid in engine cycle optimization and sizing analy- ses. Study limits involved TIN and T/P engines in the 112 kW to 7/,6 kW (150 to 1,000 hp) range and T/F engines with cores sized comparably to the T/S engines.

Eaphasis was placed on propulsors produ_i_ less than _ kW (60G hp) and on turbofans producing less than6672 N (l,5OOlbf) thrust.

TABLE I DATA BASE SOURCES GENERAL AVIATION ff#a_IUFACTURERS ASSN. (GAHA) AEROSPACE INDUSTRIES ASSOCIATION (A1A) (GA stJtistics) (lnd_tries posttim re_rdxngno_se and emissions) AIRCRAFT OWNERS AND PILOTS ASSOCIATION (AOPA) GULFSTREAN AMERICAN CORPORATION (Business .jet noise) (Cougar technical data and turbofan assessment) AMERICAN PETROLEUH INSTITUTE (APt) (Impact of automotive diesel on fuel costs_ HARTZELL PROPELLER, INC.

DEPARTHENT OF COHBERCE, (Propeller cost, weight, and BUREAU OF CENSUS, technnlogy data) INDUSTRY DIVISION HELICOPTER ASSOCIATION OF hd_ERICA, INC. (HAA) (GA statistics) (Helicopter engxne_related safety information) ENVIRONMENTAL PROTECTION AGENCY (EPA) (Aircraft noise and emission proposals) NOONEY AIRCRAFT CORPORATION FEDERAL AVIATION ADMINISTRATION (FAA) (Model 201 technical data) (GA forecasts through 1988) NATIONAL BUSINESS AIRCRAFT ASSOCIATION (NBAA) FEDERAL Eh_RGY ADMINIS1RATION (Energy influence torecasts) (Fuel cost forecasts) PIPER AIRCRAFT CORPORATION FLYING MAGAZINE ANNUAL (Aerostar 601P technical datal (GA airplane data) _"E WEE :g OF BUSINESS AVIATION ,,s'orical and stattstical FORECAST ASSOCIATES, INC.

nfornation including _nnual (GA forecasts through 1982) ,,tr.tane production figtwes) FOX JET INTERNATIONAl.

WILLI,_PI P I.EAR, SR.

(Advanced technology turbofan (SInall turbofan engirle/alrptaile assessment) as._essment ) FROST AND S[!I,LIVAN, INC.

(CA forecasts through 1983)

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NASA CR-1596013 WRC Report No. 78-I]3-15 DuriP.g the parametric work, conceptLJ:_l layouts were made of '[,P, T/S, and T/F engines, and a commot; core concept was developed. CandLdate engine perfc, t-_ance, weight, and cost estimates were made as the designs matured. F,vort:,t designs were subsequently mated with appropriate atrcraft, and al rcraft performance was calculated. The, festal t ing per_ormance was compared to des t red charac'ertstics data estahltshed during the _'larket Analysis and to the perform;it|co of piston- powered a_rplane cour_terparcs i'tnally, turbine engtne-related _tfe cycle costs were estimated anti compared _o corresponding piston engine-related life cycle costs.

Several steps were taken to develop high-confidence T/P airplane performance figures to enable value judgments to be formed on the merits of the proposed t,_w- cost T/P engine from a fuel economy standpoint. These mvolved the mat,rig _o_a paper) of the turboprop to existing airplanes, namely, the Aerostar 601P and the My 201. Limitedworkwas also d_ne with the GulfstreamAmer_can Cougar.

Prior to calculating YIP a_rplane performance, the piston airplane perfcrma.ce was calculated using the _ASA-developed General Aviation Synthesis Program (GASP). The GASP-derived data was baselined to FAA-approved airplane flight manuaI data for the airplanes of interest through appropriate input adjustments.

The resulting inputs were used to c.,lculate T/P airplane performance by me_ns of _ASP, with changes made only Lo account for airframe and powerplant differences (e.g., powerplant-, po_erplant installation-, prolmIler-,, _a4 pressurization- related weaghts; propeller and natelte size; and cooling drag) EVALbA_iON OF A COHHON CORE COM-EPT A core was sized and conceptualized and then T/P, T/S, and T/F extensions of the core were conceptualized. Performance, weight, and cost estimates were made for candidate T/P and T/F configurations and some candidates were eliminated on tht basis of these estimates. Fuel efficiency and cost were emphasized in the T/P design, with weight considered of secondary importance since T/P engines tend to be much lighter than comparable piston engines. For some retrofit applications, a weight savings can be of dubious value because of airplane balance considerations (e.g., single engine airplane retrofits). Weight, however, is considered extremely important with respect to T/F designs, because these engines are typically aft- fuselage-mounted and excess weight aggravates airplane balance and stabi ,ty problems.

Several iterations were made involving surviving concepts until a near opLHnctm core was conceptualized which was common to a fixed-shaft turboprop, a conventional two-spool turbofan, and a free-turbine turhoshaft.

TECHNOLOGY PROGRAM PLAN A master schedule, portraying the logical and sequential development of the three types of turbine engines, was developed based on the common core design concept The schedule was developed to provide the broad overview of GATE program activities necessary for successful del. ivery el certified engines which cotlI(] meet the demands of general aviation _n the late 1980 s.

NASA CR-159603 WRC Report No. 78-113-15 Besides the master schedule, several lower-tier schedules were formulated to illustrate the program planning in greater detail. Also, schedules were developed to show the proper content of a Gover_ent-sponsored program to develop and demonstrate advanced technologies for small-sized general aviation turbine en- gines. Critical components and nlgh-risk items were identified and recommendations made as to the content of an advanced engine components research program.

An overall plan was developed that identifies schedules and the projected costs for component design, fabrication, and test. An engine test program was formu- lated and long-lead technology eleu_nts identified.

NASA CR-!59603 WRC Report No. 78-113-15 SECTION 4 _T ANAI.YSIS Prior to the initiation of the market aria lvs is t(, postulate _ 198,_ _en¢'ra i aviation market scenario, a data base ot pertinent recent pai) l _cat_ons was gathered to enable the most credible projections. Specifi(aily. data was s,)ught which wouId permit a breakdown of the current general aviation powerptant mar_et into engine type and horsepower caterer tes. Ad,l_ t _ona I ly, t rends _a emtss _ ons and noise regulations, energy constraints, user fees, and a_rcraft equipment and operating requirements were sought, together w_th existtn_ _overnment and industry forecasts.

AVAILABLE FLEET PROJECTIONS i Three documents were acquired that provided general aviation fleet projection data into the 1980's. T_o of these furnished projections into the early iqgG's and were prepared by private market resemrch firms (ret. ! and 2). The third was prepared b_ the Office of Aviation Policy of the Federal Aviation AdministratLo_,.

(FAA-AVP) a_d provides projections through 1988 (ref. _). The FAA projections were _ven the most weigh_ because of the continuing work of FAA-AVP in the area of forecasts, forecasting methodologies, development of new data sources, and initimt_em for involving members of the aviation community in forecasting for decision-making. Also, the FAA has privileged _nformation through aircraft and engine type certificate applications that is not available to private market research firms.

Because FAA predictions are ased in budgeting and managing the National Aviation System and by state, regional, and local decision-makers as welI as by those _r_ the aircratt _ndustry, the forecasts tend toward being self-fulfilling. Fable il provides the forecast of general aviation aircraft production in the bn_ted States developed as a baseline scenario by FAA. Total production figures derive,: from this scenario were uItimately adopted for use in the market analys_- Because FAA has deveioped alternative scenarios (e.g., an energy scenario and a, economic stimulation scenario), factors were examined that could perturb the baseline scenario. These are discussed in Appendix A to put the market ana vsas in perspective.

ENGINE DISTRIBUTION PROJECTION In order to estimate the fleet composition and attendant engine requirement_ by horsepower and type in 1988, the composition of fleet additions made in 1975 and 1976 was determined. Also such influences as current engine deficiencies that might affect the engine mix tn 1988 were studied.

Key documents (ref. 4 through 8) were used to determine the 1975 and 1976 compo- sition of engine deliveries for new aircraft by horsepower and type These documents provide airplane delivery figures and associated engine data a:_ NASA CR-159603 WRC Report No. 78-113-15 TABLE If. FAA FORECAST OF GENERAL AVIATION AIRCRAFT PRODUCTION IN THE U.S.A.

(BASELINE SCENARIO) Piston Turboprop Turbojet Rotary Wing Twin Twin & Piston & Single Twin & Turbine Total Fiscal MuLti Engine Multi Engine Year Engine Engine 1972 132 74 444 8,856 6,901 1,305 1973 221 157 602 9,472 2,017 12,469 1974 560 199 721 11,092 2,158 14,730 1975 513 198 794 11,824 1,903 15,232 1976 616 615 12,150 1,879 15,446 199 74 182 1977"_ 559 3,742 4,786 1977" 756 279 684 12,716 2,016 16,451 1978" 743 267 656 12,640 2,136 16,442 1979" 283 684 11,692 2,271 15,691 1980" 820 329 762 2,403 12,078 16,392 | 1981 _ 909 392 2,482 12,997 17,644 886 378 833 1982" 16.222 2,178 11,947 1983 _ 860 358 798 2,130 11,916 16,062 405 861 1984" 936 2,550 12,933 17,685 464 935 1985" 2,624 18,355 13,290 1,042 1986" 13,647 2,698 1,119 509 1,004 18,977 1987" 13,986 2,789 1,194 553 1,075 19,597 1988" 2,880 19,894 13,995 1,271 598 1,150 *Forecast 19771 "_-- Is the transition quarter from i July 1976 through 30 September 1976.

NOTE--General aviation aircraft for export are included. All helicopter productzon, including air carrier transport helicopters, is included. (Total production figures and rotary wing production figures from this forecast adopted by WRC) i0

NASA CR-159603

[,_C Report No. 78-113-15

supplied by Beech, Beilanca, CessmJ, Gates Learjet, Grumman Amrricar, (now Gulf- stream _nerican), Lake, Maule, Mooney, Piper, Rockwell, Ted Smith (now Piper), and Swearingen. A total of [6,260 engim,s delivered in [975 and !7,768 e[_ginc, s delivered in iq70 were involved in the determination of engine distribution by type, (Figure 1). The engine distribution by kW (hp) involved t5,272 p_st,,n engines delivered in 1975 a_ld ,0,676 piston engines delivered in 1976.

T URBOJET-- TURBOPROP TURBOJET ------7 f--T URBOPROP z. ;o%

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,_. o4% / !

I PISTON PISTON \ \ 93.86% 93 _, \ 93. o \ N \ \ t/ USA 1975 -- 1976 A-6737 Figure I. 1975 and 1976 General Aviation OEM Engine Sales by Type Piston engine distributions by kW (hp) were determined for several combinattona of kW (hp) ranges to demonstrate the sensitivity of the distributions to the specific ranges selected. Considerable sensitivity was noted when a range was shifted to include a popular aircraft model (Figure 2).

L NASA CR-159603 WRC Report No. 78-113-15 448-746 kW ----.----- --299--447 kW 448--746 kW kW (601--1000 hp (401--600 hp _ (601--1000 hp % (401--600 hp '_ O. 29% 2. 15% O. 13% 2.43% - r - ..

,/ ,/ / / 0--112 kW / / (0--150 hp ) / _,,, 7_, 172--298 wW i ,-_..., o (23 I--.400 hp _,!

!

/ 40. 73%

113--172 kW / (151--230 hp 3 ./ ,,, (151--230 hp )\ / _27. 10% \ /

_7. 43% /

1975 USA 1976 (NOTE INFLUENCE ON THE 1976 DISTRIBUTION OF THE MORE THAN 3100 AIRPL#NES SOLD WITH 112 KW (150 HP_ ENGINES_ 0-111 kW -- 299--746 kW 0--111 kW _ _ 299--746 kW |

(o-149 hp _ | (401--1000 hp

11.64% 2. 44% (0--1491 I. 02°_hP_ t _ 2.(401--100056_,_6 hp 'l

i

i f-- \, /" 176--298 kW (236--400 hp i _ (236--400 hp ') i 38.89_/_

! 112-17sww \38.e7%

i 112--175 kW l (150-23s .p) \, \ (150--235hp) _ 47. 759/o \ \

,\

1975 USA 1976 A_10808 Figure 2. 1975 and 1976 Ceneral Aviation OE,_!Piston Engine Sales by I]or._e;_,,.,er ]'2 NASA CR-i59603 WRC Report No. 78-213-]5 A specific kW (hp) range combination was selected ;is a base_tne for turther work, (Figure 3). The selection was made on tile basis of engine design studies that indicated the feaslbiltty of pro, dating at least three turboprop (T/P) powerplant models rated (tlat-rated) between 134 alld 231 kW ( 180 and 310 hp) using variants of a single, basic (T/P) engine design. Thr basic design, i i developed to the point of be:ng competitive with approprl,_tely pov'ere'd p istc, n engines from a cost/fuel economy standpoint , would vie f(_r 58 }_t.f, ent ot tn,' genera i aviat _on new-engine sales ]ncluding the segment represented by _,2 percent of the piston engines for newly manufactured aircraft.

SALES TO BEECH, BELLANCA, CESSNA, GRUMMAN AMERICAN, LAKE, MAULE.

MOONEY, PIPER, ROCKWELL AND TED SMITH ;i 16, 676 UNITS 15, 272 UNITS g+-_ .

232--746 232--746 kW

(31I-iOOO

(311--1000 hp)

8.oI%

8. I,S%

O-- 133.5 kW 0--133.5 kW 1335 UNITS 1243 UNITS (0--179 hp) (0--179 hp) 29. 73% 29. 7 I% 4541 UNITS 4955 UNITS 13a--._ 1 kW 134- -231 kW (180- ,) "" rip) (180--3 10 hp)

28%

62. 13 `_' I0, 386 UNITS 9a88 UNITS 1975 CY USA 1976 CY A--6892 A Figure 3. 1975 and 1976 General Aviation OEM Piston Engine Sales by Horsepower

NASA CR-159603

_4RC Report No. 78-113-15

The nearly identical 1975 and 197_ plstou engiue kW (tip) distribution d,_ta was assumed similar to what might be expected in 1988 wht:,ther the then existing engines are predominant [y p_ston, tnrboorop, rotary, or other. The asstimpt ion of similarity was made in the lace o[ .-;everal contradictory Lnfiuences.

Rapidly escalating fuel _,_,s, ior instance, ace _nfluenctng buyers toward the smaller, more fuel-eft [c-ent s_,gt<'-engine aiz'planes, and this trend is being supported by a_rcraft mam.facturer act:ons to improve their airplanes aerodynam- ically. Also, actions are beipg taken to reduce the governed rpms of some engines because of a noise reduction need. The result, tn some cases, has been that output power was ceduced.

The projected increase in business-use sales relative to private-use saes, to the contrary, favor the large-engine and twin-engzne airplane market. Also, the decreasing availability of 80-octane aviation gas caused engine manufacturers to discontinue the production of low RW (hp) engines (e.g., the Continental 0-200) or to uprate some of the low kN (hp) models to enable use of lO0-octane low-lead avgas.

QUALITIES WANTING IN PRESENT AIRCRAFT AND AIRCRAFT POWERPL_TS Piston engines produced for general aviatzon by Ly_omlng and Continental are, without a doubt, the finest available anywhere, as evidenced Dy major airframe manufacturer preferences. The general acceptance of these powerplants has led to their use in over 90 percent of the general aviation aircraft being built in the free world. Despite th_s unparalleled acceptance, there are a numb_-r of powerplant-related factors that limit the usefulness and affect the saie_y ol contemporary general aviation a_rcraft. Some of these factors are unique to the piston engine and others apply equally to alternate means ot propulsion. Detrac- ting piston engine character[st_(s are listed in '[able [1[ betracting features of turboprop (T/P), turbotan _T/F), and t,rboshaft (T/S_ enRtnes are shown in Tables IV, V and and Vl. The T/S performance dell ciencJ es t i sted apply to specific U.S.-manufactured T/S engines produclng less than 7&6 kW (1,000 shp).

In the world market, there are several French-made T/S engines rated at less than 746 kW (1,000 sap) which are doing an excellent job tn helicopters and which are making inroads into the U.S. market. These include the Turbomeca Astazou IIA, Artouste IIIB, and Arriel.

ADVANCED TECHNOLOGY ENGINE POSTULATION A single factor, the ever-rising cost of loss11 fuels, stands out as the key influence in the postulation of general aviation po_'erplants for the late 1980's. As a consequence, airplane fuel efficiency and utility must be stre¢- sed. Airplane fuel efficiency is emphasized, since installed engine performance is much more meaningful than, simply, engine performance. Engine weight and cooling drag along with specific fuel consumption have significance.

Engineers are prone to evaluate airplanes for tuel efltciency lt_ terms ot tan_lbie parameters such as Mll/I _nll_/glit ) and st,at-kin/1 {seat -lun/g.tl ) . ]'h,-.qt, yard>t leas, while [IAVIII_ fll('l'l[ , t'X{ il',t_" th(' iI_ II.('[l_'_' .)l .ti !'}'l,tll_' p{ I'l,'l'll_itlIC,' ,'ii ti[I l _t\',

l

"<ASA CR-!59603 ',,'I,_C Report No. 78-11"3-15 TABLE III. DETKNCTIN(; FEATURES OF PISTON ENCINES AND PISTON-POKERED AIRPL,MNES • Co'.d w,',_!'.'t- _;t:tr: .... /. ::, ;J(_ i.u,t, a_ .',.],_r:lt.c_ cngi::_- wear. _'. ! _,:t_';, req_iirev '2I'lg[!It." l'Y¢_}/,/,'_ _I_.,i'/:,I" ,':,;_.'i'D;i! i)OWt)l'.

0 F}li!r0 [.g 1[ tlv.. t-Lis(_! iIig [<,;]tJ,_?ni'v :ltlFirlg Fglpid ,ie'_;_t*:_ia in , old .'_iF wil:i tile poasibititv o: enFh_c damage.

@ C._rburetod cng!ne,_ arc subjec_ to _toppagc dt_e i,, _ ,_rb_Jrt,tor ,,:inca.

o • Aircraft windshk¢:!d defogging a:TJ, cabin heat ar= _ ,:,_r_nonlv providec b\' an inherent- ly hazardous exhaust-gas aeat ex,:nanger.

• Fue!-injeeted engine_ are _ common cause of post-crash fires because of the dispersal and ignition of atomized fuel from broken injector ruder.

i+ • Hot fuel-inie_.ted englne.-_ are difficult to restart because of inject,,r tube 7_ vapor lock. Air restarts after the exhaustion of f.et fr,_ ,me tank and a switch t,, a second tank can be excessivelx timo-tonsuming.

X Engine vibrations ca::se airframe, fatigue era, ks, tube am: wirc chafing, and ltrait Frc pel!er TB{, ; *,_ot_t 1,000 hours. ('r/_ propellers t,;_icliiv h_e 3.0_O hm:r 1"_O'-.; • Pi._ton onginps, especia_t\' t_.'r4ocharged piston ,.,nai;_c+_, are lwavv an,i 1-.ulk'+'.

• Netal propellers arc 7e.*v'.. ,_._;Ix, ni,'ke,!, ar:d des,.rving of rcspect b?' ground personne I, p._ssenger_4, and "re_.

• rh. iex_,-ssivc ti,_ _i-;t !a!',sc rate witi7 _i :-;petal of _.r,'pci }t l-driv,,;;, i:ibt,,l-+- paw,-red airpt:_n_'>: ,'.,nsrrain_; crui._¢ :rcrf,,r;'=+/ll:, c aml limit> ti_t- v,.iume ,,f airsF+a,'e av,_i:ablt + t,, tbt. pl]ol :or wo_iher _v,,id..,+m',, ,-.n:_ p:_songer ,.,mr,,:'t.

• Prope_h.rs prod,we t,,rque, gvr,,s_.,pi, mom,,nts and '.'._wi;:a ,:!_':nt.nt_ ,It i,'x I_ it;!!t speeds ,dtit_ t,, "P" ,.:lect), making thu pi],;ting lask m,'re dill icuit.

• ('untemporar': v.ngin<.,pr,,i_ell_.r combillations ]+,l'o_Jtlct' ,%Rt't'_M_iVt * ll_71:;e, t'>pvci/lllv in the aircr,_,+l cabin _tx, picaily 8_-'_7 dB(A _, .

• Engino-m_t asvmn:,,_t-_cal thr:_t ,m twir,-t.:-.,gine, pr,,;;cll,,r-driv,.:_ .-_irl.,1..l|_,+_ creates .a conditi,:-+ requiring adcpt pilotii+oc, tN,,t,.: Reducing propelltr.ttuselagt, clearance, and the ._tten,la:K vawirlg nioment duo t,+ :t:;y.,rmletri,'ai thrust result_ in high in-cabin noise _evels,) Landi,lg gear length and weight t._ ._on,_.times [nfl_en,'ed by pr,,p, tier grotn:d clearance rcquiremt+nts with an ._ttendant influcnce _,n .'wailaI, ie wil1_ \'Olll,.'2t: for fuel _torage in rcrractnble gear airplanus.

Piston e:lgine co(:ling _lrag i_: ++xct's_ive (typi.'allv q t., 20 p,.rcent ,+I cruise drag) and propcllc.r at+rf,_t-m._n,-, • is ,Jcb:l'._ldt'd }_V a++ttq'-x,'o,iv ir.:_u,qlcc_..

l'iStoll+pox,'ol+cd aiy:,ial+t.,; l'+,qi_tr,, palJitipl, + :'.,'wt'l'pl,2llt t_,t'.tr:'[:-; Ic.,p.+ Pltr, "'.[. + .

pr_p,ller, ll:i>:tHl,'_ _,'ll'!_ql'wt_r _,/,'.IL. died d,,t0i i+;lpv:./, I ,u>. _ddi!t_2 t:' ;'_ i;'t w,-rk 1 oad.

NASA CR-159603 WRC Report No. 78-113-15 TABLE IV. DETRACTING FEATURES OF CONTENPORARY fURBOPROPS • A high first cost which tends to drive airplane cost beyond the means of most small airplane operators.

• High overhaul costs (typically 50 to 70 perce t of new engtne cost).

• High specific fuel consumption [typically 0 34 to 0.38 kg/kW-h (0 5C to 0.63 lbm/hp-h) at takeoff rating].

• High fuel consumption for flight operations at altitudes belo_ .5,000 feet where most non-instrument-rated pilots (7] percent of all p_|ots) operate.

i ,i ii i ,i i . Valnerability to inlet icing and FOD.

m • Fixed-shaft models are noisy, especially during ground operations.

In-cabin noise levels exceed turbofan airplane cabin nolse levels.

• Metal propellers are heavy, easily nicked, and deserving of respect by ground personnel, passengers, and crew.

• The thrust lapse rate wltb airspeed is excessive.

• Propellers produce torque, gyroscopic moments, and yawing moments at low flight speeds (due to "P" effect) making the piloting task more difficult.

• Engine-out asymmetrical thrust on twin-engine, propeller-driven a_r- planes creates a condition requiring adept piloting (Note: Reduczng propeIler/ fuset:_ge clearance and the attendant yawing moment due to asymmetrical thrust results in high in-cabin noise levels.)

• Landing gear length and weight is sometimes influenced by p_opeller ground clearance requirements with an attendant influence on available wing volume for fuel storage in retractable gear airplanes.

• Ground starts with a dead battery and without external power are not possible. The engine cannot be hand-propped for starting.

!

NASA CR-159603 _C Eeport No. 7_-I13-15 TABLE V. DET_\CTING FEATURES OF CONTEMPORARY TURBOFANS • A high first cost which tends to drive airplane cost beyond the means of most small airplane operators.

• Excessive fueI consumption.

• Vulnerability to inlet iclng and FOD.

• Excessive performance degradation due to high temperatures.

• Inadequate takeoff thrust for small airport operations.

• Idle thrust is sometimes excessive.

• Responsiveness for go-arounds is poorer than piston and turboprop engines.

• Thrust-reversing capability is expensive to incorporate.

• Windmilling starts can involve considerable altitude loss. j not • Ground starts with a dead battery and without external power are

J

possible.

TABLE VI. DETRACTING FEATURES OF CONTEMPORARY U.S.A. TURBOSHAFT ENGINES BELOW 746 kW (iO00 shp) • High first cost.

• High overhaul cost.

• Excessive performance degradation due to high temperature and altitude in combinatlon with inadequate performance reserve.

• Inadequate engine life because many operations are carried on at, or near, maxim_n rated power.

"i NASA CR-159603 WRC Report No. 78-I13-15 safety, the quality of ride, and the time necessary to make a trip. Indeed, the atmosphere within which airplanes fly is not homogeneous, nor is it two-dimen- sional. Flight is conducted in three dimensions, and within that airspace violent weather capable of tearing airplanes apart as well as quiescent serenity can be found. The ability of an airplane to operate clear of stormy areas is of paramonnt importance to passengers and crew alike, and this ability relates to the power available for climb and operations at high altitude. Considerable fuel is routinely saved by flying over bad weather instead of circumventing it_ The J-3 Cub can be described as very fuel-efficient in terms of the km/l (nm/gal) and seat-kJm/l (seat-nm/gal) yardsticks, but is hardly competitive as a useful people- moving conveyance except for very specialized applications. Intangibles cannot be overlooked when the subject is airplane fuel efficiency.

Because altitude capability is the key to passenger-carrying utilzty, the general aviation powerplant of the future must permit rapid climbs to high altitude whxle smgplyiq the power needed for avionics, cabin pressurization, and ice removal.

Fuel efficiency .does not become totally relevant until these prerequisites are met. _ine flat ratiq to 4,572 m (15,000 ft) or more altitude and bleed air avai1_bility with linimmperformancedegradetioncouldbe key c_siderations.

In the past, businesses have had to purchase airplanes that were, perhaps, later than needed i_ order to move about in a manner rivaling the airlines for schedule reliability and comfort. [The average business jet flight carries 3.4 passengers about _I 1_ (49i rim]. This was simply because there were no small airplanes having adequate performance capability. Airport/airways system improve- ments, together with rapid developments in avionics and instrumentation, now make possible the development of small airplanes that can move about with the surety of larger airplanes while saving considerable fuel in the process. Candidate engines for these airplanes need not have superior fuel specifics, although close-to-large-engine and piston-engine economy are desirable. Fuel can be saved simply by matching airplane size to the passenger-carrying requirements of busi- ness. Additional fuel can be saved if the small airplane can be made easy enough to fly to permit the bosinessman to fly it himself. Small airplane product- ivity, in terms of passenger seat-km/1 (seat-um/gal), can hardly be considered acceptable when professional crew members occupy as many as one-third of the available seats. With continuing advances in avionics, autopilot technology, airplane and powerplant design, weather observation techniques, and air traffic control, there should be less need for professional crews in small airplanes in the late 1980's than today. Perhaps at some point, professional small airplane pilots will be no more needed than chauffeurs for the family automobile.

In selecting a general aviation turbine engine type and size for the late 1980's, expected unit sales and sales price become key considerations. Sales-related matters involve the adaptability of the engine to single-engine and multi-engine airplanes, and to the retrofit market. The type of flying expected during the era will also influence the choice of engine type/size. Current trends suggest that instrument and night flying will be routine, with an attendant demand for systems redundancy (e.g. avionics, electrical, hydraulic) and twin engine reli- ability. Unfortunately, high fuel and engine/airplane costs will counter this demand.

NASA CR-159603

_,NCLeport No. 78-1!3-15

Single-engtne instrllment and night flying require high equ[pmenL reliability.

Purely routine operations can be shattered by con_nunication, electrlcaI system, or powerplant malfunctions, and the possibility of these has a sobering psycho- logical influence on pilots and passengers alike. )lost expel tented instrument- rated pilots would prefer the added safety provided by two engines, two alter- stets or generat_rs, two cotrdnunications transceivers, etc. Consequently, any synthesis of a fuei-effic_.enL airplane sized for the travel requirements of business and capable of airline-like ontime performance must inclade two properly sized engines. Few large-corporation presidents will travel routinely in single- engine airplanes. They may be wilIing to ride in small twin-engine airplanes, however, if trips can be made swiftly, safety, and comfortably.

In adapting turbine engines to a retrofit market as well as to new slngle- and twin-engine airplane designs, the turboprop has more merit than the turbofan.

This is because of the relatxve ,ase wxth which turboprops can be substituted for piston engines in existing designs (especially twin-engine airplanes). As a result of the light weight of candldate turboprop engines, variants of a single basic turboprop engine desig_ have the potential for use in all newly manufac- tured, four-to-seven-place, singIe-engine airplanes and twins seating up to ten persons. This kind of adaptability yields significant cost and maintainability benefits. The lightweight turboprop, in combination with new lightweight Xevlar propellers (bIade weight 60 percent of aluminum), will also cause twin-engine preferences to evolve toward the safer centerline thrust configurations such as the Rutan Defiant. These designs will be structurally more attractive (lower wing root bending moments) than thelr piston counterparts. Also, weather radars that can Iook through propellers or be wing- or tail-mounted are already providing a formerIy unavailable capability for push-puiI configurations, which is adding to their attractiveness.

Offsetting the merits of turboprops is the expected lower cabin noise |evei attainable with rear-mounted turbofans. The exact value ot a q,ieter cabin to the business executive is hard to assess against the probable higher fneI con:;ump- tion of the turbofan, but it might be an overwhelming influence. Our hypothetical executive can ill af[ord the fatigue induced by high noise levels, or the throat fatigue and hoarseness which result from conversing in a noisy atmosphere.

TURBOPROP ENGINE DESIGN CRITERIA The postulated turboprop engine is a 73-kg (160-Ibm) _, fixed-shaft unit c,tp_ble of filling the 134 to 231 kW (180 to 310 hp) market niche. It wouid be flat- rated, even for 231-kW (310-hp) applications, to enable it to compete with turbocharged piston engines. After considering weight and cooling drag advan- tapes, the installed fuel efficiency must be competitive with the piston eniine at nominal cruise conditions and excel over the piston at nigh aIt_tudes. Life cycle cost must be competitive w_th turbocharged piston engines, and first cost { must be lowered through manufacturin_ economies.

A proposed approach for re(luring manufacturing c,;sts an..'! providi_,g i:)'._ m,_i:_tenanc_, and high saleability _.nvc. lves 4 fixed-shaft, !,',w-speed. low-strauss design con- cept. This concept ut_', izes a l.'_w-speed, :nultistage axial comurefqsor w;th .1 ?

*Less starter-generator NASA CR-159603 _C Report No. 78-113-15 design tip speed of 259 m/s (850 ft/s) followed by a centrifugal compressor. fhe low-speed feature, in which all elements ran subsonically, prodlices excellent efficiency and Iow noise levels. ]'he high hub/tip ratio and the many blades put the frequencies in an easily handled category.

The compressor rotor consists of an axial component coupled directly to a low- pressure ratio centifugal compressor having all-radial element_ A two-stage inducer is a part of the axial component. The compressor is to_l,wed by a vecv lightly-loaded burner with ample volume to incorporate emissions-reducing _ont,-, vances. The fuel injection system _s a shaft-mounted nozzle system that prov_de_.

the advantages of very high injection pressures _tth low-cost fuel control ,o_- cepts. The turbine is a lightly-loaded, four-stage anxt conceived as a companl,_, for the low-speed, low-stress compressor rotor.

TL_BOFAB FJK;INE DESIGN CRITERIA The postulate_ turbofan engl-ne is an 84-kg (185-Ibm), two-spo_! unit capable of producing approxim3tely 4448 N (I,000 lbf) of thrust under static, sea level.

standard day conditions. It would use the low-cost turboprop gas generator as a major core component. The bypass ratio would be _n the 4 to 5 range and the overall pressure ratio over 20. The engine would be optimized for a 9144 m (30,000 ft), Much 0.6 mission and would fill the powerplant need of a s_x-place business jet which can be_ flown safely and inexpensively by a non-professional pilot.

The turbofan engine design employs a fan having a maximum tip speed of approxi- mately 305 m/s (1,000 ft/s), which produces pressure ratios to 1.4 under stand._,,l conditions. The fan is attached to and foIIowed by a three-stage intermedtal.- pressure (IP) compressor of about a 28°K (50°F) temperature rise per stage, rh_s modest temperature rise enables the manufacture of the IP rotor in accerdance with iow-cost construction concepts. The IP compressor and fan are driven t,v a four-stage, low-speed turbine. The engine accessories are arranged around toe waist formed by the axiai compressor rotor of the core.

TURBOq_FT ENGINE The postulated turboshaft emgine is a free-turbine design with the free turb,,e driving a simple 6,000 rpm output gearset and a high-speed accessory drive. The weight would be about 77 kg (170 Ibm) and it would produce about 373 kW (500 shp) and be flat rated to about 243g m (8,000 ft).

POTENTIAL IMPACT OF POSTULATED ENGINES As the result of the marketing study and airframer contacts, a forecast show_ng the possible influence of powerplant technology advances oa _ircraft productt,_n by type of aircraft was made. This projection, Table VI_, is considerably different from the FAA baseline scenario, (Table II), which projects a more _eisurely pace in powerplant technotogy advances. The assumptions used t, develop Table VII are as follows: 2O

I

NASA CR-159603 WRC Report No. 78-113-15

I

TABLE VI)I. GENERAL AVIATION AIRCRAFT PRODUCTION IN T}_ I_ITED STATES*

I

Rotorcraft Fixed Wing / Piston Turbofan Piston Turboshaft Turboprop

I

Twin- Twin- Single- Single- Twin- Single- Year engine engine engine engine Total engine engine =, ,,

l

1977 413 2,230 12,929 195 228 456 16,451 2,233 12,944 195 219 437 16,442 2,123 12,306 185 228 "56 15,691

I 1980

12,816 1 409 2,211 193 254 5O8 16,392 1981 4 2,411 13,756 414 195 288 576 17,644 1982 30 2,154 12,587 419 2 278 555

197 16,222

176 447 • 266 " r i 2,090 12,327 13 211 16,O62

i 2,207 743 5-58

12,997 56 263 574 _7 17,685 ,+ 11,837 859 I71 4O4 1,873 2,276 312 623 18,355 19_ 9,292 661 1 ,I(F) 5,065 1,405 381 182 822 18,977 1987 369 6,356 8,186 616 949 28 2,016 1 ,047 19,597 5,041 9,606 2,294 723 0 1,080 1 ,150 19,894 : f" P m i *Assumes eo_petitiveI_-priced, fu - efficient, T/P, T/F, and T/S engines will emerge in mid-1980's. Projections will not materialize without vigorous GATE funding.

.

Competitively-priced, fuel efficient, turboprop, turboshaft and turbofan engines will emerge starting in the mid-1980's as a result of the NASA GATE program and related manufacturer activities.

.

The FAA forecast of total general aviation aircraft and rotor('raft production through 1988 is correct (Table lI). This forecast is related to the ability of the airport airways system to safely assimilate additional aircraft. It is an FAA tool for long-range planning and for funding acquisition for system upgrading. It tends toward self- fulfillment.

Twin-engine, fixed-wing aircraft will continue to constitute only about 18 percent of total fixed-wing aircraft production despite an increasing demand for system redundancy. Economic factors, particularly fuel costs, will exert a constraining influence.

Single-engine, fixed-gear airplanes having less than 134 kW (180 hp), for the most part, will continue to be piston-powered. This is because engine manufacturer interest will focus on turbine engine replacements NASA CR-159603 WRC Report No. 78-113-15 for piston engines in the 13/. to 231 kW ([80 to 310 hp) range. Siugte- engine airplane- replacement engines will be predominant [y turboprop.

Turboprops will be preferred over turbofans for th,? following reasons.* Improved takeoff acceleration needed for operations from the many general aviation a,,_,orts having short runways.

b.

The vezy short landing capability provided by propeller thrust reversal.

The weight and balance advantages gained from a nose-mounted engine.

d. Better fuel economy than 3 turbofan.

e.

The added airplaae control|ability and responsiveness permitted by slipstreat inflmces on the tail control surfaces. Thxs ,an be pa_icularlyuse£ut tn salva_bad laudi_gs.

S_ Thirty-two point eight percent of single engine airplanes w_l! have less than 134kW (180 hp). (1916 figure ) .

ha estimted seven percent of the single engine turbine airplanes wxll be tatbotmR-power_. TheSe wilt fill the high-performance market niche.

.

An estimated thirty-two percent of the twin turbine alrplanes will be turbofan-powered. The greater fuel efficlency of the turboprop will be offset by the quieter cabin and greater ease and safety with which a twin-turbofan-powered airplane can be flown by a nonprofessional pilot, particularly under engine-out situations. The single-pilot-flown, small, twin-turbofan airplane will prove very popuJar as a business tool.

o By 1988, all newly manufactured, twin-engine a_rplanes will be turbine- powered.

.

Two-thirds of the helicopters manufactured during the 1977 through 1q8% time period will be turboshaft-powered. By 1988, all newly manufactured helicopters will be turboshaft-powered.

*Customer preference might radically affect the turbofan/turbop,op balance if the turbofan l_ves up to its expectations in minimization ot ,=n_n noise. Th_s factor has not yet been adequately assessed.

( NASA CR-159603 r, CRC Report No. 78-113-15 Table VIII forecasts the annual production of general aviation piston, turboprop, turbofan, and turboshaft engines thr_mgh [988. Quantities are shown for engines installed in new aircraft, and for the estimated total production for fulfitiing new aircraft, replacement, arid retrofit market needs. This table is sdpptemcnted by an assumptions list, and it has been developed by using the aircraft produc- tion figures in Table VII as a basis. The assumptions used to develop "Fable VIII are as follows: k. Engines For New U.S.A.-Produced Aircraft o Piston engine production for new fixed-wing, piston-powered airplanes equals the sum of single-engine airplane production plus two times twin-engine airplane production. Turboprop and turbofan engine pro- duction was figured similarly.

.

T/S engine production for mew rotorcraft was computed by multiplying

li

tile total civil T/S-powered helicopter production by the factor 1.258, where 1.258 equals the s_m of total civil T/S helicopter production :• 7 ¸ plus total civil twin-engine T/S helicopter production divided by total i civil T/S helicopter Froduction for the year 1975. (ref 10).

.

Piston engine production for new rotorcraft was assumed equal to the civil piston-powered helicopter production, since virtually all civil piston-powered helicopters produced are single-engine models that fall in the general aviation category.

B. Total Engine Production for New U.S.A.-Produced Air_lanes and for the Replacement and Retrofit Markets 1 .

Total piston engine production for fixed-wing airplanes for the years 1977 through 1985 equals 1.3 times piston engine production for ne_ fixed-wing airplanes. Total piston engine production for fixed-wing airplanes for the years 1986, 1987, and 1988 equals the piston eng,ne production for new fixed-wing airplanes plus 5,085. The 5,085 equals the annual average of forecasted piston engine production for replace- ment purposes for the nine years starting in 1977. Relatively constant replacement market is projected due to a nearly static piston fleet growth, gradually increasing piston TBOs, and static or declining piston airplane utilization rates as the result of escalating fuel costs and increasing simulator use.

.

Total T/P engine production equals 2.6 times the twin T/P airplane production plus 1.1 times the single T/P airplane production. These figures reflect the need for replacement engines for the existing T/P fleet and expected higher utilization rates of twin T/P airplanes.

.

Total T/F engine production equals 2.6 times the twin T/F airplane production plus 1.2 times the single T/F airplane production. These i • NAS& CR-159603 WRC Report No. 78-113-15 TABLE VIII. PROJECTED GENERAL AVIATION AIRCRAFT ENGINE PRODUCTIO_ _'_ A. Engines for New Aircraft Rotorcraft FixedWing Year Piston Turbofan Piston Turboshaft Turboprop 1977 228 574 390 17,389 1978 828 219 550 390 17,410 1979 786 16,552 228 574 370 1980 819 386 17,238 254 639 1981 832 390 18,578 288 725 1982 868 278 698 396 16,895 1983 1,070 266 669 435 16,507 17,411 1,859 287 722 582 t985 3,994 312 784 979 15,583 11,630 7,875 1,703 182 1,034 7,124 12,218 2,514 28 1,317 5,041 14,194 2,883 0 1,447 i ii Total Engine Production Rotorcraft Fixed Win_

farbdflh

Year Piston Piston Turboshaft Turboprop 22,606 507 296 746 1,074 22,633 1,076 5O7 285 715 1979 481 296 746 21,518 1,022 502 330 831 22,409 1,065 507 374 942 24,151 1,081 21,964 515 361 907 1,122 1983 564 21,459 1,356 346 870 22,634 2,268 373 939 20,258 &,737 I, 406 _,019 16,715 9,224 2, 261 1,344 12,209 207 14,246 3, 107 1,712 1988 676 10,126 16,531 3, 79 1,881 *Projections will not materialize without vigorous GATE funding.

NASA CR-159603 _4RC Report No. 78-113-15 figures re_lect the need for replacement engines for tne existing T/F fleet, an expected earl ier ir_troduction of the GATE F/F in twin engine aircraft than in singles, and a higher expected average twin utilization rate.

Total piston e,_gine production for helicopters for' the years 1977 through 1985 equals !.3 times the piston engine production for new helicopters. Total piston engine production for helicopters for the years 1986, 1987, and 1988 equals the piston engine production for new helicopters plus 79. The 79 equals the annual average of forecasted piston engine production for replacement purposes I _ the nine years starting in 1977. A nearly constant replacement market is projected due to an expected decline in piston helicopter fleet growth and gradual increases in piston TBOs.

5_ Total TIS engine production equals 1.3 times the T/S engiae production for new helicopters.

potential sales of turboprop engines in the 134- to 23I-k_ (180- to 310- eshp) class for the year 1988 are shown in Table IX, together with a list of asst_ptions. The reality of this number of turboprop unit sales can be influenced decidedly by the course of the GATE program and by the success achieved in overcoming turbine engine cost/fuel economy problems.

TABLE IX. POTENTIAL SALES OF TURBOPROP ENGINES IN THE 134 to 231 kW (180 to 310 eshp) SIZE CLASS 1N TREYEAR 1988" Number of Units for New Aircraft = 11,852 Total Unit Sales = 13,803 Assumptions: 1. A competitively-priced, fuel efficient turboprop will emerge in the mid-1980's which will assume the propulsion task currently being performed by 134 to 231 kW (180 to 310 hp) piston engines.

2. In 1988, 83.5 percent of the turboprop engine production wit1 be for airplanes requiring 134 to 231 kW (180 to 310 eshp) powerplants (1976 figure for OE_ powerplant deliveries involving propu!sors that drive propellers).

3. The 134 to 231 kW (180 to 310 eshp) propulsor requirement can be satisfied by three or four variants of a single, basic turboprop engine design. The models _ill be flat-rated to permit operation over a wider altitude spectrum than nonturbocharged p_ston counterparts.

*Projections will not materialize without vigorous GATE funding.

NASA CR-159603 WRC Report No. 78-113-15 Figure 4 shows the potential influence of the GATE program on the 1988 engine sales mix by type for new-pr.,duction, fixed-wing aircraft. Witho,Jt a significant technology advance, the FAA projectton wm:ld appear to be the most plausible, b_lt through the nurt,,ring of specific techno!ogy elements, a trend toward the protec- tion at the bottom of the figure is likely. The rate • t progress in this direction depends a great deal on resource allocations.

To develop turbine engtne cost goals in 1977 dollars, tne ilst prices oi sever.t!

Lycoming piston engines were esttmated using a 1975 price list and specific post- July 1977 list price data for the Lycoming 0-360 A1D four-cyltnder engine and the 0-540 A1A5 six-cylinder engine Four- and six-cylinder engxne list prices were estimated by multiplying 1975 prices by appropriate ratios as determined for the specific engines. The 1977 price list generated by this approach is shown in Table X.

TABLE X. APPROXII%%TE LIST PRICES FOR NEW AND _ACTIJRED LYCOHING PISTON ENGINES - 1977 DOLLARS New New Approx Rmfg kW (hp) Exchange Exchange Engine Outright 3,896 86(115) 5,188 4,458 0-235C, Cl 112/119(150/160) 4,446 O-320A, B 5,912 5,084 0-360A Series 134(180) 5,782 5,047 6,723 134(180) 6,692 5,751 5,032 O-360A4G, A4J, A4K 0-540A Series 186(250) 9,636 8,284 7,224 0-540B Series 175(235) 9,414 8,092 7,061 0-540E 194(260) 9,434 8,112 7,071 I0-320 BIA I19(IO0) 6,332 8,427 7,244 I0-320 CIA 119(160) 6,897 9,143 7,896 I0-320 E Series 112(150) 7,806 6,716 5,864 LI0-320 BJA 119(160) 8,689 7,470 6,527 LIO-320 CIA 119(160) 9,571 8,229 7,195 I0-360 AIA 149(200) 7,696 6,728 8,945 I0-360 B4A 0,255 134(180) 8,329 7,12q I0-360 CIC 149(200) 8,781 7,552 6,594 LI0-360 CIE6 149(200) 9,828 8,450 7,380 216(290) 12,988 11,167 9,7q7 IO-540 A_ E Series 110-540 C4B5 186(250) 10,908 9,378 8,17_ 10-540 G Series 216(290) 13,525 11,630 10,140 G0-435 C2,C2B,C2C 194(260) 18,710 16,090 14,029 G0-480 D Series 220(295) 19,765 16,997 t4,823 G0480 G Series 220(295) 19,415 14,55_ 16,696 GS0-480 B Series 254(340) 25,049 21,284 17,52_ IG0-540 261/283(350/380) 22,850 19,416 15,q90 TI0-540 A2C 231(310) 15.i13 20,154 17,332 283(380) 19,445 IGSO-540A, AIA 27,784 23,615 NASA CR-150603 I_C Report No. 78-i13-15

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T UR BOJ ET/TURBO FAN O/

I 2. I0 o

37"; EINGS

I TURBOPROP

1976 CY ,_. o_% 716 ENGS ACTUAL

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15. 105 AIRPLANES

)

L

s I%

i 196 ENGS TURBOPROP 1988 FY

1o. 8%

2542 ENGS FAA PROJECTION 18, 744 AIRPLANES PISTON / O/

\ 8_.1 o /

'_ PISTON 1988 FY

_ __._._0

PROJECTION SHOWING POTENTIAL INFLUENCE _ 504 ! ENGS OF NEW TECHNOLOGY

\

[ TURBOPROP 18. 744 AIRPLANES I 64"20/0 _- _ TURBOFAN \ 14, 194 ENGS _ --/ 13.0o/o A '.c; i 2883 ENGS Projected Influence of New Technology on OEM Engine Saies by Tvp_ Figure 4.

for Fixed-lqing Aircraft (USA) NASA CR-159603 _¢RC Report No. 78-113-]_5 Heaning/ul piston engine, turboprop and turbofan cost comparisons cannot be made without the additio[l of propeller prices to the piston and turboprop engine prices. Consequently, propeller list price data was obtained from H_,rtzell Propeller, Inc. Typical fiFures that exclude anti-ice and deice provisions are shown in Table XI. For cost comparison purposes, the cost to provide for turbotan cowl anti-icing can be ass_uned to approximate propel let ant i- icing equipment costs.

Typically, engines and propellers are sold to original equipment manutactu,ors at 60 percent of list price. The data provided in Table X, therefore, can ne translated into meaningful, albeit somewhat oversimplified, co_t goals for turboprop engine family. The combined data of Tables X and XI have to be used judiciously to estabiisb turbofan engine cost goals, however, because of the higher characteristic cruise-thrust/takeoff-thrust ratio ot turbofans relative to turboprop and piston propulsors.

TABLE XI GENERAL AVIATION PROPELLER DATA _ (SI Units)

f

List Prices Engine Dia Range Weights Type Propeller kW Range (cm) (kg) (S) 134-194 I83-249 27 7-34.9 Two-Blade, Constant-Speed I,I60-1,595 119-186 i83-208 28.1-34.0 Two-Blade, Feathering 1,380-1,520 213-261 208-249 39.9-54.4 1,685-2,3o0 Three-Blade, Constant-Speed 194-336 188-244 37.6-56.7 Three-Blade, Feathering 2,025-2,755 236-67l 2!3-274 53.5-04 9 Three-Blade, Reversible, 3,030-3,0_5 Feathering 507-533 229-254 69.9 Four-Blade, Reversible, 4,695-5,005 Feathering 835 98.9 Five-Blade, Constant-Speed, 5,805 Reversible, Feathering Anti-icing or Deicing Provisions Optional at added cost *Source: Hartzell Propeller, Inc. Piqua, Ohio, 1977.

NO]E: OEM price approximates O0 percent of list.

[

NASA CR-159603 WRC Report No. 78-113-15 TABLE Xl. GENERAL AVIATION PROPELLER DATA_ (English) List Prices Engine Dia Range We ] _hts hp Range (in) (S) Type PropeLler 180-260 72-98 6!-77 1,[60-1,595 Two-Blade, Constant-Speed 160-250 72-82 62-75 1,380-1,520 Two-Blade, Feathering 285-350 82-98 88-120 Three-Blade, Constant-Speed 1,685-2,360 260-450 74-96 83-125 2,025-2,755 Three-Blade, Feathering 317-900 84-108 <118-143 Three-Blade, Reversible, 3,030-3,965 Feathering 680-715 90-I00 154 Four-Blade, Reversible, 4,695-5,005 Feathering 1120 Ill 218 5,805 Five-Blade, Constant-Speed, Reversible, Feathering P i i ,ml i ii i i - Anti-icing or Deicing Provisions Optional at added cost *Source: Hartzell Propeller, Inc. - Piqua, Ohio, 1977.

NOTE: OEM price approximates 60 percent of list.

HISSION PROFILE CONSIDERATIONS There are several considerations meriting attention with respect to the selection of mission profiles for general aviation aircraft in 1988. These involve the expected fleet size, the national airspace system, and the level of technology.

The general aviation fleet size, for instance, is projected by FAA to increase from about 181,600 active aircraft at the end of 1976 to about 267,000 in 1988 (ref. 3). This 47 percent increase is expected to be accompanied by an 89 percent increase in hours flown and an 88 percent increase in instrument opera- tions. The 1988 airspace will therefore be more congested than it is today and the traffic separation problem will be more acute.

In the instrument environment, traffic is separated longitudinally, vertically, and laterally. Participating pilots flying unpressurized airplanes are frequently limited in their selection of suitable altitudes because of weather and the capability of their aircraft [90 percent of the earth's weather occurs below 3048 m (10,000 ft)]. Altitude options may be excluded, for example, because of the likelihood of airframe icing in specific strata. Witt_ increasing traffic, -'9 NASA CR-159603 WRC Report No. 78-113-15 fewer pilot altitude requests can be granted on a timely basis because of con- flicts with other traffic. A|so, alternate, less safe, less comfortable altltudes will sometimes be assigned by air traffic control. The reduced avallabiiity _,_ preferred altitudes in the lower strata in i988 will therefore cause airplane buyer interest to shift toward airplanes that can be operated efficiently over a wider spectrum of altitudes. These airplanes will have pressurized cabins and cruise speeds sufficiently high to minimize the effects of headwinds at h_gh altitude. The popular air_iane of the era will also have been certified tot flight in known icing conditions because icip_ is a common occurrence during cloud penetrations at temperatures below 273°K (32 F).

The 89 percent increase in expected hours to be flown tn 1988 will cause more congestion of traffic flying under visual flight rules, especially at the lower altitudes, since all aircraft start and conclude their operations at ground level. A plot of traffic density versus altitude would show the greatest density at traffic pattern altitudes with progressively lower densities at the higher altitudes. In recognition of traffic density trends, GATES Learjet re(entl_ certified its Century III models (24E, 24F, 25D, and 25F) for operations at altitudes u V to 15,545 m (51,000 ft). These aircraft will literally have thf sky to themselves above about 13,716 m (45,000 ft). In this same spirit, g,-neral aviation aircraft manufacturers will be building a greater percentage of airFianes capable of routine operations in positive controlled airspace where separation from other axrcraft is assured [currently 5486 m (18,000 ft) and above].

General aviation has had a long-standing interest in emulating the scheduled airlines, but there have been obstacles. Besides cost, these have had to do with the size of general aviation airplanes in relation to the size and weight of available instrumentation and avionics. In recent years, the larger corpo_,te jets and turboprops have mastered the emulation goal and the mastery is extend_n_ to downsized aircraft. The technology-related pace has been acceIerated t_y turbine engine, avionic, and a,topilot developments, and additional progress s foreseen for the decade ahead. The key technoiogy elements wilt involve small turbine engines in combination with digital, integrated avionics systems. Scme of the market applications and corresponding mission profiles Identified on the pages that follow reflect expected progress in these areas, both from a technology and a cost standpoint.

CANDIDATE AIRPLANES AND MISSION PROFILES FOR TRADE STL_IES The established major airframe manufacturers are a very conservative iot, and this is for a good reason. According t.o James N. Lew, senior vice president (now retired) for engineering of Beech Aircraft, development costs, inclt_dlng production tooling, of a proposed new airplane may range from 54,500 to S6,000 per pound of airframe (1976 dollars), where nonpressurized airframes weigh about 66 percent the empty weight and pressurized airframes average 73 percent o! _he empty weight of the airplane. Clearly, a large investment is involved, and the manufacturer must be very sure that he has se!ected the proper engine for the new design, or at least that there is a powerplant alternative should the selected engine prove unsatisfactory. He must also conslder product !iabititv wheo choosing a new engine.

3O i- NASA CR-159603 78-113-15 WRC Report No.

TABLE XII. PERFORMANCE CAPABILITIES OF PISTON AIRCRAFT BEFORE RETROFIT 1 (SI Units) A i_ PI,AN E AEROSTAR t,81P >ItIONEY fOl CIRJGAN ¢) PeFsolts 01i ['_oard 4 4 { _£6 2722 Gross Wetght, kg 1243 YES Pressur t zed N(J NO iLq Engine Rated kW, each l,_q Takeoff Distm(e (SL, ST[)Day, dW) " St]oft N_$rma-] it++> Ground Run, i:; 271 284 h ,_,q Over 15.2 m Obstacle, m 4_ 540 564 ' " 312 _b6 "_4ql Rate of Climb (SL, STD Day. GW) m/mt,n 'rime to Climb to Indicated 30 to 4572 m 21.O to 7620 m Altitude, - Min 25.0 _o &572 m Maximum Cruise Speed km/h 324 5578 8534 Serv/ce Ceilin_n m .5700 Range (45 Min Reserve) 2591 4572 7620 A[titt_le. m 2438 859 1109 12t7 B_nge, km 995 _28 t.41 Speed, km/h 300 _a_x Fuel with Full Seats & Bags 2, _50 g50 I 170 3.23 +_7g kmll 7.72 17.t, t9.4 22.1 Seat -km/I 30.9 .., Landing Oistanc_ (SI., STD Day, GWJ Maximtlm Performaneg Over 15.2 m Obsta(le, m 491 F,I:) Z7/* Grot_ Roll, m 235 3U2 |from Atrcraft Handbook Data 2Assumes qO.; kg f.r eat'h perso, on board and their baggage TABLE XlI. PERFORMANCE CAPABILITIES OF PISTON AIRCRAFT BEFORE RETROFIT 1 (English) A,,,2 RPLANE ,_oI HOo_,EV , C()t GAR AER)51AN ¢_01l > Persoll._ I)ll _++.+l-(+ o Gross Weight , Ibm 2,7_0 3. 800 (_, t)()O Pressur i zed N() NO YES 200 I bO EngH_e Rated hp, each 290 Short Normal Takeoff Distance fSL. F,TD l)ay, GW ) 000 Groun<! R,ln, tt _90 931 Over 50 it t)bst:t;le, ft 1.518 1.7/I i, 800 ' 8t)O "'Rate of Climl') (SL, STD Day, GW), I ,02.3 I, :UO ft/min Time to Climb to lnd)c:lted Altitude, - Hit) 25.0 to 15,000 It _0 to 15,000 f( 21.0 t,) 25,00) It 175 lb8 ttaximtm_'-C'ruise Speed, knots 257 18,700 Service Ceiling, ,_ 18t300 2_,d 001) Range (45 Hin Reserve) 8,5 O0 Altitude, ft 8,000 15,0OO 25,000 4o4 Range, rm_ 537 5qq 057 160 231 "38 Speed, knots 162 Hax Fuel with l".li Seats & BaRS 2, ?5 t lq i lg ga 1 f 45 8.8q f). t_O 7. 72 33.8 Seat-nm/{6al 63.1 _'I_X ItIIl,lJ_ Per+ ormance I.;l, ndi':tg Dist;)n(t" (SL, STI)'D,)y, G_" l,e, O0 . ,, '> Over 50 it obstacle, it t,°lO I ilm/ga I t i r). 7B Ground R(>;I. tt r ;0 )

I

'!00 ",_1 i lFr)m Airtralt ttdll.lb,'ok l)4t., eAssumes 200 it, t,_r (+,,tth W._s,_r) .. h.._td a,.! the. Jr t_.t_.,_<, j I :31 .7 NASA CR-159603 WRC Report No. 78-113-15 TABLE XIll. AVERAGE NI_IBER OF PERSONS TRAVELING IN GENERAL AVIATION AIRCRAFT (1975 Civii Air Patrol Survey) Average Number Of Travelers Crew Aircraft Type (tncluding Crew) Requ irement s 2.1 I Single-engine piston 2.9 I Helicopters 3.8 I Multi-engine piston 5.7 I Turboprop =r 5.4 2 _harbojet/turbofan Because of airframe development cost and product liability considerations, th,- introductory use of a new turboprop engine would probably be. by engine substttu- lion in an existing model. A new turbofan would probably also be Introdu,_,d in an existing airframe if an appropriate airframe became available by the txme of introduction. The re-englned alrplanes would have to provlde performance , cost advantages (or both) over preceding models.

The candidate fixed-shaft turboprop engine was retrofitted to, and evaluated in, three contemporary piston-powered airplanes. Two of the airplanes were twins and one was a single. One of the twins, the Gulfstream American Cougar.

was selected to permit engine merit evaluations at derated powers where the introductory risk could be minimized and in-service experience gained. The other twin, a Piper Aerostar 601P, was chosen for evaluating the merit of _om_- nally-rated engines. The single-engine Mooney 201 was seletted to satisfy the need to substitute the turboprop for a four-cylinder piston engine in lieu of a heavier six-cylinder engine. The weight difference between the turboprop and a six-cylinder piston engine produced airplane balance and stability perturbations with attendant retrofit complications.

_ission goals were to provide, at comparable gross weights, equal or better performance and economy of operation in terms of seat-km/l (seat-nm/gal) than the airplane being retrofitted. An expanded airplane operating envelope, together with a lll2-km (600-nm) stage length capability with all seats occupied, was also desired. Table XII lists the performance capabilities of the aircraft being retrofitted (before retrofit), and Figure 5 depicts these aircraft after retrofit. A candidate twin turbofan-powered configuration is also shown in Figure 5.

The twin turbofan airplane was sized to accommodate the average number of travel- ers shown in Table X[II, i.e., 5.4. Six seats were believed more than adequate, since the requirement for two _:rew members on all business jets has beeT_ relaxed.

and because the 5.4 figure provided by the Civil Air Patrol Survey lacl,,,te,_.

according to a private survey, an average of one traveler having no assot_,,t_;;_

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NASA CR-159603 _RC Report No. 78-113-15

I

with the business at hand. This person was traveling si,lp[y because extra seating was available that wou!d otherwise go unoccupied. Also, with only six passenger seats, a crew of two would be unlikely for most missions. Only one

I

professional crew member would be used, or the businessman would fly the a_rplane himself.

I

The selected mission tor the twin turbofan airplane tnvolved a lll2-km (600-nm) range requirement at a _lach 0.6 cruise at 9144 m (30,000 ft). Adequate fue_ would be needed to travel to the destination 1112 km (600 run) away, hold for 45 minutes and then proceed to an alternate destination. Consequently, a 1852 km

I

(1000 run) range capability with reserves would be desirable.

The Hach 0.6 cruise speed was selected because it fails short of the speed where

I

compressibility influences become significant. Once near the compressibility flight regime, lifting surfaces need to be made thinner and this adversely affects available wing volume for fuel storage, a problem unique to small jets since fuel storage volume is reduced by the scale factor cubed while the thrust

I

requirement reduces by the scale factor squared. Also, the potential for airplane stability and control anomalies begins to influence control system design. To avoid adverse influences from local shockwaves on the lifting surfaces, added control system complexity and cost can be anticipated (e.g. from a Mach trim device that includes actuator, computer, air data sensor, and aural _ach overspeed warning components). The added cost and complexity of Mach trim, a yaw damper, etc., are not believed warranted in small jets that will be flown by nonprofess- ional pilots.

NASA CR-159603

WRC Report No. 78-113-15

MOONEY 201 TURBOPROP AEROSTAR 601P TURBOPROP i

E ]

TWIN TURBOFAN STUDY AIRPLANE GULFSTREAM AMERICAN COUGAR TURBOPROP A I0, 0_t0 Figure 5. GATE Study Air,,lanes

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NASA CR-159603 _C Report No. 78-113-15

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SECTION 5 BROAD SCOPE TRADE-OFF STUDIES

!

Turboprop and turbofan parametric studies for projected 1988 state-of-the-art general aviation turbine engines were conducted using as a basis, data generated during the market analysis. SpecificalLy, the market analysis shows a need for a low-cost, flat-rated turboprop capable of replacing four- and six-cylinder, hori- zontally-opposed, turbocharged piston engines producing from 134 kW to more than 224 kW (180 hp to more than 300 hp). A significant market for a lo_cost, iuel- efficient turbofan in the 4448 N (l,O00 Ib) thrust class was also foreseen. Both the turboprop and turbofan, when installed in appropriate aircraft, would have to provide performance, fuel economy, and life cycle cost (LCC) benefits comparable to piston airplane counterparts. Attendant safety, utility, and environmental improve- ments are additional prereqaisites.

trade st_ies described in this section provide exiles of work that was clone ?

to define candidate turboprop (T/P), turboshaft (T/S), and turbofan (T/F) engine v- • concepts and layouts for consideration relative to common core compatibility.

TURBOPROP/TURBOSBAFT Par_t ri¢ Stady Parametric performance data for shaft engine cycles was calculated using compressor and turbine efficiency, compressor temperature rise, and turbine inlet temperature (TIT) as variables. The data was prepared for flight speeds and altitudes which were determined representative of operational conditions expected for the T/P airplanes defined in the marketing study, i.e., SL/Mach 0, 4572 m (15,000 ft)/Mach 0.3, and 7620 m (25,000 ft)/Mach 0.3.

Samples of the parametric performance data generated during the T/P design point study are shown in Figures 6 and 7. The curves show relationships between specific fuel consumption (SFC) and specific power at seven compressor temperature-rise values between 167°E and 500°K (300°F and 900°F) and five turbine inlet tempera- tures between 1144°K and 1589°K (1600°F and 2400°F). To illustrate the effect of component efficiency on specific performance, additional plots were prepared for compressor and turbine efficiency values incremented by four percentage points in plus and minus directions from nominal values (q_a_eO.78, qt : 0.86), Figures 8 and 9. These curves, for a 1367°K (2000°F) TIT , illustrate the import,:nce of aerodynamic component development for SFC and specific performance gains. A four percentage point improvement in compressor and turbine efficiency (Figure 8), for example, can produce an SFC reduction of 18 percent. Potential improvements of this order emphasize the necessity for working to maximize efficiencies for vi- ability in a low-cost T/P propulsion unit.

Turbopro p Design__Po_i_gt "hoice Rationale The choice of the optimum design point of a turbine engine is a compromise inw_lx;- ing mmerous variables. Some, such as performance, can i)e th(_roughly ,tuantifiec! at the early stages of design; and others, such as costs and mechanical refinements, are less easily quantified. The process of cycle selection involves, in general, a combination of design point analysis focused around a preliminary design concept NASA CR-159603 WRC Report No. 78-113-15 A J= (o. 9)- C.£) MPONENT DEFINITION S (LPP_,_,L,= 0 J= -0.50 (o. 8_- d U.

I (o. 7)- Z -0.40 p.

el (0. 6)" Z .1 L (o. S, .o.3o _0.4) (320_ (o) SPECIFIC POWER -- ESI, IP/WI, /w-=tlg (l_pstl_) Turboprop Parametric Design Point Study - SllMach 0, STD Day, Figure 6.

q = 0.78, _, = 0.82 C t • 0.50 (0. S_" (0.7)" -0.40 (0.6).

(o. 5).o._o (o. 4_ - '0, 20 l O0 200 300 400 500 (0.3) I I , I, l _ I , I (0_ (4'03 (8'0 _, (1203 ( 160"_ (200/ (2403 (2803 (3201 SPECIFIC POWEF_- ESHP/WI, kW's/kg (hp.s/Ibm) A--I0838 Turboprop Parametric Design Point Study - SL/Mach O, 5"173 Day, Figure 7.

= 0 78. = 0.86 qc " Ct < / • ,• _•, .... :4, NASA CR-159603 78-113-15 k_C Report No.

I

!"COMPONENT DEFINITION : -0.50 =: (LP. P_,Nt _.r: 0 - (0.8)- (£;P'P_.L,.,_FR = 0. 035 E LEAKAGE _ I.0_>_, v

I (£_p, pX.r__oz =0.02

.C _/- = 0. 995 LHV 42, 800kJ/k 9 _, (0.7_- _|8, d0OBTUllbml .I 9"-., : 0. 985 ,_T -0.40

I 6

V_ _ 152m/s b.

7 °K '300°R) =_500 ftl$_ m bJ (0. 6_- Z kx ,

l

o_ F- D.

_. _, \ '_, _ 2zz(400) _E (o. s)- o. 3o m z J (o. 4)-

L

I, U '0. 20 W D. tOO 200 300 4OO 500 (0.3) U} I ,I , I , • I , I , (o') (4'0'_ (8'0) (|20) (160} (200"_ (Z40) (280% (320'_ SPF_c4FIC POWER -- ESHP/WI. kW-s/kg (hp-$11bm) A--IO82B Figure 8. Turboprop Parmetric Design Point Study - SL/Mach O, SID Day, TIT = 1366°K (2000°F) A _- (0. ;)- E .0 COMPONENT DEFINITION -0.40 .¢ (_R"P_BL._,,_,, : O. 035 LEAKAGE :: 1.0% (_P/P'_-_oz :: 0 02 _1, = O. 995 d , (_ P/P_ ..... = 0 h (.%T, LHV = 42, 800kJIkg U) =(18, 400BTU/Ibm) I (0. 5)" -0.30 V_w]= 152m/s Z (400"_ I'- 82 _._,____ 67 °K 300°R] t/c,,, = 0. 9_35 m (0. 4)- Z 333 (600> =(500 ftls) _/' ; 389 i700_ .J • 0. 20 bl _, :0.82; 444 (800") 9, :.0.9(3 SO0 (9O0) " (0.3'_" L_ I O0 200 300 400 500 m (o. 2_ (40] (8e'_ (120 _ , (160 _, {200_ (240h (280_ (320"_ SPECIFIC POWER -- ESHP/WI. kW-s/kg (hp.sllbm) A--I0839 7620 m Figure 9. Turboprop Parametric Design Point Study - (25, {)r_O ft) / lath 0.3, STD Day, TIT = 1366°K (2000°F) _7 NASA Ck-159603 WRC Report No. 78-113-15 together with a great deal of judgment in interpreting the impact of cycle param- eters. In the case of GATE powerplants,previous studies of turbine engines utilLz- ing low cost manufacturing techniques based on low rotational speeds indlcaLed % that substantial cost benefits can be accrued if reasonable performance (:an be obtained from aerodynamically simplified components These components will, at course, have to be compatible w_lh mechanical arrangements having good dynaml_ characteristics.

The performance characteristics of a simplified-geometry axial compressor were quantified and shown as a function of pressure ratio, as indicated in Figure 10 Notice that these data indicate that at pressure ratios over 3.9:1 some form ol stability control device might be necessary. Test performance of a sample com- pressor of this type of construction indicated that such a requxrement might incited be real The taking of bleed air for cabin pressurization from the axial compres- sor might obviate the need for a separate, engine-mounted stability control device, however.

The performance of a low cost, geometry-limited, centrifugai compressor which would be co_aatihle with a "low cost axial" was quantified as a function of its pressure ratio and the pressure ratio of the leading compressor element at constant absolute flow. This information, shown in Figure 11, indicates that a severe performan¢e penalty might occur if the axial compressor element produces a pressure ratio exceeding 4:1. Combined compressor performance was then estimated and is shown in Figure 12. These data indicate that lead compressor pressure ratios between _:1 and 4:1 do not significantly change overall compressor performance, but a lead compressor element having a 5:1 pressure ratio will significantty degrade component performance.

To utilize this information in assessing the merits of a compressor, a preliminary estimate of the probable cruise turbine inlet temperature range must be made Cruise turbine iniet temperatures of 1200°K (1700°F) and 1255°K (1800°F) wore chosen because it is believed that a low cost turbine design which has desirable life characteristics and utilizes semi-noncritical materials will be limited to the 1255°K (1800°F) cruise temperature leveI.

With the decision to use these turbine temperatures, stage turbine efficiencies of 86 percent for a four stage turbine were estimated and trade studies of specific fuel consumption versus pressure ratio were made. The resulting parametric per- formance, evaIuated at 74 percent compressor efficiency and the selected turbine inlet temperatures, is shown in Figure 13 for the 7620 m (25,000 ft), 0.3 _ach number turboprop airplane flight condition. Superimposed on these data is the performance of the engine concept with compressor efficiency adjusted for overall pressure ratio as shown in the 3:1 and #:1 lead compressor configuration data of Figure 12. This information shows that specific fuel consumption will be a minimum at a 15:I pressure ratio at this condition. However, with a penalty of less than two percent in specific fuel consumption, a pressure rat_o as low as 11:1 can be used, particularly if cost benefits are to be obtained. From this trade study, it was concluded that a compressor of approximately 10:l pressure ratio at sea level static conditions is optimum. Significant increases of pressure ratio over !0:1 will not be advantageous as long as the cruise turbine temperatures are held within the 1200°K (1700°F) to 1255°K (1800°F) range because compressor reJated eligint • costs will escalate with increasing pressure ratio.

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NASA CR-159603 WRC Report No. 78-113-15

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Turboprop Performance - Specific Fuel Consumption vs Compressor Pressure Ratio at 86% Turbine Stage Efficiency 25,000 ft 0.3 Mach Number NASA CR-159603 WRC Report No. 78-]13-15 The low cost fabrication concept pr,'.posed for this engine introduces anomalies into the conventional understanding of tests. This is because of the minor influence of pressure ratio and the number of stages on actual compressor cost. A higher pres- sure ratio compressor requires a larger number of turbine stages, however, to properly expand the gas and this in turn requires a longer shaft system. The length extension introduces cost increases into the shaft and shall suspension tn increments which are functions of the particular mechanical design concept.

Figure 14 shows the effect of the choice of the number of axial stages when com- bined with a low specific speed centrifugal of approximately 167°K (300°F) tempera- ture rise at different temperature rises per axial stage. Expertence has shown that the limitation of axial stage temperature rise to 28°g (50°F) per stage per- mits low cost fabrication methods to be used, and the limitation of the number of stages to six increases the probability of eliminating stability control devices.

Since a six-stage simplified low cost axial is less expensive than an axial of fewer stages at higher pressure ratio and will run at considerably lower rotational speeds and stresses, it was decided that the turboprop compressor design would be a six-stage simplified axial of approximately 167°K (300°F) temperature rise followed by a low specific speed, dose-coupled, centrifugal compressor of slightly less than 167°R (300°F) temperature rise.

Turboprop P7757 After completing the T/P parametric analyses, conceptual layouts were made of candidate T/P engines, and engine performance, weight, and cost trade-offs were made. The surviving concept was configured for good performance and minimum devel- opment and procurement costs. It uses an axial/centrifugal type compressor with the axial component based on low-cost design and manufacturing concepts developed for the _33 low-cost turbojet. The turbine concept is based on manufacturing techniques proprietary to Williams Research Corporation 0aRC} for low-cost rotor and stator construction. These techniques result in manufacturing costs which characteristicall_ are relatively independent of the number of stages in the compo- nent. For satisfactory results, a low stress level design or low specific speed component is required that uses low-speed aerodynamics.

The goal of a long operating life led to a time between overhaul (TBO) design objective of matching a,rframe life (arbitrarily asst_ed to be 10,000 hours). It the objective is achieved, a user of the turboprop would no longer have to set aside a reserve for overhaul or engine exchange allowance (typically $5 to $10 per flight-hour for piston engines) in his direct operating cost accounting. The impact on airplane alid engine LCC would l,e remarkable.

The engine size selected would allow for a flat rating of approximately 22/, kW to about 6069 m (300 hp to about 20,000 ft) altitude and the specific performance level would be better than competitive piston engines when proper allowances are made for the lighter weight [typically a 45 to tgl kg (100 to 400 Ibm) advantage], reduced cooling drag (piston engine cooling drag is 5 to 20 percent of the total cruise drag of the airplane) and the lesser frontal area.

The recommended engine configuration is characterized by a six-stage. [o_sp_.ed axial compressor followed by a low specific speed centrifugal compressor which supplies air to a shaft nozzle-fed corl_bustor. A four-stage axial tt_rbine is driven by combustor eff!ux. The core engine is reversed, with the propelle_ drive gearb,'.:, NASA CR-] 59603

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Various Worl<-levc:] St4?_c'.q of ,,3,,_,l,1 ._,_lon NASA CR-159603 WRC Report No. 78-113-15 mounted on the turbine end of the shatt system and the compressor _r_let at. the aft.

end of the engine. 'the exhaust is dis_[;3rged trom two p,_rts on either side of the engine and turned aft to recover ..is m_('h residual thrl_st as possible. The dual exhausts are to balance side fo_(es, thereby precluding exha;Jst contrib_ltions to possible airplane spin recovery problems.

The compressor pressure ratio ts over lO:}, and _[th conservative efticiencies the turboprop could provide an g or 9 percent fuel efficiency advantage over competing piston engines. A 25 percent instal fed fuel efficiency advantage could be obtaine¢l through an aircraft specifically configured to capitalize on the lighter, more compact, turbine power unit and through more extensive engine development. Figure 15 is an installation drawing for the fixed shaft P-7757 turboprop. Estimated component performance is shown in Table XIV. Figures 16, 17 and 18 show some of the engine performance data that were used in conjunction with the NASA-developed General Aviation Synthesis Program (GASP) to predict the performance of Piper, GulfstreamAmerican, and Hooney aircraft retrofitted with the turboprop.

Airplane Studies Three existing airplanes were chosen for study relative to T/P retrofit possi- bilities. T_o of these were twins and one was a single. The twins were a Gulf- stream American Cougar and a Piper Aerostar 601P. The single was a Mooney 201.

Other airplanes could have been selected for study and some would have shown the turboprop in a better light from the fuel efficiency standpoint vis-A-vis the piston than did the Cougar and Mooney. Nevertheless, the selections were made according to the following rationale.

At the time the airplane studies were initiated, only a cursory T/P cost analysis had been made. This analysis indicated that the turboprop could be produced for an OEM price falling in the the $i0,000 to $30,000 range. Because the cost was in- fluenced by production rate and quantity, and _luantity, in turn, by the types of airplane capable of productively using the engine, a twin and a single at the lower end of the retractable gear airplane cost spectrum were selected for analysis. If these could be shown to benefit from a turbine engine retrofit, the potential for" the engine would be very great indeed. Thus, the Mooney and Cougar were selected as representative of the type of airplane that could have a derided influence on the demand for the turboprop and the resultant cost. The more expensive Aerostar was selected as one of the more appropriate airplanes tot retrofit from the stand- point of demonstrating fuel efficiency and LCC advantages.

There were other reasons for selecting the Cougar and Mooney. The four-place Cougar, for example, would be an excellent test bed for engine introduction, be- cause it is a twin and very good performance can be achieved at a conservative introductory engine gw (hp) rating. Service expe'rience could thus be obtaihed wiLh a minimum of risk at an engine life meeting ¢ustumer expectations. Also, the very large Cougar cabin could easily accommodate two addittonal passenger s_'ats, and the potential for gross weight growth is excelier.t. Note, however, that thv Co_g_r is limited with respect to cabin pressurization potential [differential limit about 21 kPa (3 psi)l because of the fuselage shape and strt_ctural makeup.

The single-engine Moor:ey was selected for analysis for sew'ral reasons. First, the _erodyt_amic drag was we]l known because c,f the ve,,-y excellent drag redu<tto,_ pra- gram that preceded Lhe introduction of the Nodet 20[. This facil *t_t_.,t piston airplane performance-matching using GASP and the later turboprop airplane pertor-

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NASA CR-159603 WRC Report No. 78-113-15

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TABLE XIV.

TURBOPROP P7757 COMPONENT PERFORMANCE SUMMARY (SHEET I OF 2) (SI Units)

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Takeoff Cruise Altitude - m

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4572 4572 Flight Velocity - km/h 0 0 519 519 Ambient Temperature - OK 288 288 258 258

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35O0O 35000 Shaft Speed - rpm 35000 35000 Exhaust Gas Temperature - OK 766 824 739 796 0.400 BSFC - kg/kW-h 0.366 0.332 i O.319

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2.79 2.80 r 0.726 0.731 0.726 0.730 _c BURNER: 0.03 0.03 0.03 0.03 APt/P t _B @ hf = 42800 kJ/kg 0.99 0.99 0 99 0.99 F/A 0.016 0.018 0.016 0.019 NASA CR-159603 WRC Report No. 78-I13-15 TABLE XIV. ENGINE P7757 COMPONENT PERFORMANCE SUFL_ARY* (SHEET 2 of 2) (SI Units) Takeoff Cruise FIRST STAGE TURBINE: 0.057 0.057 0.057 0.057 (W_--_t/Pt)in - (kg/s) 4_/kPa 1.60 1.59 1.60 1.59 Pb 0.860 O. 860 O. 858 O. 859 qt i/ I n, i O. 049 O. 049 0.087 0.087 ckg/,) jW/ , 1.66 1.66 1.67 1.67 P r O. 860 O. 859 0.860 o.859 THIRD STAGE TURBINE: O. 136 O. 136 0.136 0.136 (W_'_t/Pt)in - (kg/s) _/kPa 1.74 1.74 1.78 1.78 P r O. 860 O. 859 0.860 0.859 qt FOURTH STAGE TURBINE: 0.224 0.224 0.229 0.229 P 1.82 1.84 2.17 2.17 r O. 86O O. 859 0.855 0.850 qt EXHAUST DUCT : O.O52 O.O54 0.070 0.073 APtlP t .i NOZZLE: A - cm 2 64. 376 64. 376 64.376 64.376 P 1. ii0 1. I19 I.i00 1.130 r O. 985 O. 985 CF 0.985 0.985 *Assumed losses - 1.5% gearbox, i. I19 kW parasitic.

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NASA CR-159603 WRC Report No. 78-113-15

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TABLE XIV.

TURBOPROP P7757 COMPONENT PERFORMANCE ShiVeRY (SHEET 1 OF 2) (English)

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Takeoff Cruise Altitude - ft 0 0 15,000 15,000 0 0 280 280 Flight Velocity - knots

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59 59 5.5 5.5 Ambient Temperature - OF 306.8 375.7 272.5 324.3 Shaft Output Power - hp

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Net Thrust - lb 77.3 80.1 30.3 33.9 202.0 226.1 148.9 170.0 Fuel Flow - Ib/hr Tu__aelnletT---emperature - oF 1,700 " 1,850 1,700 1,850 35,0_O !35,000 35,000 35,000 871 974 _t Gas Temperature - OF 920 1,024 0.658 0.602 0. 546 O. 524 BSFC- Ib/(hp-hr) O. 488 0.471 O. 598 0.555 a_ F_FC - Ib/(h_-kr) mr H , , , INLET DUCT: 0 0 APt/P t AXIAL COMPRESSOR: 5.519 5.402 5.950 5.860 (Ibm/s) _/psia (W_-_t/Pt)in - P 3.815 3.854 4.074 4.133 r 0.810 0.802 0.810 0.810 qc CENTRIFUGAL COMPRESSOR: 1.81 1.77 1.85 1.80 (W_-_t/Pt)in - (Ibm/s) _/psia 2.67 2.68 2.80 P 2.79 r O. 726 0.731 0.726 0.730 qc i J, BURNER: 0.03 0.03 0.03 0.03 APt/P t 0.99 0.99 0.99 0.99 riB @ hf = 18,400 Btu/lb 0.016 0.018 q.o[6 0.019 F/A •i_¸ _;_ _ -_,_,_ _ _ : _, ,_ ......... __. ,. • • NASA CR-159603 78-113-15 WRC Report No.

TABLE XIV. ENGINE P7757 COMPONENT PERFORMANCE SUMMARY* (SKEET 2 OF 2) (English) Cruise "l_keoff FIRST STAGE TURBINE: 1.16 1.16 1.16 1.16 (W_-Tt/Pt)in- (ibmls) 4_Ipsia 1.60 1.59 1.60 1.59 Pb 0.860 O. 859 0.860 0.858

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WRC Report No. 78-113-15

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NASA CR-159603

WRC Report No. 78-I13-15

mance predictions. Secondly, the Hooney structure has proven to be rugged, and a high diving speed has been demonstrated. The airframe is therefore well suited for a higher kW (hp) application. Thirdly, the lightweight turboprop can be more easily substituted for a four-cylinder piston engine such as the Model 201's Ly- coming IO-360-A3B6D than for a six-cylinder piston engine. Substitution for the six-cylinder engine requires a much longer nose section for balance and this per- turbs airplane stability, i_hile this is certainly not an insurmountable obstacle, the work involved would have limited the extent of analyses of other airplarws during the trade studies. Finally, a potential exists for pressurizing the compact Model 201 cabin.

Prior to initiating performance analyses involving piston and turboprop-powered versions of the Aerostar, Cougar, and Mooney, three-view drawings of each airplane were obtained from the respective manufacturers together with aerodynamic and weights data. P_erostar and Mooney FAA-approved flight manual data were also ac- quired. Cougar flight manual data was unavailable at the time.

Three-view dram_ngs of the turboprop-powered version of each airplane were genera- ted by modifying the piston airplane drawings as shown in Figures 19, 20, and 21.

Twin-engine airplane piston/turboprop nacelle comparison drawings were also genera- ted (Figures 22 and 23) to illustrate the much more compact and streamlined turbo- prop engine nacelles. The three-view drawings, aerodynamic data, and the weights information were then used to develop input data lists for the GASP progrnm.

The initial GASP runs involved the piston-ln_wered Mooney and Aerostar airplanes and attempts to match airplane performance data published in the flight manuals. After several iterations during which input adjustments were made, a suitable performance match was achieved for each airplane; i.e., takeoff distance, climb rate, maximum speed, and landing distance were matched.

T/P engine data were then inserted in GASP together with input changes to account for powerplant, propeller, pressurization, and aerodynamic differences. Gross weights of the retrofitted Aerostar and Mooney were held to the piston airplane values. Because the retrofitted Cougar was overpowered at the piston airplane's gross weight, two seats were added and the gross weight increased by 408 kg (900 lbm). Each turboprop-powered airplane, at gross weight, carried substantially more fuel than the piston counterpart when payload was held constant (see Tables XVI and XX).

Results of the GASP computer runs are shown in Tables krv thru XXI for the Aerostar, Cougar, and Mooney, respectively. Note in Table XV that the turboprop-powered Aerostar takeoff is shorter, climb rate faster, ceiling higher, and cruising range substantially greater. Its fuel efficiency at altitudes above 4572 m (15,000 ft) is improved. At 7620 m (25,000 ft) the gain is about 15 percent and at 10668 m (35,000 ft) it is about 40 percent better then the piston at optimum cruising altitude.

Cougar performance, too, is greatly enhanced by the turboprop retrofit. In fact, the conversion transforms the airplane into a wholly new performance class as shown by Table XVIlI. Note the 250 percent increase in climb rate and 44 percent improve- ment in seat-km/[ (seat-nm/gal). Engine-out climb rate and single engine ceili,g (not shown) are also decidedly higher with obvious safety benefits.

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NASA CR-159603 78-I13-15 WRC Report No.

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NASA CR-159603 WRC Report No. 78-113-15 J GROSS WEIGHT = 2132 kg (4700 tbm) EMPTY WEIGHT = 1107 k0 (2441 Ibm) PASSENGERS = 6 t / A-9205 Figure 20. Turboprop Version of the Gulfstream American Cougar

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6] NASA CR-159603 WRC Report No. 78-113-15 TABLE XVI. AEROSTAR 601P WEIGHT BREAKDOWN COMPARISON (GASP DATA) TURBOPROP PISTON COMPONENT/CROUP kg Ibm kg Ibm Propulsion Group Primary Engines 161 354 504 1112 Primary Engine Installation 45 100 96 211 Fuel System 30 67 15 32 Propulsor Weight 82* 180" 86 190 Total Propulsion Group Weight 318 701 701 1545 Structures Group Wing 261 576 261 576 llorizontal Tail 33 72 32 _ I Vertical Tail 17 37 16 36 Ftmelage 239 528 228 502 IJmlimg Gear 122 270 122 269 Primary Engine Section 52 114 I27 280 "r Total Structures Group Weight 725 1598 786 1734 Filet Controls _o_ Cockpit Controls 10 23 I0 23 Fixed Wing Controls 40 88 35 77 Total Controls Group Weight 50 111 45 i-O0 Weight of Fixed Equipment 291 641 291 641 Weight Empty 1384 3051 1823 4020 Fixed Useful Load (Inc. Crew of 1) 125 275 125 275 Operating Weight Empty 1509 3326 1948 4295 Payload 454 !O00 454 I000 Fuel 759 1674 320 7O5 Gross Weight 2722 6000 2722 6000 * 1988 Technology Propellers (composite blades) assumed.

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NASA CR-159603 78-113-15 WRC Report No.

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TABLE XVI I. AEROSTAR 601P DRAG COEFFICIENT BUILDUP COMPARISON (GASP DATA)

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COMPONENT PISTON C D TURBOPROP CD o o 0.00837 0.00795 Wing 0.00702 0.O0732 Fuselage 0.00088 0.00083 Vertical Tail 0.00218 Horizontal Tail 0.00230 0.00208 0.00294 Engine Nacelles 0 0.00150 Cooling Drag 0.02095 Total 0.02242 m i Turboprop CD = 0.0209 + 0.0506 C.L2

M,te_ c v = 0.0224 + o.o5,o eL"

7¸ Flight conditions for Reynolds number and skin friction calculation: 1 Turboprop - Bi = 0.400 at 10668 m (231 knots at 35,000 feet) 2 Piston - lq = 0.300 at 4572 m (188 knots at 15,000 feet) NASA CR-159603 WRC Report No. 78-113-15 TABLE XVIII.

GULFSTREAH AMERICAN COUGAR PERFORHANCE COHPARISON (SI Units) TURBOPROP I PISTON Passengers 6 4 Gross Weight - kg 2132 1724 Pressurized, (&P) - kPa Yes (20.7) No Engine Rated k_, each 227 Takeoff Distance (SL, Std Day, GW) Ground Rum - m 237 305 Over 15.2 m Obstacle - m 521 ii, lrt_ Setti_ - r_ .... 0.349 0.262

I"

:&re of Climb (SL, Std Day, G_ - m/sin Tim to Climb to 4572 m - min 6.2

30 (appx) !

B_ximumCruise Speed - km/h 491

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Service Ceiling - m 11582 5578 Range (45 min reserve) Altitude - m 4572 7620 Range - km 1241 1791 Speed - km/h 326 380 296 Max. Fuel With Full Seats & Bags 2 - 1 537.5 537.5 283.9 km/l 2.94 4.18 4.35 Seat-km/l 17.6 25.1 17.4 Landing Distance (SL, Std Day, GW) Haxlmum Performance'.

Over 15.2 m Obstacle - m Ground Roll - m

405 it

Flap Setting - tad 0.698 0.419 TIT for takeoff and climb = 1200°K. Maximum Cruise TIT = I144°K.

Fuel flow penalty for power extraction and bleed assumed at 4.25%.

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_tP Assumes 90.7 kg for each passenger and his baggage.

Can be substantially shortened with reverse thrust.

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NASA CR-159603 W_C Report No. 78-113-15

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TABLE XVIII. GULFSTREAM AMERICAN COUGAR PERFORMANCE COMPARISON (English)

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TURBOPROP I PISTON

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b Passengers 4 Gross Weight - Ibm 4,700 3,800

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No Pressurized, (AP) - psi Yes, (3) 3O5 Engxne Rated hp, each

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Takeoff Distance (SL, Std Day, GW) Grouad Run - ft 119 1,000

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Over 50 ft Obstacle - ft 1,710 1,850 20 15 Flap Setting - deg i Rate of Climb (SL, Std Day, 6W) - ft/min 3,000 1,200

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6.2 Time to Climb to 15,000 ft - min 30 (appx)

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265 168 tL_immsCraise Speed - knots 18,300 Service Ceiling - ft 38,000

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Range (45 mln reserve) Altitude - ft 15,000 25,000 8,500 670 967 Range - run

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176 205 160 Speed - knots 142 142 75 Max. Fuel With Full Seats & Bags 2 - gal 6.00 8.55 8.89 nm/gal

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360 51.3 35.6 Seat-nm/gal Maximum Performance Landing Distance (SL, Std Day, GW) Over 50 ft Obstacle - ft 1,402 1,330

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Ground Roll - ft 710 40 24 Flap Setting - deg

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TIT for takeoff and climb = 1700°F. Maximum Cruise TIT = 1600°F.

Fuel flow penalty for power extraction and bleed assumed at 4.25_.

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Assumes 200 lbm for each passenger and his baggage.

Can be substantially shortened with reverse thrust.

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_5 NASA CR-159603 _C Report No. 78-113-15 1 w 1 _ ¢'q u" 0 0 O Z _._ O 00 u'_ 0 r_ _ ,--_ u_

_oo

o,I O _._ _ _'_ ..._ Oi 0 ¢q ¢q _0_ [.-., [.-, ffl O _rm _D 0 "--' c_-._ r_C _-_ °°

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.,._.._ ! _ _ _,_ O 0

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_C 0._ < _.-_ O _J _ N ¢,,I ¢'0 NASA CR-159603 WRC Report No. 78-1].3-15

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TABLE X!X. MOO_NEY 201 PERFORMANCE COMPARISON (English)

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POWZRPLAN, T. TYPE MqD [SOURCE, OF DATA) PISTON (I_DBOOK) TURBOPRO[ '1 (GASP) [ PlsroN (GASP) , ',' _ _"....... .,- . ,, ,

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4 4 4 Passengers Gross Weight - Ibm 2,140 2,740 2,740 Pressurized, (AP)-psi Yes (7.5) No No

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305 2O0 2OO Engine Rated hp ,. J ,, , Takeoff Distance (SL, Short Normal Std Day, GW)

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Ground Run - ft 934 890 931 Over 50 ft Ob- stacle - ft t,649 1,51R.. 1,771 Rate of £timb'{$L,' StdDay, _) = ft/min 2,445 1,020 1,023 T_ to Climb to /.6 26 I5,000 ft - min 25

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2_3 175 175 Naximut Speed - knots Service Ceiling - ft 40 ;O00 1_,709 .... 18,700 Range (45 Min Reserve) Altitude - ft 25,000 35,000 8,000 8,000 576 t26 524 537 gange - mm 177 195 162 162 Speed - knots Max Fuel with 52 52 45 45 4 Pssgrs 2 - gal 15.95 20.49i 15.78 nm/gal 63.8 82.0 63.1 Seat-nm/$al i, Landing Distance (SL, Maximlm_ Performance Std Day, GW) Over 50 ft Ob- stacle - ft 1,603 1,602 1_610 Ground Roll - ft 7803 780 770 TIT for takeoff and climb : 1700°F. Maximum cruise TIT = 1700°F. Fuel flow penalty for power extraction and bleed assl_ed at 4.25%.

Assumes 200 lb for each passenger and his baggage.

Can be substantiall.y shortened with reverse thrust.

I ,4 NASA CR-159603 WRC Report No. 78-113-15 TABLE XX. HOONE¥ 201 WEIGHT BREAKDOWN COMPARISON (GASP DATA) TURBOPROP PISTON ibm COMPONENT/GROUP kg Ibm kg ,, , Propulsion Group 80 177 179 394 Primary Engine 23 50 24 53 Primary Engine Installation 42 16 19 35 Fuel System 40 8g 29 64 Propulsor Weight 162 357 248 547 Total Propulsion Group Weight Structures Group 132 290 132 290

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Wing llerizoatal Tail 26 57 24 54 Vertical Tail 12 27 12 194 427 149 328 Fuselage 47 103 47 103 Landing Gear 9O4 364 410 802 Total Structures Group Weight Plight Controls Group 9 2O 9 2O Cockpit Controls 15 34 14 3O Fixed Wing Controls 25 55 23 50 Total Controls Group Weight 109 240 109 240 Weight of Fixed Equipment 706 1556 743 1638 Weight Empty 230 104 230 Fixed Useful Load (Inc. Crew of 1) 810 1786 847 1868 Operating Weight Empty 272 600 272 600 Payload Fuel 354 123 272 1243 2740 1243 2740 Gross Weight

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NASA CR-159603 WRC Report No. 78-I13-15

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TABLE XXI. MOONEY 201DNAG COEFFICIENT BUILDUP COMPARISON {GASP DATA)

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COMPONENT PISTON C D TURBOPROP CD

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O 0.00835 0.00870 Wing 0.00546 0.00580

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Fuselage 0.00067 Vertical Tail 0.00069 0.00176 Horizontal Tail 0.00184 0.00130 0 Cooling Drag

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0.01703 0.01753 Total

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Turboprop CD = O.O170 + 0.0514 CL2 Piston c o = o.o175 + o.o515 CL2

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Flight conditions for Reynolds number and skin friction calculations:

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1 Turboprop - H = 0,32 at 7620 m {193 knots at 25,000 feet) ] 2 Piston - H = 0.249 at 2438 m (160 knots at 8000 feet)

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NASA CR-159603 WRC Report No. 78-113-15 The rate of climb of the retrofitted Hooney has been more than doubled as has the altitude capability (Table XIX). A modest range improvement is also shown which would have been better had the cabin design AP for pressurization been lower and the corresponding fuselage weight increase less (Table XX). The turboprop Hodel 201_ has sufficient excess power to permit an increase in the certificated gross weight and an increase in allowable fuel load. This action could lead to a sub- stantial increase :in range capabi[kty.

These airplane trade studies have demonstrated that the turboprop can provide benefits to general aviation through enhanced airplane performance. The general aviation pilot must be willing to operate his airplane at altitudes above 4572 m (15,000 ft) to maximize the gains, however, and this requires that most operations be by instrument-rated pilots under controlled flight conditions. By 1988 the percentage of controlled flight operations is expected to be substantially higher than today, and the instrument-rated pilot population is expected to grow from today's 222,000 to about 380,000. Thus there should be little shyness among the pilots of that era about high-altitude operations. The conceptual engine can therefore be considered able to surmount the fuel efficiency impediment that has constrained small turbine engine sales (ref 11), at least for business-use airplane applications.* It must also be abie to pass through the first cost and LCC bar- riers. The feasibility of this is discussed in the LCC subsection.

TURBOFAN Parametric Stud__ Parametric performance data were generated to facilitate T/F performance optim- zation studies. Data plots were used to show the relationship between TSFC and specific thrust for several compressor temperature rise values between i67°K and 444°K (300°F and 800°F) and several bypass ratios. The AT of compression was selected to be a variable rather than the conventional compressor pressure ratio, because this form of presentation is believed more representative of actual opera- tionat modes of turbine engines. Also, a better understanding is gained of how specific compressors will operate at varying inlet temperatures. Figure 24 is a schematic of the parametric study T/F engine and includes a list of study assump- tions, variable values, and nomenclature. Samples of the type of parametric curves generated are shown in Figures 25 and 26.

TurbofanDesign Characteristics Selection The selection of a design concept for a turbofan is as complex a process as the previously discussed turboprop design selection procedure. First, the maximum takeoff and cruise turbine temperatures are selected. The maximum takeoff turbine temperature and the maximum cruise temperature chosen here were based on the design phiIosophy that the achievement of very long life hot section rotors as man,tac- lured by low cost fabrication techniques will require compromise in the d_.sign *The piston engine i:; probably superior for instructional, recreati,mal, and most proficiency flying since these are usua}ly done at low altilude an,"., i_lw_]vo lJ-e- quent changes in altitude and direction of ilight. Such changes ar_" nr,t desirable in enroute airspace _,t altitu,tes above 3048 m (lO,O00 ft) _'Lcre t.hc 4_, km/'h_ (25!)

knot) speed limit is not in effect, instructional flying alone accounts for about 25 percent of all general aviation iiight hours (re[ t2).

7O

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NASA CR-159603 WRC Report No. 78-Ii3-15

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"T F I

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ENGINE ASSUMPTIONS PARAMETRIC STUDY ASSUMPTIONS 1) Fuel heating value = 1) No power extraction or bleed 42800 kJ/kg

t

2) Ram recovery = 1.0 (18400 Btu/lb) 3) Bypass ratio (BPR) variable 2) Bypass duct pressure loss = between l and 7

!

3.5 percent ._) .7I" cor.:pr_s.si_,n (/ic'., vari:ib]<, 3) Burner pressure loss = between 167 K (300o!: ) and 444OK o .

1 4.0 percent

4) Burner efficiency = _) (h.,el-;t} ] _'_ mprt'ssion pYt'_llYt 99.5 percent ratio (P ) defip, cd bv ." lc a;_d 1- "Various comprL, ssor input._ 5) Thrust coefficient = 0.985 _) Fan t_:p,'_L.raturo fist, (ZI_) :_ • 0 ,O., ' 33°K t,_O F) , 4-+ _, t,<_(l'_P ) , and 56°K (h)O°F).

6) HP and LP turbine efficiencies = 0.86 ;) ']urbii_._' in.let tLm_.pt,raturc (i It) = 12OO°K (l?O0°F), ]31 l°K (it_l!O°i :) I .'_ _O , aI'ld ['4...- K (2100°F).

s) At "1"_ = 44°K (SO'_l :] ; fan, liP I compr<'ssor ,tnd turbim, t.!fi,i,mci,,_ (':"' 'C ' '-] ',,',,r_ v,lri,.d to de_ormln<, iilf]ucncv ,.:o,. t:icicnt> 9 ) St.S <.,_d AJt crui::,. ,ti ,.i_ ?'!,:oil ,_mi c_].',',:.. _. :,,,.i)i_ I it _ ',..t.r, ,',,i_;jd<._,.d r ._ , / Figure 24. ,qt'hematic _!- the P:lrmr:t,trl.c ,qt_dv iurbc, f:m Kq.vin,, 7] NASA CR-159603 WRC Report No. 78-I13-15 (Io 1_-,0. 11 SLS q, = 0.87 ki. 0:- • 0, 10 tl, =0.76 _T P.

167°K (300('F '' 3 71 l/, = O. 86 _TF= 44°K (80°F) PR_ = I. 554 .0.09 (0.9)- TIT=1311°K (1900°F3 222 (400; 5.20 i % 4,: J¢ - (o s_ -0, 08 BPR 278 (500_ 7 10 E ,o 333 (600" 9 49 m v ,0, 07 (o. 7Y J_ 389 (700 12 ,t7 444 ('800_ 16. 14 (o. 6)- -0.06 5 F- -0. 05 (o, 5`)- (o. 4") -0.04 3OO 400 (0.3") I

1--1

T (20) (30) (4o'_ (50"_ SPECIFIC THRUST -- F.AN2, N-s/kg (lbl-s/Ibm) A--I0857 • f _ 'T• TF Turbofan Pqrametrlc Study £urves Figure 25. -- ,%L bl_) Dn':', c r..,O

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NASA CR-159603 WRC Report No. 78-113-15

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I

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ALT = 9144 m

(30,ooo.)

_Tc PR MN=0.6 (1. 1)- .o. 11 1GT°K (300°F"_ 4.43 (RAM PR = !. 273)

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222 (400) 6. 42

,/,_= o.87

//c =0.76 _r = 0. 86

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&TF= 44°K (80°F ') P*r "= t. f47 O. 10 (SOO) 9.02

m = m_K (I_PI_

L

/

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(GO0') 12. 34 • i O._J E .o

(7oo) 16.52

u v .= Z (800) 21.73 ol I

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0.08

p. (o.83

m i-

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BPR •0. 07 (o. 73- 6 5

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1o0 200 300 400

(o. 6)

i I L ,,, I I0 2'0 30 4O 50

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SPECIFIC THRUST -- FN/_/2, N-slkg (Ibf-sllbm) A--10858

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Figure 26. Turbofan Parametric Study Curves - 9]44 m (30,000 ft)/_!acll 0.6 Nc = 0.76, DT = 0.86

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7_

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NASA CR-159603 WRC Report No. 78-113-15 stress levels of the materials used. Figure 27 illustrates existing and expected strength versus temperature characteristics of candidate turbine materials for the 1985/1990 timeframe. Conventional high speed spool turbine design generally puts blade stress in the 207 to 24_ .MPa (30,000 to 35,000 psi) region with blade tempera- ture approximately ll!°K (200°F) below gas temperatl_re. Once a determination is made of the maximum turbine entry temperature, the selection o_ design co,npressor temperature rise and bypass ratio is considerably simpllfied. As an example, inspection of Figure 26, which is based on a 44°K (80°F) AT fan design at 9144 m (30,000 ft) and 0.6 Much number shows that a bypass ratio of 4.5 to 5 with a pres- sure ratio of 20 approaches the optimum for a 1311°K (1900°F) TIT.

Presently used materials ate exemplified by IN-100 or MAR-M 246. Very long TBOs (3000 to 6000 hr) can be attained by using these materials whea gas temperatures are held to the 1255°K (1800°F) range. The low speed, low _tress design (approxi- mately one-half the blade stress of conventional designs) will permit higher temp- eratures, longer TBOs, or less stringent mannfacturing techniques. Figure 27 shows several advanced materials which are of interest for improved versions of a basic -t_t_ime _ _esi_. The mterials are HA 600Og, DS EIJYECTIC, and RSR. If ckaraeterization of HA 6000E proves to be as advantageous as it looks now, it will be of great interest for the low stress concept because it can operate at stress levels compatible with 1478°K (2200°F) temperatures with good life expectancy.

The advantages of cooling a small blade turbine were judged to be more than offset bycooling lo_ses and blade shape compromises unless gas temperatures can be raised above 1478°K (2200°F). Thls wo_Id push optimum pressure ratio and bypass ratio ........... to maintain the same specific fuel consumption. The result would be an increase in the cost of manufacture of both the hot end and the cold end of the

L

engine and an increase in development cost and risk. For this reason it was judged more cost effective to utilize the low speed, low stress concept to produce low cost turbofan engines. These engines could operate at moderate turbine tempera- tures and have a potential for growth as advanced materials became real commercial engineering entities.

A turbine entry temperature of 1311°K (t900°F) would indicate that an optimum bypass ratio would be in the 4.5:1 to 5.5:1 range and require overall compression temperature rise values cf 389°K to 444°K (700°F to 800°F). The 44°K (80°F) temp- erature rise fan upon which Figure 26 is based was judged to have the h_ghest practical pressure rise for a single stage that would have good part speed per- formance and stability when designed with a low hub/tip rati,J. The low hub/tip ratio is desirable to permit a high low-pressure-spool speed and thereby minin_ze intermediate compressor design and fabrication problems and reduce the _equired number of low pressure turbine stages. A further constraint on the temperature rise achievable in the fan arises from the desire for a low tip speed to minimize noise. This, of course, influences permissible rotational speed.

Parametric data was prepared for fans of other temperature rise capahiliti_s 3nd for engines having various component performance levels. Three basic turb,_fan design concepts, in addition to an evaluation of a "common core concept," were explored during the cost/performance trades. These ;nc]uded a geared fon F]O7- derivative engine, a tandem spool d_sign, and a concentric-shaft, two-_poot '|'/F.

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NASA CR-159603 WRC Report No. 78-113-15

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ALLOY DENSITY g/cm 3 (Ib/in _)

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7.75 (0° 2.80) S N--100 MAR--M 246 8.44 (0.305) [] DS MAR--M 200 8.53 (0.3 08) (6o)- • SINGLE CRYSTAL 8.53 (0. 308) -400 (MAR--M 247) A MA 6000E 8. I 1 (0. 293) A }yr%_ DS EUTECTIC

1 8.50 (0. 3 07")

l RSR (so)- '300 (4o).

0.

IE i f -" (n (3o)- (D (2o)- ('I0)- I150 1200 ! 250 13 00 1350 1400 1450 (03 , I . I I l, i ., I (1600) (18"0_0 ' ' ' (1700) ( 19oo_ (2ooo> ,(21 oo" TEMPERATURE, °K _F) A-- 1086 1 F£gure 27. Nl.ckel Ba_e Blade At].ov_ ' _ _ , "' _=,,_.,_ NASA CR-159603 WRC Report No. 78-113-15 Geared Fan FI07 Derivative A sophisticated design based on the WR19 series of fan jet engines, of which the F107-WR-I00 and F107-_-400 for the ALCM and Tomahawk cruise missiles are the most notable versions, was evaluated. This design was characterized by a gear-driven fan stage and a four-stage intermediate pressure compressor with a three-stage I_ speed turbine in place of the preseuL two-stage LP turbine. Figure 28 shows the external configuration of one version of this engine.

Tandem SFool T/F In an effort to exploit core commonality and WRC low cost construction techniques, an unconventional turbofan design was generated. The concept was referred to as a "tandem spool fan" because the high pressure spool and the low pressure spool are located on completely independent shafts displaced axially. All liP spool accessor- !

ies are located in the engine tailcone. A very low speed, two-stage fan of rela- tively _ hub/tip ratio was chosen because it extended the low speed, low stress deep _lo_ tO_ prs_icaI limit.

The engine cycIe was optimized through selection of the number of stages of IP compression. The generaI arrangement of components is shown in Figure 29. The performance level was found to be approximately 7 percent below that of the geared fan design, primarily because of additional duct losses resulting from the compli- cated flow passages and cross overs necessary to make the independent axial spools wo_.

Concentric-Shaft_ Two,Spool T/F The third design approach involved a more conventional, two-spool T/F constralned in configuration to enable use of the low cost, low stress design philosophy of the turboprop engine. A fan pressure ratio of l.Z_ was selected for compatibility with the desired airplane performance, and the number of IPC stages was traded against design complexity and performance as shown in Figure 30. In this analysis the aerodynamics of the turboshaft core was held constant and changes to the cycle were achieved by varing only the low speed spool components.

Because manufacturing cost is insensitive to the number of stages when using the low cost design concept, decisions as to optimum stage numbers must be based on other considerations involving dynamics, aerodynamic stability, and bearing sus- pension complexity. The design point chosen was at a bypass ratio of 5.2:1 and a maximum cruise turbine temperature of 1283°K (1850°F). The logic behind the chosen design TIT is the expectation that 1283°K (18500F) turbine temperatures can be tolerated in uncooled low cost/low stress components in the 1988 timeframe.

Stress levels approximately one half that of conventional design practice will ensure a very long life. As material temperature tolerance and cooling technology advance, increases in ,_rbine temperature and engine performance can be anticipate,l while maintaining the low life cycle cost environment generated by the origxnal long life components. Required adjustment in cycle pressure ratio for the improved engines would be accommodated by improved compressor efficiencies and increases in work level of the compressor components.

NASA CR-159603 WRC Report No. 78-i13-15 % F---'_ -_ • ! T

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t,\,\TL _ _ 7,/

Jt -f 6.750 DIA -- ENGINE EXHAUST NOZZLE A--7885 Figure 28. Geared Fan - I000 Pound Thrus _ Class FIO7 Derivative Turbofan g p , NASA CR-159603 WRC Report No. 78-i13-15 O OO m- L C L_ Lr_ C .J _D % ?

Cv = O. 985 _F:)/PDu CT =0. 015 i ('0,,...,.,, .7_T'O" 080 , _P/'PBURN-----O. 035 (o.

o.o,, NOTE: HPC PR, r/ & Q SELECTED

(0. 74)t0. 075 _k_ TO HOLD N2 C_ 35::)00 RPM I _ (0.72)T0.073-- _[_X BASED ON COMBINED AXIAL AND _ _e,_ CENTRIFUGAL LOW COST HIGH SPOOL (0. 70)+0. 071 IT= 1283°K (1850°_ =, [ z (0. 687T0. 069 (0. 66)1"-0. 067

_ (°"47T°'°6s '-s_,_ 3-'3--_

[0. 62)T0. 063 4G_" s 4 [ (0.60)t0. 061 l I , I I (0. 58,-t-0. 059 800 1300 I I I I I (150; (200) (,250) (300) k350) A--1086z NET THRUST- N (Ibtl

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Figure 30. Low Cost Conventional Turbofan Design Concept Optimization Stud,/ - 9144 m (30,00C ft)/Mach 0,6

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NASA CR-159603 WRC Report No. 78-113-15 Preferred ConceEt Analysis of the three design concepts showed that the performance and weight of the geared fan design and the low cost concentric-shaft conventional spool design were comparable. The cost of the geared fan design was more than twice that of the low cost conventional spool design, however. The performance of the tandem spool design was 7 percent below that of the low cost conventional spool engine, _ts weight was b0 percent greater, and the cost was 7 percent higher. On the basis of this comparison, the decision was made to narrow the study to the low cost conven- tional spool design. The final version of this preferred design, termed the P7808 turbofan, is shown in Figure 39.

Turbofan P7808 The P7808 turbofan is configured to use the T/P power section as a gas generator with a conventional, concentric, low speed-spool shaft system. As sized, the T/P LJ Ooze is appropriate for an efficient T/F engine in the _48-N-thrust (lO00-1b- that) class. The core is comprised of a six-stage axial compressor followed by a centrifugal compressor, annular combustor, and two-stage turbine. The low-pressure (LP) shaft system includes a fan that produces a pressure ratio to 1.4 followed by a three-stage, intermediate-pressure (IP) compressor and a four-stage, low-speed tarbine. The three-stage IP compressor and four-stage LP turbine were selected through a trade study involving shaft dynamics and data as shown in Figure 30.

Table XXlI is a component performance summary for a nominally-rated P7808 engine at the SL/static condition and the Hath 0.6 cruise condition at 9144 m (30,000 ft).

Note the excellent cruise SFC for this under-4448 N (under-lO00 lb) thrust engine.

Figures 31 and 32 show the net thrust, Hach number and SFC relationship for SL, standard day, and 9144 m (30,000 ft) flight conditions. Figure 33 illustrates the clean lines of the engine's external surface resulting from internally-mounted accessories and plumbing. This arrangement obviates the need for, and saves the weight of, a separate engine nacelle.

Ai____lane Studies To aid in the evaluation and selection of the most suitable T/F engine of the three that were considered (tandem spool, geared fan, and two spool), airplane perfor- mance analyses were made using each candidate. The analyses involved a basic six-place twin T/F airplane weighing 2722 kg (6,000 Ibm) as defined by the market study. Limited work was also done with an eight-place version of the same air- plane. Because the baseline airplane had aft pod-mounted engines, weight and balance computations were made to ensure engine compatibility. One engine, the tandem-spool configuration, was judged too heavy and a liability from the weight/ balance/stability standpoint for the type of airplane design being considered.

This finding and engine cost considerations led to the eliminatio_ of the tandem- spool candidate. The high cost of the geared fan candidate led to its elimination.

Figure 34 illustrates the baseline airplane with the two-spool P7808 engine in,;tai- led. The GASP-derived performance of this airplane/engine combination is shown in Table XXlII. For comparitive purposes, the performance of T/P and piston versions of the Aerostar 601P are also shown in this table. Note with r_,_pect to fuel efficiency, that the numbers favor the T/P Aerostar. Note also that the turbo- fan-powered airplane compares very favorably with the p_ston-powered Aerostar 601P.

Its range is milch greater with the same payload because the lighter-weiRht t,3rhofaI_ engines permit more than twice as much fuel to be carried, (Table XXIV).

NASA CR-159603 WRC Report No. 78-113-15 TABLE XXII. TURBOFAN ENGINE F7808 COMPONENT PERFO_IANCE SUMMARY (Sheet [ of 2) (SI Units) Takeoff Cruise Alt. - m 0 9144 0 0.6 288.2 228.7 Pa - kPa 101.325 30.089 Fo - N 4346 1112 0.040 0.069

sFc - ksl_-h

TIT - °K 1333 1264 5.32 5.22 18.1 8.92 /i i: 11660 II870 111 - 34320 N 2 - rpm 15-52 19.82 P Compressor r Compressor,Performance F_dm: 3.025 3.640

_i. - (ks/s) 4_/kPa

1. 336 1.407 r O. 863 0.862 q AT - oK 28.9 29.3

]

IPC: O. 376 0.44O _in - (kg/s) °_/kPa i. 858 2.069 M r 0.791 0.797 AT - OK 77.3 79.5 HPC: 0.226 0.241 _in - (kg/s) °4_-K/kPa 6. 806 6.251 0.744 e .749 q r AT - OK 348 Burner Performance 0.050 0.049 _in " (ks/s) °_kPa 0.965 0.965 r 0.995 0.995 Mr@ 42800 kJ/kg 0.0971 Fuel/Air

NASA CR-159603

WRC Report No. 78-113-15

TABLEXXII. TURBOFAN ENGINE P7808 COMPONENT PERFORMANCE SUMMARY (Sheet 2 of 2)

(SI Units) / Turbine Performance HP Turbine: 0.070 0.070 _in " (kg/s) _/kPa 3.596 3. 496 r 0.86 0.86 q LP Turbine: 0.221 0.216

_in - (kg/s) _'_l_a 3.573

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WRC Report No. 7_-I13-15

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TURBOFAN ENGINE P7808 COMPONENT PERFORMANCE SUMMARY (Sheet 1 of 2) TABLE KXll.

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HPC: 4.600 a in (Ibm/s) _rSRIpsia 6.251 r 0.744 n AT - OF ° Burner Perfomance 1.008 1.010 in - (Ibm/s) ,/'6"R/psia 0,965 0. 965 0.995 0.995 qr@ 18,400 Btulib 0.0971 Fuel/Air

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NASA CR-i59603

WRC Report No. 78-113-15

TABLE XXII. TURBOFAN ENGINE P7808COMPONENT PERFORMANCE SUMBARY (Sheet 2 of 2)

(English)

Turbine Performance HP Turbine: 1.427 1.428 (Ibm/s) _dR/psia Qin 3.496 3.596 r 0.86 0.86 q LP Turbine: 4. 404 4.511 - (ibm/s) _/psia Qpin 3.161 3.573 r 0.86 0.86 q 0.011 0.015 (_P/P) sPD 1.0 1.0 _(NL/Wcore) O. 982 O. 990 CF-Primary O. 980 O. 989 CF-Secondary NASA CR-159603 WRC Report No. 78-113-15

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0--1 _-_ o _ O-_N _'_ O_m_o 0 E-,-i NASA CR-159603 l_C Report No. 78-113-15 TABLE XXIV. TURBOFAN/TURBOPROP/PISTON WEIGHT BREAKIX)_ COHPARISON (GASP DATA) 601P Piston AIRPLANE A-9737 T/F 601P T/P COHPONENT/GROUP kg Ibm kg lbm kg Ibm Propulsion Group 504 1112 168 370 161 354 Primary Engines 96 211 36 80 45 100 Prilary Engine Installation 27 59 30 67 15 32 Fuel System 86 190

Propulsor Weight 18o*

231 509 318 701 701 1545 Total Propulsion Group Weight Structures Group 226 498 261 576 261 576 Wing 42 92 33 72 32 71 Horizontal Tail 28 62 17 37 16 36 Vertical Tail 340 749 239 528 228 502 Fuselage 87 191 122 270 122 269 Landing 6ear 127 280

Primary Engine Section 34 78 52 114

75s 725 1598 786 1734

Total Structures Group Weight Flight Controls Group 10 23 10 23 10 23 Cockpit Controls 39 87 40 88 Fixed Wing Controls 35 77 m _m_ 49 110 50 111 45 100 Total Controls Group Wei&ht 291 641 291 641 291 641 Weight of Fixed Equipment 1327 2927 1384 3051 1823 4020 Weight Empty 125 275 125 275 125 275 Fixed Useful Load (Inc. Crew of 1) 1452 3202 1509 3326 1948 4295 Operating Weight Empty 454 1000 454 1000 454 lO00 Payload 816 1798 759 1674 32O 705 Fuel 2722 6000 2722 6000 2722 6O00 Gross Weight * 1988 Technology propellers (composite blades) assumed.

"- 89 NASA CR-159603 k_,C ]_port No. 78-113-15 Drag buildup data for the three airplanes are compared in Table XXV.

ENGINE-RELATED LIFE CYCLE COSTS (LCC) The Nooney and Aerostar turboprop retrofit studies show that the retrofitted air- planes are competitive with piston-powered counterparts from a t,,el efficiency standpoint provided that the turboprop airplanes are flown at altitudes above about 4572 m (15,000 ft). The twin turbofan-pc:ered study airplane is also as fuel- efficient, or more fuel-efficient, than the piston-powered Aerostar for some mis- sions. The "real-world" efficiency advantage dorives fr,_ the greater ability of the turbofan-powered airplane to surmount fronta_ _,eat_er and fly a straight-line course to the destination.

Fuel efficiency, of course, is only a part of the LCC picture. Engine first cost as _ell as inspection, maintend,t,e, and overhaul costs are also important. The influence of engine-generated vibration on the airframe and propeller also affect ownership costs.

A limited examination of the foregoing was made using the Hooney, Aerostar, and twin turbofan designs as representative airplanes for determining the viability of the conceptual turbine engines. This was done by calculating turbine and piston engine-related ownership costs over a 20-year period assuming a 185 200 bm (100,000 um) annual airplane utilization rate. Engine-related fleet LCC's were also calcu- lated to enabIe a comparison of possible turbine-fleet benefits with the investment costs required to develop and certify the P7757 and P7808 turbine engines. In addition to predicting the relative cost impact of introducing nominally-rated turboprop and turbofan engines, cost tradeoffs involving the introduction of higher technology units (i.e., engines with improved aerodynamic components and higher temperature capabiIity) were evaluated.

Several simplifying assumptions were used to facilitate the LCC determinations.

Because of these assumptions and the fact that only major cost drivers were con- sidered, the LCC data provided in the present report should be considered "figures of merit" only with the relative values having more meaning than the absolute values. Ground rules, assumptions and costing methodology are discussed in the paragraphs that follow.

LCC Analysis Assmsptions The singIe-engine Hooney and twin-engine Aerostar were considered representative of the airplane classes where the introduction of an "optimum-type" GATE turboprop engine would be most likely. Although there are numerous other possibilities, these classes were considered typical for individual-airplane and fleet cost an- alyses, and for turboprop/piston engine-related LCC comparisons. The twin turbofan / s_udy airplane was considered typical of the type that could use the P7808 turbofan engine, and it was used as the basis for turbofan LCC analyses. Because there was no piston-powered counterpart for the twin turbofan design, individual-airplane and fleet LCC comparisons were made with the piston-powered Aerostar 601P.

It was assumed that each airplane analyzed would travel 185 200 km (I00,000 im) per year. Also, for computation purposes, a typical trip of llll km (600 nm) was assumed. Trip block times and POL (petroleum, oil, and lubricants) usage and cost were calculated using the GASP program for turbine-powered airplanes and flight .NASA CR-159603 WRC Report No. 78-113-15 TABLE XXV. TURBOFAN/lqJRBOPROP/PISTOX DRAG BUILDUP CONPARISOH (GASP DATA) i AIRPLANE 601P T/PZ[ 601P Piston 3 A-9737 T/F j m 2 Equivalent Flat Plate Area (C D Sre f) COMPONENT o (ft 2) 0.1016 0.1384 0.1314 Ning ( 1. 094) (1.490) (1.414) 0.0997 0.1211 0.1160 Fuselage (1.073) (1.303) (1.249) Vertical Tail 0.0201 0.0145 0.013/ (0.216) (0.156) (0.148) Horizontal Tail 0. 0308 0.0380 0.0360 (0.332) (0.409) (0.388) O. 0186 0.0344 0.0485 Engine Nacelles (0.2O0) (0.370) (0.522) Incremental 0.0035 0 0.0248 (0.038) 0 (0.267) Total 0.2743 0.3464 0.3704 ° (2.953) (3.728) (3.988) A-9737 TIF C D = 0.0236 + 0.0696 CL2 (Sre f = 125 ft 2) 601P T/P CD = 0.0209 + 0.0506 CL2 (Sre f = 178 ft 2) 6OlP Piston CD = 0.0224 + 0.0510 CL2 (Sre f = 178 ft 2) Flight conditions for Reynolds number and skin friction calculations: (Note: Second iteration runs were not made at speed for best specific range) XA-9737 T/F - M = 0.450 at 10973 m (259 knots at 36,000 it) 2601P T/P - H = 0.400 at 10668 m (231 knots at 35,000 it) 3601P Piston - H = 0.300 at 4572 m (188 knots at 15,000 ft) NASA CR-159b03 NRC Report blo. 78-113-15 manual data for the piston-powered airplanes. _ Turbine-engine yearly operating times and associated engine inspection, maintenance, and overhaul frequencies and costs were determined on the basis of trip block times and the number of trips per year (166.7). Because helicopters are used for many tasks where trip mileage has little significance (e.g., cargo transfers using a sling), the turboshaft LCC figures are based on 500 hours per year operation.

An underlying premise for the prediction of GATE-engine Iife cycle costs is the attainment of a 10,O00-hour time between overhaul, a figure consistent with antici- pated airframe life. Although attainment of this high a TBO is an ambitious ,nder- taking, it is believed feasible, at least with respect to the proposed turboprop engine, due to the low speed/low stress design characteristic. Many thousands of engineering hours and millions of development dollars will, of course, be required to achieve this goal. For ease of calculating, the lO,OO0-hour TBO objective was assumed achievable by early prod_.:tion engines planned for introduction at substan- tially derated power levels in airplanes of the Cougar and Hooney type. The later production, more mature, higher horsepower engines for airplanes like the Aerostar were also assumed to achieve a lO,COO-hour TBO.

The 20-year LCC predictions are based on constant year economics (calendar year 1978 dollars). In the case of aviation gas costs, which increased by about ten percent during 1978, mid-1978 costs apply.

LCC Hethodology and Predictions Engine-related life cycle costs were assumed to be influenced by four major cost drivers: a Initial Investment • Production Unit Price _p • POL • Inspection, Haintenance and Overhaul There are, of course, other influences on life cycle cost, but these were not _b considered for the comparative purposes of the present study.

• Initial Investment Investment costs were determined through the establishment of a development and 4, certification plan for each turbine-engine type assuming maximum core-engine parts couuonality. The initial-investment estimate was made using a "bottoms up" or build-up estimating approach that considered acquisition of production tooling sufficient to meet the annual delivery rates sho_rn in Table XXVI. The estimate was prorated among the various T/P, T/S and T/F applications on the basis of parts commonality and total engines within each category.

*GASP and flight manual performance data can be used synonymously for the piston powered airplanes since GASP airplane performance was made to match flight manual airplane performance.

NASA CR-159603 NRC Report No. 78-113-15 TABLE XXVI. ANNUAL PRODUCTION QUANTITY ESTIHATES FOR ENGINE PRICING Turboprop Engines Projected total unit sales per year in 1988 [134-231 kW (180-310 shp) class}, Table 1X = 13,803 Number of units assumed produced annually by one engine manufacturer (used for pricing estimate). 7,555 (Includes 5137.44 units for single engine applications and 2417.56 units for twin engine applications.)

Turboshaft Engines Projected total unit sales per year in 1988 [134-746 k_ (180-1000 shp) class], Table VIII = 1881 Humber of units assumed produced annually by one engine manufacturer [134-231 kW (180-310 shp) class] (used for pricing estimate).

l ii l L l J ii Turbofan Engines Projected total unit sales oer year in 1988 (all thrust levels), Table VIII = 3676 Number of units assumed produced annually by one engine manufacturer [4448 N (I000 Ibf) thrust 1,115 class] (used for pricing estimate).

NASA CR-159603 N1RC Report No. 78-113-15 There were no investment costs associated with the piston engines used for the LCC comparisons because these engines are fully developed and in service.

• Production Unit Price Production turbine-engine unit prices (equivalent to OEM prices but less product liability insurance allowances) were estimated using the following methodology: 1. Industrial engineering estimates were made of direct labor for fabrica- tion, assembly, and test based on cross sectional drawings of each engine configura- tion. These estimates were developed in terms of "standard hours," hours that do not consider shop efficiency. Tooling concepts in keeping with relative11 high production delivery rates were assumed.

2. Improvement curves were developed for the appropriate quantities, and from these curves variance factors were computed and applied to the standard hours to predict total hours for fabrication, a_embly, and test.

3. Hanufacturing and engineering support hours and direct cost dollars (i.e., sustaining manufacturing engineering, tooI maintenance, inspection, etc.)

were estimated using cost estimating relationships (CElt's) deveIoped from histori- cal data. These CER's are based on a percentage of fabrication, assembly, and test hours or a percentage of material cost.

4. Once the total direct labor and direct cost dollars (DC$) were projected, a total price was developed using _RC CY 1978 direct labor rates, burden rates typical of a production mode, and a profit. Table XXVII summarizes production engine pricing information.

• Petroleum, Oil, and Lubricants (POL) POL costs were projected by assuming all airplanes would travel 185 200 km (I00,000 ha) per year. The stage length of a typical trip was assumed to be llll km (600 nm)_, and each trip included 0.2 hr ground maneuvering time. Turbine airplane trip fuel usage was determined using the GASP program and piston airplane fuel usage was calculated from flight manual data where optimum flight profiles could be deter- mined more expeditiously. As already discussed, tests were made of the GASP pro- gram to ensure the accuracy of the output.

Fuel costs were based on a survey of prices being charged at local airports in mid-1978. At that time Jet A prices averaged $0.207/liter ($0.784/gai) and 100- octane avgas was priced at $0.227/liter ($0.86/gal). POL costs were based on these figures plus an allowance for oil and lubricants that amounted to $0.005/ liter ($0.02/gal) of Jet A us q and $0.00811iter ($O.031gal) of avgas used. Table XXVIII shows how POL costs we_ _etermined for the twin turbofan study airplane_ Table • XIX gives an example o the methodology used to determine the 20-year fleet _The most recent nationwide survey conducted by the FAA and the Civil Air Patrol indicated the average stage length of all business-use jet aircraft to be 891 km (481 rim). Piston and turboprop airplane stage lengths can be expected to be less.

The IIII km (600 me) assumption was made with the expectation that, because of the energy situation, short range, inefficient flights (tha_ lower the stage-length average) will be curtailed in 1988.

NASA CR-159603 WRC Report No. 78-113-15 TABLE XXVII. PRODUCTION ENGINE PRICING Production unit prices are based on: Industrial Engineering estimates of direct labor and material/sub- contract dollars using preliminary cross-sectional drawing_ for each engine • Direct labor and burden rates based on an LCC study for a similar direct labor base. Prices are expressed in terms of C¥ 1978 dollars.

Projected engine quantities are for a 20-year ti_e period using a 20 percent--per-year build-up rate until the maximum annual production r,_te shown in Table XXVI is reached. The rate then remains constant for the subsequent 15 years.

The theoretical price of the first production unit and the average production unit price for each engine type are shown below: Number of Units Price of Average Production First Unit Price Per Unit At Lot Midpoint Engine Type | * $19,515 67,995 Turboprop " $23,000 5,580 !Turboshaft 26,163 , 35,000 :urbofan 25,352 10,035 I 40,100 • L, -,._ ._ J _a II II lU

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• Inspection, Maintenance, and Overhaul Mooney 201 and Aerostar 601P piston airplane inspection, matntenance and overhaul costs were obtained from manufacturer-supplied data. For the Hooney, piston en- gine-related inspection and maintenance costs were assumed to equal one-half of the airframe-plus-engine figure. The fraction for the twin-engine Aerostar was assumed to be two-thirds. Shop labor rates were adjusted upward to $20.00/hr for consis- tency and to match turbine-airplane shop labor rates. The assumption of equal turbine/piston shop labor rates is believed reasonable in the light of a clrr_-1988 market scenario for general aviation wherein turbine-powered airplanes are ass_ed to dominate the fleet additions. Tables XXX and XXXI show the cost data supplied by Hooney and Piper and the adjustments made to this data to arrive at a yearly allowance for piston-engine inspection, maintenance, and overhaul.

An underlying premise in projecting the life cycle costs for the turbine-engines is the achievement of a TBO of the order of that of the airframe (10,000 hours assumed). At a travel rate of 100,000 nm per year for each year of the 20-year LCC study, the twin-engine, turbine-powered airplanes would accumulate less than 10,000 hours. The turboprop-powered Hooney would accumulate about 13,000 hours when flown with a derated powerplant. Because no engine can be expected to run flawlessly for 20 years without some kind of parts replacement, a contingency reserve was set aside to permit the replacement of deteriorated parts and parts damaged by foreign objects. This reserve is sufficient to permit one complete engine replacement in 20 years and it has been prorated over this period.

Specific inspection and maintenance actions as well as associated frequencies were identified for each turbine-engine type. With the exception of occasional filter and ig:iter plug replacements, the proposed engines were considered relatively maintenance-free. An isotope inspection was included at 500-hour intervals to verify the absence of cracking or other deleterious conditions in critical parts.

A list of the specific maintenance actions identified and the corresponding cost estimates are given _n Tables XXXII and XXXlII. Twenty-year piston- and turbine- fleet inspection, maintenance, and overhaul cost summations are provided in Tables XXXlV through XLII. These summations were made using the engine fielding factor described in Table KXIX.

Turbine/Piston LCC Comparisons In order to get at the cost benefits, if any, of going to turbine power, the direct operating cost (DOC) and engine production cost estimates previously discussed were combined and piston/turbine cost comparisons made. The 20-year summations were based on 20 percent per year turbine-engine production build-up rates to the pre- dicted peak values shown in Table XXVI. Thereafter, the annual production rates were assumed constant at the peak values. The engine-related cost predictions were made on the basis of the total engine population in the particular year of inter- est.

As an additional aid for evaluating the monetary and fuel economy implications of a general aviation industry movement toward the expanded use of turbine powerplants, information of the following type has also been provided in the LCC summaries: NASA CR-159603 WRC Report No. 78-113-15 TABLE XXX.

ENGINE-RELATED INSPECTION, MAINTENANCE, AND OVERHAUL COSTS PISTON MOOHEY 201 Airplane Hanufacturer Data $ Per Operating Hour • Airframe and Engine Inspection and Maintenance 3.48 (includes a small allowance for parts replacement) • Engine Overhaul (O/H) Allowance 3.13 Adjusted Data _ Engine Inspection and Maintenance (assumes engine- 2.68 related inspection and maintenance cost equals approximately one-half of airframe-plas-engine figure) 0.5 x $3.48 x (20/13) = $2.68 Engine O/H Allowance (assumes field O/H) 4.81 $3.13 x (20/13) = $4.81 Total Eagine Maintenance and O/H Allowance S7.49/hr Yearly Allowance for Ensine Inspection T _aintenance_ and O/H $5063 • 185 200 km/yr (100_000 nm/yr) 274 km/h (148 knots) = 676 hr/yr 676 x $7.49 = $5063 *Manufacturer data based on $13/hr for shop labor. This figure adjusted to $20/hr for consistency with shop labor rates applicable to turbine-powered aircraft (equal piston/turbine labor rates assumed for 1988).

NASA CR-159603 NRC Report No. 78-113-15 TABLE ]orxI. ENGINE-RELATED (ONE ENGINE) INSPECTION, MAINTENANCE, AND OVERHAUL COSTS - PISTON AEROSTAR 601P Airplane Hanufacturer Data $ Per Operatin$ Hour $13.00 Airframe and Engine Inspection and Haintenauce (includes $1.00 all>wance for propeller and governor O/H) 12.01 Engine Exchange Allowance (two factory remanu- factured engines) Adjusted Data _ 17.33 Inspection, Haintenance and Propeller/Governor O/H $13.00 x (20/15) = $17.33 - 1.33 Less PropellerlGovernor OIH $I.O0 x (20115) = $1.33 Airframe and Engine Inspection and Haintenance $16.00/hr Engine Inspection and Haintenance (assumes engine* related inspection and maintenance cost equals two- thirds the airframe-plus-engine figure) 0.667 x $16.00 = $10.67 • Inspection and Haintenance Cost Per Engine = 0.5 x $10.67 $ 5.33/hr $2448 Yearly Engine-Related Inspection and Haintenance Cost • 185200 im/yr (100_000 na/yr) 403 I_/hr (218 knots) = 459 hr/yr 459 hr/yr x $5 33/hr = $2448 $2759 Engine Exchange Allowance (one engine) • 459 hr/yr x 0.5 x $12.01/hr = 2759 (adjustment for factory labor rate not required) $5207 Yearly Allowance for Engine and Propeller Inspection, Haintenance# and Overhaul (one engine) $2448 + $2759 = $5207 *Hanufacturer data based on $15/hr for shop labor. This figure adjusted to $20/hr for consistency with shop labor rates applicable to turbine-powered aircraft (equal piston/turbine labor rates assumed for 1988).

NASA CR-159603 _rRC Report No. 78-113-15 TABLE XXXII. TURBINE-ENGINE SCHEDULED INSPECTION AND t/AIFFENANCE COST Scheduled Estimated Parts Frequency Maintenance Event (hr) Hanhours Costs, $ IOO 0.8 35 Oil, Oil Filter and Fuel Filter (replace oil and clean or replace filters as required) IOO 0.5 2 x 265 Check Igniter Plugs (replace at 5OO-hr) o .

0.5 Inspect Wiring, Tubing, Connections, and Screws 100 0.5 Chip Detector (inspect and clean as necessary) 500 1.5 75 Isotope Inspection lO0-hr total 2.3 $ 35 500-hr total 3.8 $64O

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_J _ .,-_ _.1 I-0 0 ::3 .,-'* I _ "_ _ .,-, ca X U -,'_ ",-_ _J _ ,_ 0 _'+ -_-_ _J _ [-- 0 _J _ _ _" •_ _ 0 _.-_ .,.._ ,._ , r./3 . '.-.-" 3_ III NASA CR-159603 WRC Report No. 78-113-15 • Life Cycle Cost/Engine • $/seat-km (seat-nm)/Airplane (engine-related dollars only) • Liters (gals) of Fuel/Year/Airplane Table XXXIV provides 20-year turboprop/piston engine-related LCC comparisons for single- and twin-engiue airplane applications, and Table )CZXIX provides turbofan/ piston LCC comparisons for twin-engine airplane applications only. Nominal turbine engine component efficiencJes and turbine inlet temperatures are asst_ed. Note that turbine engine LCC benefits are exhibited for all applications, while fuel economy benefits are exhibited for the twin-engine turboprop airplane application only.

Because the potential for bettering piston engine economy appeared good for all applications, an investigation was made of the influence of improved turbine engine component efficiencies and turbine inlet temperature capabilities. The following improvement combinations were studied: • Engines with components having nominal efficiencies and uprated turbine inlet temperature capabilities ITurboprop AT Z 194°K (350°F)Turbofan AT " 145°K (260°F)].

• Engines with components having improved efficiencxes attd nominal turbine inlet temperature capabilities.

• Engines with components having improved efficiencies and uprated turbine inlet temperature capabilities.

Tables XLIII, XLIV and XLV compare the performance of the several engine variants. The estimated impact of the performance improvements on invest- ment requirements, production engine pricing, POL cost, fuel efficiency, etc., is shown in Tables _'v_IV through XLII. Note that component effic- iency improvements (with or without gains in TIT capability) are sufficient to tilt the turboprop/piston furl efficiency advantage in favor of the turboprop for the single-engine airplane application.

Turbofan TIT gains are required before the twin turbofan study airplane can match the fuel-efficiency of the Aerostar, however.

Airplane Life Cycle Costs The primary cost impact of introducing turbine engines to airplanes in the under- 2722 kg (6000 lb) weight class can be expected to be engine-related. There will be airframe-related cost influences also due to the lessened engine-generated vibra- tions. The lowered vibration environment will reduce the airframe fatigue cracking and chafing problem common to piston-powered airplanes and prolong the life of controllable propellers anti avionics equipment. The potential for additional LCC savings through the introduction of new-design airplanes is good if the designs take advantage of the characteristic light weight and compactness of turbine- engines. The new airplanes can be made smaller, for example, because of the re- duced engine weight, cooling drag, and nacelle drag. The lighter, smaller air- frames will not require as much propulsivt energy and there will be attendant fuel cost benefits.

NASA CR-159603 NRC Report No. 78-113-15 The higher cost of turbine engines can offset such advantages, however, by influ- encing airplane insurance costs. The added annual insurance burden can be expected to amount to about two percent of the higher hull value.

It would be a very difficult task for an engine manufacturer to develop a truly meaningful piston/turbine airplane LCC comparison, especially a comparison invol- ving 1988 airframe/avionics propeller technology, and no attempt has been made here to do this. Some insight along these lines will, perhaps, be obtainable from Beech T-34/T-34C experience after the T-34C has been in the field for several more years.

BENEFITS OF TECHNOLOGY ADVANCEHENTS Technology advancements which may benefit small engines can be derived from two sources: (I) the technology being developed by the manufacturers of large engines, with or without government support; and (2) technology development programs con- ducted specifically to benefit small engines. The objectives of currently active large turbine engine programs are shown below: Slay OF FORECASTED ADVANCED TECHNOLOGY AI_S FOR LARGE TURBOFAN RESEARCH • Increase turbine entry temperature • Increase pressure ratio • Increase bypass ratio • Increase component performance (fan, low-pressure compressor, high-pressure compressor, high-pressure turbine, low-pressure turbine, fan exhaust, core exhaust) • Noise and emissions reduction • Accessories improvement • Variable cycle designs (split exhausts and fan flows, variable turbine nozzles) • Engine life improvement • Weight reduction • Cost reduction • Increased durability

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Some of the listed technology areas are formalized and attacked through specific programs, while others of general concern are addressed by all turbine engine manu-

! facturer_ in order to stay competitive. Some are dealt with through comoinations

of the foregoing motivations.

While all of these developments are of value to the smaller general aviation tur-

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bine, their relative payoff is somewhat different since the utilization of general aviation aircraft is typically much lower than that of airline aircraft. The lower utilization rate puts greater emphasis on first cost relative to operating cost.

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The specific technology areas that promise the greatest returns for general avia- tion with respect to the economics of purchasing and operating turbine powered aircraft appear to involve: .3 $ NASA CR-159603 WRC Report No. 78-113-15 • Cost reduction throughout • Increased component performance (stage efficiencies and turbine work levels) • Durability improvement and increased life • Accessory miniaturization and reliability improvement Programs to increase component performance (particularly compressor and turbine ef- ficiency) and increase the work capability of turbines are worthy candidates for bringing about the operating cost, fuel consumption, and acquisition cost reduc- tions needed to permit the emergence of general aviation turbine engines. An increased turbine work capability, for example, would reduce acquisition cost by reducing the number of turbine stages required for a specific cycle (assuming he maintenance of good efficiency). Alternatively, a more sophxsticated cycle could be permitted at the same cost.

Increases in turbine entry temperature benefit turboprop and turboshaft engines by improving both specific output and specific fuel consumption. For these engines, the use of temperatures which demand 31ade cooling are limited by the increased first cost associated therewith. In certain cases, benefits, such as being abl_ to cover a larger power range with the same basic engine, may justify the expense of developing and producing the small cooled blades. In any event, the turboprop/ turboshaft engines will benefit in size, weight, and fuel consumption from any probable increase in temperature which can be attained with only a minor, or zero,

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cost penalty. Heans of Lncreasing TIT which are of interest are: improved mater- ials, better corrosion resistant coatings, lower blade stresses, etc.

In the case of small turbofans, turbine inlet temperature must be considered in

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relation to pressure ratio and bypass ratio as discussed previously. The optimum temperatures tend to be low enough to permit the use of present state-of-the-art alloys. Use of temperatures above the optimum always increases thrust but at the

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expense of higher specific fuel consumption. The extraction of blade cooling air further increases specific fuel consumption. In spite of this, the use of cooling my be economically desira_-le in some cases to permit agivenengine design to cover a larger thrust range.

b Because of the large difference in size, the cooling techniques presently being utilized in large engines are not directly usable on small engines. Hence, the problem of raising the turbine inlet temperature of small engines will require special attention to develop better alloys and/or coatings or to develop cooling techniques suitable for small blades.

Ti_e cycle pressure ratio for a small turbine is a compromise between theoretical thermal efficiency, losses associated with excessively small parts, and the costs arising from additional aerodynamic elements and the increased complexity asso- ciated with surge avoidance over the speed range. The last consideration involves both the mechanical complexity of variable vanes and/or blowoff valves and the increased complexity of the system elements to control the variable features. The net result is that small engine pressure ratios wxll probably fall in the range of 10:1 to 15:1. Such ratios give reasonably good fuel consumption together with a tolerable level of complexity.

The fan bypass ratio of a small turbofan is also a trade-off between a number of factors of which the most important are: cruise specific fuel consumption, engine size and associated external drag, _rigine weight, engine cost, and sensitivity to NASA CR-159603 WRC Report No. 78-113-15 inlet and exhaust duct losses. Generally, for cruise speeds of about Hath 0.6, specific fuel consumption is improved by increasing fan bypass ratio for bypass ratios under 8:1. Howeve:, increasing the bypass ratio has an undesirable effect on all of the other considerations. An area of particular concern is the interre- lation between bypass ratio, permissible fan speed, and low pressure turbine load- ing, If the complications and expense of a fan drive gear reduction are to be avoided, the low rotational speed required by a high bypass fan to avoid excessive tip speeds results in a requirement for multi-stage low pressure turbine elemeuLs.

Hence, it is apparent that the development of a moderate temperature rise fan (_T = 28 ° to 45°C) able to tolerate a higher value of _i 2 (W = through flow, N = speed) without an unacceptable loss in efficiency or excessive noise generation would be beneficial.

Improved and reduced cost accessories are needed for small gas turbines, and work toward these ends can be potentially profitable. Accessories such as the fuel control and starter-generztor do not scale down in proportion to engine size and this creates weight and _acelle drag penalties on some small turbofan powered airplane designs. The problem could be alleviated by durable, small, high-speed accessories that take advantage of available high rotational shaft speeds. Remote mounting is a second approach, but reliability problems can result.

Fuel control cost and reliability are major areas of concern. The fuel control on a small turbine engine typically costs as much as a new mid-size automobite and contributes from 5 to 15 percent to engine cost. This situation is further com- plicated by the need for minimum pilot attention since the contemplated aircraft will be mainly operated by a single pilot. A promising avenue for development is an electronic control using state-of-the-art electronic techniques. Problems with this approach are meeting the reliability requirements and attaining sufficient sales volume to justify the nonrecurring expense of the large scale integrated circuits necessary to make the unit cost acceptable.

Another disproportionately expensive item is the starter-generator and its asso- ciated drive train. The use of a high speed alternator and rectifier together with an inverter for starting deserves further study and development.

Large engine development work in the following areas has a more or less direct ap- plication to small engines.

. Noise reduction and suppression. In the small engine, the fre3uencies are higher and the acoustic energy is much less but the same princip]es apply. This knowledge must be applied to make the general aviatlon turbine socially acceptable.

e Engine life improvements and life cycle cost reductions.

Large engine technology programs oriented toward improving engine life and reducing life cycle costs could benefit small engines from the life extension standpoint.

Small engines have problems unique to their size, however, in terms of high specl- fic bearing speeds and an inherently high number of stress cycles due to higher rotational speeds. Bearing, gear, rotor and static structure improvement work, and accessory life increase programs would be attractive.

Large engine work that involves so-called variable cycle designs is now under way.

Some of this work could benefit small engines. Because of the more limited operat- ing envelopes of general aviation aircraft, however, the cost of the complex mech- NASA CR-159603 WRC Report No. 78-113-15 anical arrangements probably is not warranted and cannot be offset by fuel saving economies. As noted above, general aviation operations tend to have a lower sensi- tivity to fuel cost than coJmercial operations because the annual operating hours are typically lower.

Programs to reduce aircraft engine weight are always important, and this is es- pecially so for small turbofan engines. One of the best locations for mounting tIJrbO_ i_ nn th_ _ft fugp]z_e whprp thp _lznp of rotatin_ parts i_ behind the cabin pressure bulkhead and the wing fuel tanks. Aft mounting creates airplane weight and balance problems, however, and excessive engine weight aggravates this.

On the other hand, weight saving on a small turboprop may be somewhat less impor- tant since the turboprop will always weigh less than the piston engine it replaces Furthermore, on single engine airplane designs, the light weight of turboprops sometimes necessitates excessively long nose sections for balance that tend to impair visibility from the cockpit during climb and the landing flare.

The major requirement for small turbine engine marketability ,s cost reduction Any programs for reducing the cost of large engines should be monitored for their possible applicability to small engines and cost reduction programs specifically aimed at small engines should be undertaken.

Large engine programs to improve resistance to damage from foreign objects such as birds and ice as well as simplified anti-icing schemes must be monitored for pos- sible application to small engines.

_o In summary, large engine technical and manufacturing developments should be care- fully monitored to identify and apply those items which can improve the SFC and weight without increasing cost or which can simplify the engine and reduce its cost. Additionally, programs which are aimed at these same objectives but which are appropriate to the peculiar features of small engines should be vigorously pursued. If both of these things are do_e, competitive general aviation turbine engines can be anticipated.

NASA CR-159603 WRC Report No. 78-113-15 i # .

SECTION 6 EVALUATION OF THE COMNON CORE CONCEPT The generaI aviation field of engine applications consists of approximately 180,000 existing aircraft with 200,000 engine installations and nearly 20,000 annual engine installations in new-production aircraft. By 1988 the total installations are expected to grow to more than 300,000 with a proportionate number of engines coming up for overhaul and replacement. Many owners of engines requiring overhaul would be in the market for a more advanced powerplant that would upgrade the utility of their aircraft if such a unit were available. They would look to turbine power if a cost effective installation were offered. With the potential annual market for new and retrofit turboprop installations at more than 16,000 units, the fielding of a low-cost turbopower engine is feaaLble if a method can be devised for reducing development and production costs and achieving competitive fuel efficiencies.

If it were possible to use turbopower generators as the core or power source for not just the turboprop (T/P) engines which are potentially so numerous, but also as the critical power core of turboshaft (T/S) and turbofan (T/F) engines, production cost benefits would accrue to all three engine types. The ability to utilize a common core depends on achievement _f a design concept which can permit the core to be configured substantially independent of those components, which, by their addition, transform the core into a T/S or T/F engine. Because of the much larger numbers involved in core production for T/P engines, if design compromises are necessary, these compromises should be to the advantage of the T/P engine to assure its acceptance. This consideration introduces difficulties with respect to core thermodynamic cycle optimization.

A well performing T/P engine runs at a fairly high pressure ratio. When the T/P core is converted for use in a TIF engine, the pressure ratio becomes excessive for practical turbine inlet temperatures, i.e., when enough compression is added through the additional stages needed to raise airflow to a value which will produce acceptable thrust levels. By judicious design of the core, however, it is probably possible to provide enough flexibility in compressor geometry and shaft speed to enable a common core to be used as an optimum T/P engine component [224 kW (300 hp) class] as well as the high-pressure section of a T/F engine [4448 N (1000 ibf) thrust class].

A significant aspect of the common core concept is the potential for use of a common set of engine accessories such as fuel pumps, oil pumps, starters, genera- toes, and accessory drives. The accessories constitute a 15 to 30 percent cost fraction of turbine engines, especially of the smaller size engines. If accessories can be made truly common, or only minor modifications are necessary to adapt them to the more complex cycles, a large saving can accrue to the benefit of T/F powerplants.

FAMILY OF ENGINES CONCEPT General aviation can be decidedly influenced by the availability of high-perfor- mance, low-cost propulsion. Low cost is influenced by the requirement for o_

NASA CR-159603

WRC Report No. 78-113-15 development, non-recurring production, and maintenance or recurring costs. Low development costs are extremely important, as they have a direct bearing on the willingness of business to invest after considering the degree of risk and the expected return. Low production and maintenance costs are important because of the interaction with the production number base and the market sensitivity to maintenance burden. Typical piston engine lifetime maintenance costs can exceed three times acquisition cost. The negativ_ attitude toward high existing mainte- nance cost could significantly expand the market for low-maintenance engines, thereby improving the rate of retL_rn possible. Also, it is unrealistic to assume that high volume production could significantly bring turbine engine costs down, since the national airspace system cannot support such rates.

The low-cost goals of the GATE program require departure from classic turbine design procedures. Several methods can be employed to reduce non-recurring costs in turbine development. One such method is the use of commonality in the design of multiple product lines. The use of parts and assemblies common to several different engines reduces performance and reliability development costs, and common-design manufacturing and assembly procedures introduce further economies.

Also, the choice of prototype fabrication methods compatible with low-cost aero- dynamic development and low-cost producibility can considerably reduce the cost of achieving objective performance levels by reducing the cost of test and develop- sent hardware.

Turbine engines characteristically exhibit high rotational speeds and high temperatures, creating sensitive design parameters (blade stress, disk stress and vibration, shaft dynamics, material properties, control characteristics, shaft suspension, lubrication, tolerances, and quality controls). Reducing the intensity of these design-induced problems permits the development of high-performance, low-speed components using low-cost production methods. Designs for low-cost fabrication must be considered from inception, and these will be far more effective in a low-speed, low-stress environment than in a conventional aircraft gas turbine.

Figure 27 illustrates the high temperature capabilities of some advanced turbine materials and compares them with two commonly used alloys, IN-100 and HAR-N 246.

One material that is apparently ideally suited for low-speed, low-stress rotors because of its strength at elevated temperatures and its compatibility with the fabrication techniques being explored for low-cost, dual-prcperty rotors, is designated HA6000 E. This material lends itself well to blade forging.

Operating costs and fuel economy are also highly important and must be attacked by matching the low-cost components for optimum performance and by developing compatible controls and accessories.

The following pages describe a low-cost set of tufbine propulsion systems with a wide range of general aviation aircraft applications (light single-engine turbo- props, light and medium twins, helicopters, and small turbofan-powered craft).

The engine concepts presented (Figure 35) are a T/P, T/S, and T/F based on a common core and designed around low-speed, low-stress, low-cost approaches.

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NASA CR-159603 78-113-15 WRC Report No.

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TURBOFAN TURBOPROP TURBOSHAFT A-- 109,_3 Figure 35. GATE Fm.ily of Engines i NASA CR-159603 NRC Report No. 78-113-15 CO_K}N CORE CONCEPT DESCRIPTION The common core concept for this family of engines is a turbine gas generator with the significant parts common to all three engines. As indicated by Figure 36, the major common components are the high-pressure axial compressor, centrifu- gal compressor, and the turbine.

_ll 11 11_ Figure 36. Common Core The leading element of the hiRh-pressure compressor is a six-stage axial compressor exhibiting approximately a 2_K (50°F) temperature rise per stage at approximately 259 m/s (850 ft/3) tip speed. The axial compressor is designed for low-cost production technology and, as a consequence, has compromised aerodynamic blade configurations. Constant camber, constant twist, and constant blade sections introduce aerodynamic limitations (e.g., via the first stages being set at choke at the tip and stall at the hub, and other blade rows being set aerodynamically at arbitrary incidence).

The axial compressor feeds a centrifugal compressor of relatively low specific speed whose inducer employs a two-stage bladed design planned for manufacture similar to the axial rotor. The radial portion of this design utilizes another low-cost approach wherein blade span and blade numbers are compromised to match the outlet requirement of the axial compressor and overall pressure ratio of 10.3:1 at standard conditions. The radial diffuser is fabricated in accordance with a low-cost concept which regards efficiency and stall margin as design optimization objectives. For fabrication, a shell-molding process involving a no-bake sand molding technique is used.

NASA CR-159603 WRC Report No. 78-113-15 The high-pressure compressor feeds a cosbustor designed for fabrication from basic pressed components. Long-life requirements will be achieved by use of thermal barrier coatings and film-cooling techniques. The first stage turbine nozzle will be of relatively conventional configuration but adaptable to a coated and cooled design to achieve long life at increased performance. The turbine rotors and second stage nozzle are of medium-to-low stress design with somewhat compromised blade shapes adaptable to low-cost production methods.

Aerodynamic limitations include high taper ratio, non-arbitrary twist, and compromised hub and tip incidence angles and loads.

The core is designed to run at relatively low rotational speeds at standard conditions (as encountered in T/P or T/S applications) and induce relatively low stress and cyclic loads on the rotational parts. This allows simplified, econom- ical bearing and shaft designs. The low initial design speed can be increased for T/F performance optimization. Structures peculiar to the core will be fabricated to be compatible with each engine configuration and the fuel control sad starter/generator. These will be designed to allow operation for refinement of aerodynamic, combustion, and mechanical properties. In terms of part costs, the common core components comprise _3.8 percent of the prop and shaft engines and 30 percent of the fan engine.

Preliminarydesignpo_nts for the core co_onents are as follows: T P = 10.3 - Compressor Pressure Rstio r _c = 73 percent - Compressor Efficiency W__= 1.59 kg/s (3.50 Ibm/s) - Airflow n b = 0.99 - Burner Efficiency qt = 0.88 - Turbine Efficlency Significant to the core design is the utilization of a starter/generator compat- ible with all engine configurations. It is anticipated that the starter/genera- tor will be a hybrid permanent magnet motor/generator of a relatively high-speed brush type. The fuel control contemplated will be an electromechanical type employing a zero-pressure-rise pump metering system controlled by an electronic computer. The design would utilize integrated microcircuits and standard microprocessor modules. This control element for the core would be designed to be compatible for functioning as the primary segment of the control for all three engine types, with enough inherent sophistication to enable its adaptation to the range of control functions required for each application.

TURBOPROP ENGINE DESCRIPTION The conceptual design of the T/P engine shown in Figure 37 is basically a 224/298 kW (300/400 hp), flat-rated, fixed-shaft turbine. It is designed for medium performance at very low production cost. The fixed-shaft concept was chosen NASA CR-1_9603 WRC Repert Ro. 78-113-15 o o E-+

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NASA CR-159603 NRC Report No. 78-113-15 because it is approximately one-third lower in cost than free turbine designs and is compatible with lower cost control systems. The constant speed characteristics of engine operation tend to reduce problems of supplying electrical power and bleed air for cabin pressurization. Also, the engine response is fastPr in landing modes due to the fixed-shaft design, and this enhances go-around capabil:ty as compared to free-turbine-powered airplanes. The turboprop's predicted perfor- mance is shown for two levels of component efficiency and two turbine inlet temperature levels in Table XLIII.

The engine consists of a gas generator composed of the compressor and burner elements described previously in the common core account. Two stages are added to the turbine section for power delivery and a simple single-shaft power reduction gearbox is incorporated for propeller drive. The engine is designed to run at low rotational speed and low stress levels, both in the compressor and turbine.

This low rotational speed simplifies the gearbox, which is designed to benefit from low-cost manufacturing processes.

The engine is designed with the inlet at the rear and the exhaust directly behind the propeller reduction gearbox. This geometry eases the problems of foreign object ingestion, induction system icing, distortion, noise, accessory access for service, and installation in single-engine and some twin-engine airplanes.

The compressor blading and drum material will be titanium. Predicted engine weight is 73 kg (160 ibm) without starter/generator but with all other equipment including control system and interstage bleed valve [gO kg (177 Ibm) with starter- generator]. All gears will employ powdered metal fabrication technology.

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o NASA CR-159603 WRC Report No. 78-113-15 TURBOSHAFT ENGINE DESCRIPTION The conceptual desig_ of the T/S engine shown in Figure 38 is a free turbine with the aerodynamics of the gas generator basically identical to that of the common core but with a free turbine driving a simple 6,000 rpm output gearset and a high-speed accessory drive. A compressor bleed valve is provided at the aft end of the axial component to avoid surge problems under part-speed conditions. The fuel control is a version of the common core control with features added for free turbine overspeed protection. The starter/generator would be the same as for the T/P engine with similar performance.

A -10944 Figure 38. Turboshaft Engine The engine was conceptualized for the case where the power turbine runs at close to gas generator speed. Further optimization could change this in favor of a lower speed power turbine to improve the gearing and overall power turbine design. Another version of the engine could use the shaft system from tne T/F engine with a through-shaft gearbox at a modest increase in accessory cost and complexity of construction. Such a design would require a fuel management system similar to the T/F system described on the following pages. Obviously, this configuration would have the characteristic of greater commonality to the T/F at the expense of commonality to the T/P. As a free turbine, the engine would also be less responsive to transient inputs than the fixed shaft engine.

Manufacturing technology methods and processes are expected to be essentially the same as those described previously for the T/P engine construction.

NASA CR-159603 _C Report No. 78-113-15 TURBOFAN _GIME DESCRIPTION Figure 39 illustrates in cross section the selected concept for the T/F engine.

The design utilizes a conventional two-spool shaft configuration with character- istics amenable to envisioned low-coat manufacturing technology, processes, and methods.

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The design points for the T/F would be considerably different if a turbine of approximately 1478°K (2200°F) were developed and a higher bypass fan of approxi- mately the same fan pressure ratio were substituted. Tables XLIV and XLV summarize the performance improvement that could be achieved with the higher temperature design point and with component efficiency improvements.

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NASA CR-159603 WRC Repo_c No. 78-113-15 TABLE XLV. LOW COST TURBOFAN PERFORMANCE COMPARISON - 9144 m (30,000 Ft)/ Mn 0.6/Standard Day Nominal Improved Component Nominal Improved Efficiency Turbine Inlet Nominal Nominal Improved ImproveO 1333°K 1478°K 1478°K 1333°K Temperature (2,200°F) (2,2Oh°F) (1,940°F) (1,940°F) Uninstalled 8.0 4.80 4.89 8.0 BPR 27.4 31.2 OPR 23.2 25.0 Engine Inlet 15.06 9.08 9.93 13.29 Airflow-kg/s (29.3) (33.2)

(lbmls) (20.01) (21.9)

1.0 1. 167 1.252 1.524 Relative Fn 0.854 Relative TSFC 1.0 0.941 0.913 Installed 8.3 8.23 BPR 4.89 4.97 24.1 25.8 29.6 OPR 22.3 Engine Inlet 8.98 9.84 13.06 14.88 Airflow-kg/s (21.7) (28.8) (32.8) (Ibm/s) (19.8) 1.120 1.161 1.433 Relative Fn 0.950 1.028 0.960 0.943 0.874 Relative TSFC TURBOFAN COHPONENTDESCRIPTION The T/F engine design employs a fan that operates to an approximate 305 m/s (1,000 ftfs) tip speed and a pressure ratio of 1.4 under standard conditions.

The fan is attached to, and followed by, a three-stage intermediate pressure compressor producing about a 28°K (50aF) temperature rise per stage. This modest temperature rise enables the intermediate pressure compressor to be designed in accordance with the low-cost construction concepts intrinsic to the NASA CR-159603 NRC Report No. 78-113-15 o.

,w technology program. The IP compressor and fan are driven by a four-stage, low- speed turbine also based on low-cost construction concepts.

The high-pressure spool is basically derived from the common core elements de* scribed previously. Provision has been made for surge protection via a flow control device between the axial and centrifugal compressor components.

me engine accessories are arranged around the waist formed by the axial compres- sor rotor of the core. Hajor fuel control components and the starter/generator are common to the core design. Low-cost construction methods utilizing compro- mised aerodynamic shapes and low-speed :omponents are used throughout the design.

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NASA CR-159603 NRC Report _Io. 78-113-15 SECTION 7 TECHNOLOGY PROGRAH PlAN This Technology Program Plan presents an approach to a research, design, and development program to provide iow-cost turbine power for general aviation in the next decade. The plan is intended to demonstrate an orderly and logical process culminating in the FAA type certification of turboprop (T/P), turboshaft (T/S), and turbofan (T/F) engines linked by a conmon core. The results of life cycle cost (LCC) studies presented herein indicate that turbine ownership in general aviation can produce significant economies over piston engine ownership and serve to reduce the depletion rate of world petroleum reserves. Engine design concepts baked on a common core for economy are feasible and exhibit the potential for future growth in performance and fuel efficiency with advancing technology. The combination of these concepts with realistic program plannin$ and systems engineering control offers promise that implementation will result in successful achievement of the stated objectives.

This plan is arranged in two-page displays presenting textual descriptions on each left-hand page and supporting graphics on each right-hand page. A sugary of the total plan is provided on the following pages. The plan is sequenced to present develol_ent r_tionale and schedules (pages 136 through 151) followed by program cost projections (page 152). The scheduling is presented in two levels. The first level is a program overview supported by another level of detail for preliminary design, common core development, the development of each of the three engine types, and additional detail for core COll_nent and turbofan engine component development.

The second level schedules were generated by iterating tasks against the major silestonesuntil realistic and viable detailed plans resulted.

NASA CR-159603 WRC Report No. 78-113-15 PROGRAN PIAH SI_ARY Design concepts for the GATE technology program (detailed in Section 6) are based on low-cost approaches for component construction and the development of a comBon core adaptable as the nucleus of the T/P, T/S, and T/F engines. Basic accessorxes (e.g. the starter/generator) will be compatible with all three engines. Low-cost construction techniques to be used require aerodynamic compromises in the designs based on blade configuration constraints, but benefit from low operating temper- atures, modest temperature rises per stage, low pressures, and low rotational speeds at standard conditions.

The common core is composed of a six-stage axial compressor feeding a single-stage centrifugal compressor. The high-pressure compressor feeds a combustor of simple construction employing thermal barrier coatings and film-cooling techniques Turbine rotors and the second stage nozzle are of medium, low-stress design with aerodynamic limitations due to high taper ratio, non-arbitrary twist, and compro- mised hub and tip incidence angles and blade span loading. In terms of part costs, colmon core components comprise 43.8 percent of the propeller and shaft engines and 30 percent of the fan engine. Common core development details are given on pages 140 and 141.

Planning for development of the colnon core and the propeller, shaft, and fan engines is displayed in a series of eight schedules with accompanying descriptions (pages 137 through 151). This display begins with a summary-level master schedule and progresses with coverage of the preliminary designprogram, common core develop- nent, and each of the engine development programs. Additional detail is provided for conmoncore component and fan engine coRponent develol_ent.

The nmster schedule provides an overview of the sequencing and logic for the ll- year tern of the overall program. This level of planning illustrates a concentra- tion on aerodynamic and thermodynamic analysis combined with extensive manufactur- ing technology investigations involving both the common core and engine development efforts. Also, considerable development test hours will be dedicated to refinement of fabricated hardware (3,000 hours on the common core, 17,000 hours on the pro- pellet engine, 9,000 hours on the shaft engine, and 6,500 hours on the ian engine).

The entire effort will be monitored and controlled by an in-depth systems engineer- ing activity against the LCC discipline and performance requirements imposed on each engine type.

Preliminary design activity forms the basis for departure into the development activity. Preliminary design will build on the conceptual approaches presented in this plan and will result in initial designs which will circumscribe the perfor- mance requirements and development parameters for control of the development pro- grams.

Common core development will be influenced by early design activity on each of the engines to allow the best mix of core components for optimum engine performance at low cost. However, once the common core component designs have been selected, the core will be a driving factor in all subsequent engine design and development work.

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Each of the engines will undergo essentially the same kind of development proce- dure. Based on common-core and refined-component hardware, engine designs will go through a series of three release cycles. The first release will result in fabri- cation of hardware for development tests; the second will result in hardware for endurance tests leading to preliminary flight rating (PFRT) qualification; and the third release will provide hardware for endurance, environmental, and qualification tests leading to FAA certification. A technical manual activity is also included to provide operation and maintenance data for support of flight tests and certifi- cation.

A twenty-year, turbine-fleet cost benefit summary is shown on pages 103 and 108.

This summary compares projected engine-related ownership costs for single- and twin-turboprop and twin-jet airplanes with comparable piston engine related costs.

Note that a total savings of more than $3.52 billion is possible through fleet turhinization. Data on page 152 shows that a GATE Technology Program investment of about $0.11 billion (1,261.5 man-years of effort and $48,619,000 material dollars) is required to enable the $3.52 billion savings. More on the LCC benefit and GATE Q° program cost subject is given in Section 5.

# 4o NASA CR-159603 v'__C Report No. 78-113-15 PROGRAM OVERVIEW The Master Schedule displayed in Figure 40 portrays the logical and sequential development of the three types of turbine engines based on a common core design effort. This schedule provides a broad overview of how the GATE program stimulus can lead to successful delivery of certified engines which will meet the demands of general aviation in the late 1980's. Subsequent schedules in this plan will serve to illustrate the program planning in greater detail.

Major emphasis has been placed upon economy of operation and producibility. This goal can only be met by a carefully executed program of aerodynamic/thermodynamic analysis, manufacturing technology development, and exhaustive testing, with design iteration controlled by a pervasive systems engineering discipline.

Development costs for this program will represent a substantial investment and the importance of a high level of planning and control cannot be overemphasized.

Significant in this planning are 3,000 hours of development testing of the core design to establish a baseline for the three engine development efforts. Prior to completing this test activity, each of the three engine designs will have influenced the developing core design. However, once conq_lete, the core design will bea driving influence on the engine designs.

Concurrent with core development, manufacturing technology imvestigations will be undertaken to establish optinmm methods and materials for development of the propeller engine gearbox. The evolved gearbox design will be phased into the propeller engine design along with control and accessory designs, and the engine will undergo approximately 17,000 hours of development testing. Followin& PFRT, the development engine will undergo tests leading to FAA certification and delivery.

The shaft engine will undergo essentially the same development cycle as the pro- peller engine but with approximately 9,000 hours of development test activity.

Much of the accessory development and test data derived from the propeller engine program will be available for refinement of both the shaft and fan engines.

The shaft engine fuel control is expected to be applicable to the fan engine because of the similar free turbine characteristics. Additional development of fan components will be undertaken to ensure optimum matching under operational conditions. Subsequent sections will deal with the details of development and matching of the fan, axial compressor, turbine, burner, and core.

The overall GATE technology program presented is based upon intensive investi- gations and the resulting concepts developed during the conceptual design phase.

Activity during the study phase reported here has shown the design approaches presented and the goals of achieving economical production and operation to be feasible during the prescribed time period.

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L'U_IL1T_ OP TI-IP, NASA CR-159603 WRC Report No. 78-113-15 PRELIMINARY DESIGN TO REFINE CONCEPTS FOR TECHNOLOGICAL DEVELOPMENT The technology program will begin with a 9-month effort to refine basic core and engine concepts into sufficient definition to allow commitment of development effort. This preliminary design activity will serve to focus subsequent design and development efforts so that the program can be controlled and directed toward a unified goal. This activity will result in a preliminary design report in the tenth month which will serve as the initial analytical baseline for the program.

There will be no intent to limit design flexibility or to ignore technological breakthroughs or significant advances in the industry state of the art. When these occur, they will be investigated and utilized when appropriate. Planned iterations in every phase of the prcgram will accommodate these actions.

Preliminary design will begin with cycle analyses and design optimization studies of the three engines. Performance sensitivity analyses will be made and aero- dynamic flow paths described. Design layouts will be initiated for the core, the three engines, and the engine controls and accessories. These will be subjected to mechanical, maintenance, and safety analyses. Critical components will be defined and initial aerodynamic flow paths and performance data will be updated.

Initial engine layouts will be used to drive the core design. As the core design evolves, it will be used to iterate the three engiue designs toward final pre- liminary configurations. This data will be reviewed at the end of the sixth month and decisions made for update of the preliminary designs in each area.

Systems engineering will perform an independent audit of this activity throughout and will define requirements and develop specifications to integrate the analytical and design activities. LCC studies, maintenance and safety requirements, comnon- ality considerations, and initial specification development will be used to impose design requirements.

Major emphasis will be placed upon core and core component design. The core design layout as described by the sixth month will undergo manufacturing technology investigations and core component aerodynamic/thermodynamic analyses continuing into the subsequent development phase. This activity, through the ninth month, will be _eflected in the preliminarydesign report.

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o o l-i -'-:, =. :.UCV3ILrI'Y OP THE ._ .,,N.,_L PAGE IS POOR NASA CR-159603 NRC Report No.-78-I13-15 CO_R)N CORE DEVELOPHEN_ The development of a common core for the three engines is driven by the objectives of low cost for each component. Durability and optimum performance for the hot sections of the design will be given prime consideration. As aerodynamics and thermodynamics are obviously constrained by the requirement for low-cost, careful matching of speed, temperature and pressure is indicated. These factors will be aided by intensive investigation and development of manufacturing technology and p:ocess/procedures refinement. As the design progresses, redesign and upgrading of core hardware will be accomplished based on test results.

Basic to development of low-cost designs are the evolution of an economical burner and approaches to bearing and suspension of the high-speed shaft. Hate- rials, coatings, processes, shapes, and cooling concepts will undergo iterative investigation and development as shown by the schedule. The preliminary design baseline will lead into cooperative aerothermodynamic analysts, manufacturing technology development, and turbine/compressor/burner development. These will define the core design activity leading to approximately 3,000 hours of development testing. The compressors, burner, turbine, shafts, bearings, and accessories will be matched during this activity to form the nucleus for the subsequent engine development work. The L_ots planned for the core wlll be unpressurized but will yield data sufficient to harden the designs for use in the engine build- ups. Details of these activities are given on the following pages.

Systems engineering activity during this time period will include audit and control of the core program and continuing LCC study and refinement, airframe integration studies, propeller and control interface and design requirements, and the identification and update of propeller, shaft, and fan engine specification content.

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}'" ;.-NAL PAGP_ l,q p v ,_ NASA CR-159603 WgC Report No. 78-113-15 CORE COHPOk'ENT DEVELOPHEICr The axial compressor, centrifugal compressor, and turbine will undergo concurrent investigations for advanced manufacturin& technology. Control design and the manufacturing technology approaches will continue as initiated during preliminary design. Tooling, methods, and process development will be undertaken as shown on the schedule. These activities will interact with alloy and material selection, and research of design properties, resulting in the development and buildup 9f prototype parts for integration into the common core for testing. Prior to release to the common core, each component will go through appropriate fabrication tests, nondestructive investigation, spin tests, and dimensional checks. The manufacturing processes will be refined and designs coordinated with engineering.

Final-configuration fabricated parts will be available for propeller engine test at the end of the core test serles.

Aerodynamic and thermodynamic testing to establish airflow patterns and heat and stress characteristics will take place in concert with manufacturing technology activity. One fabrication/test cycle is planned for the burner and two cycles are planned for the turbine and compressor. As these test-redesign-test activities are completed, results will be fed into the common core design and to the manufac- turing technology activity, as required to maintain the best overall approach.

The common core design will progress based on inputs from the manufacturing and aerothermodynamic efforts. Fabrication and procurement of refined designs (inclu- ding requisite controls and accessories) will be undertaken to assemble hardware for core tests. Hardware from these tests will be fed to the engine development efforts and design updates will be maintained. Aerothermodynamic analysis and manufacturing technology development will undergo concurrent refinement integrated by systems engineering control. The results of the two activities will be super- imposed on the second design iteration for final proof parts fabrication and test.

A fully-integrated effort between design engineering, manufacturing technology, and aerothermodynamic analysis will be maintained by project engineering control throughout the process. LCC studies will continue to impose design and development constraints toward the prime objective of low-cost final hardware. Maintain- ability, reliability, durability, and safety data will be taken from all tests and will be considered in the evolving designs and controlling specifications.

Systems engineering will update LCC and design requirements as required to maintain an integrated overall program.

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]4] NASA CR-159603 WRC Report No. 78-113-15 PROPELLER ENGINE DEVELOPMENT Shortly after the start of common core development, manufacturing technology investigations into the propeller engine gearbox and components will commence.

Also, based on common core work, the development of engine accessories, including a fuel control, can begin. These development activities will continue until a low-cost fuel control and other engine accessories are defined and the technology is released to permit the gearbox to be designed. The gearbox will be fabricated and undergo independent tests concurrent with engine buildup. Following engine accessory definition, aircraft accessory development will commence. The gearbox, controls, and accessories data along with inputs from the common core development effort will feed the initial engine design.

Final results from the gearbox qualification tests will be fed into the second design iteration for integration with engine development tests. These tests will continue and will provide data for design refinement. A next generation of gearbox, controls, and accessories will be fabricated for an endurance test leading to PFRT. Data from this testing will serve to further refine the engine design. Data from both the development test and the endurance test will be used to update the engine design including the gearbox design. A final design release of the complete engine will permit fabrication of hardware for the final endurance, environmental, and qualification tests leading to engine certification. A total of approximately 17,000 hours of development testing are planned.

During the final test phase, assembly procedures and test, operation, and mainten- ance data and analyses will be assembled into manuals for support of the engine.

These manuals will be evaluated during PFRT. Data from flight tests following PFRT will also be used for manual update and design refinement during this period. Six months following certification, final manuals and the first certified engine will be delivered.

As in all other phases of this program, project engineering will coordinate the development activity, and systems engineering will control the integration.

Propeller engine specifications will be updated and released to support each development milestone. The propeller interface will be released at the start of development activity. Propeller and propeller-control specifications will be issued and the final turboprop engine specification released at PFRT concurrent with initiation of final tests.

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_=1 w oO Wt_ s _° _u >- o M _< S31aO£S3_V _N3 I_ -.._ = U _0 >" W v - ]45 NASA CR-159603 WRC Report No. 78-113-15 SHAFT ENGINE DEVELOPNE/_r Design inputs from the common core development activity and the work on the propeller engine will initiate design of the shaft engine. The initial design activity (basically the preliminary design updated to accommodate changes in the common core and propeller engine requirements) will provide a baseline for gearbox and fuel control development. The common core used for the propeller engine will have a free turbine section added and require additional fuel control develop- ment. The fuel control development and bench test would apply to both shaft and fan engine concepts. Manufacturing technology and engineering analysis previously undertaken for the propeller engine will be applied to the design for the shaft engine. A gearbox design will be required to reduce free turbine speed to shaft drive speed. Once the gearbox and fuel control are ready for testing, engine design will resume based upon those configurations. The gearbox design, although differing from the Fropeller gearbox design, will employ much of the same manufac- turing technology. Gearbox qualification testing will be oriented toward englne design-matching and compatibility with subsequent engine development test hard- ware. Test data will be injected into the design process, and fabrication of a shaft engine will be started for development testing. Data from this testing and from the continuing fuel control rig tests will be fed into the design update activity, leading to release of engine drawings for parts fabrication and subse- quent engine endurance tests for PFRT. A total of 9,000 hours of development tests are planned for the shaft engine.

Data from the PFRT endurance test will affect the final design release for fabrication of hardware for endurance, enviromsental, and qualification testing.

This testing will lead to certification. The first issue of shaft engine manuals will be available at PFRT for support of flight tests. These manuals will be updated based on results of flight tests and will be validated during the tests leading to FAA certification. Production-released drawings, specifications, and manuals will be available concurrent with delivery of the first certified engine.

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NASA C_-159603 WRC Report No. 78-113-15 FAN ENGINE DEVELOPMENT Preliminary design work and aerothermodynamic analysis, combined with results of the common core development activity, will define the baseline for start of the fan engine design. The culmination of this initial design work will coincide with the start of the fuel control development and rig test which were previously described as applicable to both the shaft and fan engines. Also beginning concur- rently will be manufacturing ,echnology investigations for the low-pressure spool design and a parallel =ffort to develop the spool. Component development efforts are detailed on the following pages.

Burner improvements and core updating will also be undertaken before resumption of engine design activity leading to fabrication release for the first engine development test. The fuel control rig test results and the manufacturing tech- nology work on the low-pressure spool will be fed into another design iteration which wiU also benefit from development test data. This design will be released for fabrication of hardware for the endurance test leading to PI_T. PFRT hardware will be delivered for flight test. Approximate\y 6,500 hours of development tests are planned.

Design activity will continue with final test d-'.a from the development test and interim data from the endurance test used to define hardware for the endurance, environmental, and qualification test series leading to engine certification. As in the previous two engine development programs, engineering data and manuals will beavailable for support and deliverywith the first certified engine.

The technology program displayed on the preceding series of schedules shows an orderly process for the design, development, test, and delivery of high-perfor- mance, low-cost engines meeting the requirements of general avietion. The iterative development process described, based on a common core design with maximum control and accessory similarity, enhances the probability of meeting the goals of the program.

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,.,I NASA CR-159603 NRC Report No. 78-113-15 FAN ENGINE COMPONENT DEVELOPMENT The development of fan engine components is based on core development, manufac- turing technology, and aerothermodynamic investigations. This additional level of detail indicates the emphasis that will be placed on the design and matching of the components for efficient fan engine performance. Also shown is the sequence and logic of the development superimposed on the engine development cycle.

Beginning with the completion of the initial fan engine design phase, component design will commence on the fan and turbine. Layout and design of the fan will be aided by flow path analysis activities, blade design, stress and dynamic analysis, and performance analysis. This design will result in hardware fabrication and testing for two cycles. The fan design will then be updated and the hardware will be subjected to testing on a spin/shake rig. Completion of this activity is coordinated with the required fan introduction into the engine design effort.

Initial turbine design will supply configurations for tooling, process, and methods development. Three base designs will be fabricated and subjected to test.

This test activity will result in the selection of one optimum design which will be updated based on manufacturing development work, fabricated, and tested again.

Data from this test will be fed to engine design and to process refinement activity for the development of engine parts for engine fabrication and testing. Update of the turbine designwill continue as engine test data is received.

The axial compressor and burner will be subjected to design investigation to allow fabrication and one test cycle each. Data from these component tests will be reflected in updated designs to allow engine buildup and testing. Data from the common core development activity will be used to update axial compressor process development.

Core design matching for the fanjet engine will also be undertaken and the updated core fabricated and tested. Modifications from the common core are required because of the close aerodynamic coupling of the high- and low-pressure spool sections. The modified core, along with the other developed components, will result in the buildup and testing of a carefully matched, balanced, and efficient fan engine.

i NASA CR-159603 WRC Report No. 78-113-15 ]: W_J .c o W_ or} tP_ ¢p o _p _J 4J _p o o f_ REPRODUCIBILITY OF T!t L ORIGINAl, PAGE IS P()_;]t NASA CR-159603 WRC Report No. 78-113-15 PROJECTED PROGRAM COSTS The Technology Program Plan described herein is estimated at 1,261.5 manyears and $48,619,000 material dollars (1978 economics). These projections are based upon experience with similar development activities. Cost estimates were collected (based on the specific activities presented in this program plan) from the func- tional work groups that will be required to participate. These data were reviewed and revised by management to present the most realistic budgetary and planning information possible.

Costs and material dollars as shown in Figure 48 are distributed according to basic hardware elements. It is felt that this illustration provides the best indication of projeLt_d requirements at this time. However, it in no way depicts a work breakdown structure, cost accounting scheme, or any other method that may be required for cc_t c, llection and control. The illustration is provided solely for future planning pucp,_es. • The manyear data displayed is based on inputs from Design, Systems Engineering, Aerothermodynamic Analysis, Test Operations, Fabrication and Assembly, Logistics, Design Assurance, and support organizations. Material dollars represent direct costs for hardware and services projected as required for the program.

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ENGINE ENGINE CORE ] l PROPELLERENGINE ] 21.5 MANYEARS 136 MANYEARS 471 MANYEARS 241 MANYEARS 392 MANYEARS PLUS _152, 000 PLUS _2,831, 000 PLUS _16, 5|4, 000 PLUS SIt, 841, 000 PLUS _|7, 281,000 TOTALS - |, 26 1.5 _vlANYEARS PLUS _ 48,619, 000 A.-9555 Figure 48. Projected GATE Program Cost Estimates

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NASA CR-159603 WRC Report No. 78-113-15

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SECTION 8 I CONCLUSIONS AND RECOMMENDATIONS The following conclusions resulted from the Market Analysis, Trade Study, and Common Core Evaluation: I. A modest growth in annual aircraft unit sales is predicted for the decade ahead. This will be accompanied by a steadily growing fleet, because fleet additions will exceed retirements by a factor of about five to one. Most new aircraft sold will be for business use.

2. Because of fleet growth and characteristic weather influences, there will be a crowding in the flyable airspace below 3810 m (12,500 ft) that will influence buyer preferences toward pressurized aircraft that can operate in a greater volume of airspace.

3. Along with the need for pressurization will be a requirement for higher climb rates to facilitate flight at the higher altitudes, and higher cruise speeds to minimize the influence of the strong, high-altitude headwinds normally encountered by westbound flights. Deice/anti-ice capability will also be required.

4. There will be a demand for down-sized, turbine-powered business air- planes due to the high cost of fuel. These new airplanes will have adequate, but minimum, seating capacity (four to six seats), and they will be designed to be easy to fly with minimum crew requirements.

5. Flat-rated T/P engines in the 134 to 261 kW (180 to 350 hp) range and T/F engines producing about 4448 N (1000 lbf) of thrust will play an important role if fuel efficiency can be improved and cost constraints eased. Bleed air and power extraction requirements will exert an important influence on engine design.

6. The largest market will be for T/P engines because they can be more easily adapted to FAA-certified, single- and multi-engine, new-production and previously-owned airplanes, airplanes that will be better suited for small airport operations than turbofan-powered craft. There will be many more pressur- ized and icing-certified candidate airplanes for conversion in 1988 than today.

7. Because of the substantially larger market for T/P engines, if core design compromises are necessary, the compromises should be to the advantage of the T/P to assure its acceptance.

8. By judicious core design, it is probably possible to provide enough flexibility in compressor geometry and shaft speed to enable a common core to be used as an optimum T/P engine component as well as a component for the high- pressure section of a T/F.

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WRC Report No. 78-II3-15 9. A fixed-shaft, constant-speed T/P is preferabl_, to _ variable-speed, free-turbine design because it can be produced at a lesser cost, is compatible with lower cost control systems, is better able to supply bleed air for cabin pressuri- zation during letdowns from high altitude, and enhances airplane go-around capa- bility from aborted landings.

10. A concentric-shaft, t_:o-spool T/F engine with the high-pressure spool derived from common core elen:cnts is the least costly ot the T/F designs consi- dered. Its performance potential is excellent, and it is lighter in weight than the second-choice tandem-spool contender. Other contenders included two geared- fan, FI07 derivatives.

11. A free-turbi.e T/S ongine design that uses the common core as the gas generator and the T/P fuel control (with features added for free-turbine overspeed protection) was selected because of cost and commonality benefits.

12. An innovative approach to the design and manufacture of small turbine engines is required to meet cost goals. The approach should not involve plans to run the engine hotter and faster, but rather slower and cooler. The advantages from the resulting stress reduction should be exploited.

13. Cost savings are possible through the augmented use of powdered metal- lurgy and non-precision casting techniques, and as a result of the geometric con- straint of rotating components. Low-cost, multiple-blade-row axial compressors can be manufactured, for example, if the blading has uniform twist and constant camber, chord, and cross section.

14. Research oriented toward improving the durability of engine components is required in areas involving conductive cooling with thermal barrier coatings, coatings for environmental protection, hot isostatic pressing, etc.

15. The cost of engine accessories (fuel control, starter generator and propeller control) must be reduced and the durability improved. Accessories alone can equal the cost of a piston engine.

16. The productivity of the turbine-powered conceptual airplanes investigated that had piston-powered counterparts was competitive in terms of seat-km/1 (seat- nm/gal) when flown at high altitude with the counterpart piston airplane. The performance improvement due to turbine engine substitution was remarkable in each case.

17. Although the candidate turbine engines were more costly than the piston counterparts, engine and airplane LCC were competitive because of reduced airframe, propeller, and engine maintenance costs, and because of the price difference be- tween Jet A fuel and 190-octane aviation gas.

18. Calculated C£M prices for the T/P, T/S, and T/F engines in 1978 dollars were $19,515, $26,163, ,,nd $25,352, respectively, excluding product liability influences.

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NASA CR-159603 WRC Report ho. 78-113-15 19. Turboprop, turboshaft, and turbofan development programs based on a common core would require a minimum total investment of $0.1i billion before each engine type could be FAA-certified. Subsequent engine procurement and use over a 20-year period should result in a user savings of more than $3.52 billion _hen compared to the continued use of current-technology piston engines. Substantial fuel savings [up to 4.9 billion liters (1.3 billion gallons)] could be realized through turbine engine component efficiency and temperature-tolerance improvements.

The investment cost would be 46 percent higher than for the nominal engines, however.

20. Because of the substantial investment required and the considerable development risk, the small turbine engine will probably continue to elude the small airplane without Government technology support. These small airplanes ore already twice as productive as the newest airliners from a seat-km/1 (seat-nm/gal) standpoint, but need expanded operational capabilities for added utility, comfort and, especially, safety. The small turbine engine offers promise in all these categories.

RECOMMENDATIONS Turbine engine manufacturers interested in the small general aviation engine market have ongoing technology programs to advance the state of the art. For the most part, these are at a low level of activity because of the risk/return situation and available funds. These programs must be accelerated to produce more immediate results. This study report has defined the proper content of a Government- sponsored program to accelerate this activity through a blend of analytical and experimental work. Now is the time to start the hardware-oriented research (which can provide greater returns) and Government support for this effort is strongly recommended.

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NASA CR-159603 WRC Report No. 78-113-15 APPENDIX A KEY INFLUENCES ON A 1988 GENERAL AVIATION MARKET SCENARIO The general aviation market is very sensitive to a number of factors that could cause any long-range forecast to deviate far from the mark. Con;equently, the assumption has been made that market perturbations will be evolutionary, not revolutionary. In this respect, it is assumed that there will be no major global conflict, no breakdown of the U.S. monetary system, no depression, no energy depletion, etc. Rather, brushfire wars in distant lands, moderate infla- tion (about 7 percent), moderate unemployment (averaging about 5.2 percent), continued high interest rates for airpiane financing (II.0 to 11.5 percentJ, and increasing aviation fuel costs (due to inflation, taxation and scarcity) can be expected. On the basis of the evolutionary perturbation assumption, a modest general aviation market growth is foreseen for the decade ahead.

The prediction of modest growth can mean many things to many people, and it would be well to put this term in perspective. The confusion comes with respect to the index used for growth measurement. Two popularly used indices are annual unit production and annual dollar sales. Other indices include fleet size, aircraft operations logged by FAA control towers, flight services logged by flight service stations, and total hours flown by general aviation airplanes. It is possible to have a growing fleet and operations at record levels while new airplane sales are declining drastically. This occurs because aircraft retlre- merits, which will number about 2,000 in 1978, fall well short of the number ot new aircraft being added to the fieet, even during a very poor sales year. A Frost and Sullivan estimate (ref I) of net additions to the total active general aviation fleet in the United States is shown in Table XLVI.

TABLE XLVI. FROST AN_ SULLIVAN ESTIMATE OF AIRCRAFY ._DDED TO ACTIVE US FLEET 1974 1975 [976 1977 1978 1q79 1980 lq8l 1982 [ 1983 Factor3 Shipped 14,166 14,270 1.5,Olb 16,186 17,754 1q,|87 20,324 [ 21,489 22,785 24,13_ US Harket 8,499 8,562 9 7bO 10,521 12,428 1t,_,31 f 14,227 t 15,042 15,9_)0 1h,8')4 Sh ipl)ed to I Fo re i gn t ] Market 5,667 5,708 5,256 5,b65 5,32b 5,756 6,097 [ t,447 6,835 7,240 !

Retired 1,449 I .562 I ,760 2,000 2,000 2,400 2,6()0 3,000 i 3,300 _,500

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Ne2 Added to Fle,,t 7,000 7,000 8.000 f, 521 10,428 1t,031 1i.b27 12,042 ! 12,650 1 _._94

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Tota 1 I Activ, ] t Fleet 152,950 159,o50 1t67,950 176,471 186,899 198,330 209.957 221,999 i;'_'t_9 ,:48,(}43

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NASA CR-159603 _RC Report No. 78-113-15 The prediction of modest growth for the decade ahead is made here with reference to new airplane and new engine unit sales for domestic and foreign markets by U.S. manufacturers. The pace of the growth will be governed by the energy situation, airport/airways development, the regulatory environment, economic regulations, new technology, and product liability considerations. The paragraphs that follow discuss each of these influences in relation to a 1988 sce,ario for general aviation powerplants including those for rotary wing aircraft.

ENERGY IN-FLUENCES The 1973/74 fuel crisis and the attendant Emergency Petroleum Allocation Act had a decided influence on general aviation that set the stage of activity for years to come. The full effects of the crisis were not obvious because of cross influences and the limited duration of the immediate problem. Nevertheless, the seriousness of the situation and the potential consequences of a recurre,ice has caused general aviation industry leaders to focus on the fuel economy of their present and planned products. Table XLVII lists observations of the effects of the 1974 fuel shortage on aircraft operations and sales. Some of the actions initiated to conserve fuel are listed in Table XLVIiI. Additional actions which will probably be implemented with the increasing preciousness of fuel are listed in Table XLIX.

AIRPORT/AIRWAYS DEVELOPHENT During the late 1960's, airport/airways development fell far short of what was required for the safe and expeditious handling of air traffic. The result was a series of air traffic controller slowdowns (traffic was handled "by the book"), traffic backups, and long holds on the ground and in flight, with attendant fuel wastage. The effect on general aviation fixed-wing aircraft sales was devasta- ting, with sales falling from more than 11,000 in 1966 to less than 5,000 in 1971. Wichita, indeed, became an economically depressed area, and corrective action was clearly called for.

To reverse the downward trend in sales, the general aviation industry, in concert with the air transport industry, recommended legislative action leading to the establishment of an Airport/Airways Trust Fund. Monies for the Fund would come from excise taxes on aircraft tires and tubes, aviation gasoline, airline tickets, international head taxes, waybills, aircraft registration fees, and aircraft weight taxes. Proceeds from the Trust Fund were to be used to provide for the expansion and improvement of the nation's airport/airways system.

After the successful enactment of Trust Fund legislation into public law in may 1970, the job of airport/airways system upgrading was begun. As progress was made and the productivity of the National Aviation System increased, aircraft sales improved until today a more than 18,000 unit sales year is predicted for 1979. In December 1978 the Trust Fund balance was nearly 4 billion dollars.* _-_r_he Weekly of Business Aviation, S March 1979, page 79.

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NASA CR-159603 k WRC Report No. 78-113-15 TABLE XLVII. IMPACT OF THE 1974 FUEL CRISIS AND EHERGENCY PETROLEUM ALLOCATION ACT ON GENEPAL AVIATION • Multi-engine piston airplane sales declined during the 1975 and 1976 fiscal years while single-engine piston airplane sales increased.

• Turbofan/turbojet airplane sales leveled in FY'75 and declined in FY'76, reversing the strong growth trend of prior years. T/P airplane sales turned downward during FY'75 and then resumed an upward trend. The decline in FY'75 was due to a slowdown in engine deliveries resulting from an engine supplier strike. There was no lessening of demand for turboprops.

• Because of the imposition of the national 88.5 lun/h (55 mph) speed limit on highways and motorist fuel acquisition problems, aircraft sales remained strong despite the fuel crisis and a sagging economy.

• Aviation fuel was allocated to fixed base operators on the basis of prior year sales, thereby creating spotty shortages. Preferential treatment in fuel dispensation was given to regular customers and locally-based aircraft Itinerant aircraft experienced fuel acquisition probl:_ms.

• Flights had to make more fuel stops to obtain adequate gallonage, thus wasting fuel during descents, holding, approach, landing, taxi, takeoff and climbout.

• Local flying continued at a high level as did itinerant operations within the round-trip capability of aircraft.

• Long-distance flying, where aircraft are the most fuel-efficient, was curtailed.

• Flights were made with inadequate fuel reserve, and safety was l compromised.

• Airplanes had to be left at destination airports for several weeks until they could be refueled. This required passenger and crew shuffling by

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other modes of transportation, which resulted in a waste of fuel.

• Condensation of water in empty fuel tanks created potential in-flight engine stoppage problems.

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NASA C_-159603 WRC Report No. 78-113-15 e_ TABLE XLVIII. FUEL CONSERVATION ACTIONS INITIATED DURING 1974 FUEL CRISIS • Aircraft manufacturers initiated airplane drag reduction programs.

• Business jet redesign activity was focused on the substitution of turbo- fans for turbojets.

• Flight profiles were optimized and more use was made of area navigation (RNAV) equipment for making direct flights.

• Air traffic control issued corner-cutting and time-saving clearances.

• Traffic flow control procedures and airport quota systems were implemented by FAA to improve the flow of air traffic into congested airports and reduce delays.

• Training and proficiency missions were conducted on deadhead or position- ing flights. Flights were consolidated where possible. Unnecessary and nonproductive flights were curtailed.

• A fuel reservation plan was initiated whereby aircraft operators could telephone ahead to a stopover point and reserve fuel before leaving on a flight.

• Engine use for ground operations was curtailed through tows from hangar to ramp, use of fewer engines for taxi, hold-taking at the gate with engines off, and engine shutdowns during short stopovers.

Vexation in the aviation community has been caused by the accumulating Trust Fund surplus and executive attempts to divert large sums for other purposes (e.g., FAA operating expenses, urban mass transit, etc.). These attempts have been challenged successfully to date on the basis of not being in accordance with the intent of Congress in establishing the Trust Fund. Also, considerable system upgrading is still in order and needed to ensure continued safe and orderly handling of air traffic and minimization of fuel wastage. No other expenditures, including those for improved engine fuel efficiency, will make as significant an impact on the efficient use of aviation fuel as those for airport and air traffic control improvements. Airplanes simply cannot be parked at the "side-of-the-road" when there are airborne traffic jams, and considerable power and fuel are used to sustain flight while in a holding pattern.

_m NASA CR-159603 WRC Report No. 78-113-15 Qe TABLE XLIX. ADDITIONAL ACTIONS TO BE EXPECTED WITH THE INCREASING PRECIOUSNESS OF FUEL % • Durxng the new design phase, trade-offs of airplane aerodynamic efficiency versus production cost wiil be weighted more heavily on the side of aerodynamic efficiency.

• Allow_ather capability will be emphasized in downsized, pressurized, turbine-powered aircraft. Small, high-performance T/P and T/F airplanes sized to the passenger load requirements of business will emerge on the market.

General aviation airplane cruise speeds will continue to increase, facilitating terminal area traffic flow and yielding large-airplane fuel savings.

Airplane flying ease will be emphasized so that the businessman can fly himself and his associates to business meetings in a minimum-size airplane. A highly automated, simple-to-use, efficient air traffic coatrol system will play an important role.

• Area navigation equipment will proliferate and its use will become the norm. Airways use will decline.

• Pilot training curricula will evolve toward a greater use of simulators and away from in-flight activities. Procedural training including navigation, instrument flying techniques, and operations within the air traffic control system will be emphasized, with the simulator playing an increasingly important role. In-flight maneuvers having little training value, such as lazy eights and chandelles, will be deleted from the curricula.

• Instrument proficiency and currency will be maintained through increas- ing simulator use.

Turbine fuel specifications will be relaxed to allow more fuel to be obtained from a barrel of oil. Due recognition of this eventuality will be reflected in new engine designs.

1hi NASA CR-159603 _rRC Report No. 78-113-15 The 8eneral aviation market prediction for the late 1980's assumes the wise use of Trust Fund proceeds for the benefit of aviation through the timely imple- mentation of a well-conceived National Aviation System Plan, a plan that is responsive to new technology developments with respect to general aviation airplanes, engines, and avionics.

THE REGULATORY ENVIRON}/ENT As the airport/airways system was upgraded to increase productivity, a need surfaced for new and upgraded airborne equipment and for periodic equipment checks for accuracy and proper functioning. Prior to the upgrading, aircraft operators voluntarily purchased new equipment in accordance with their financial abilities to do so and in order to take advantage of the added capability provi- ded. To accelerate the upgrading, however, new equipment was mandated and unequipped aircraft were restricted from specified airspace. Examples of mandated • _PPROVED= DISTANCE MEASURING EQUIPMENT (FAR st. :3:_) AT OR ABOVE 7315m (24,0000t) _L i i, • IFR FLIGHT PLAN REQUIRED • IFR CAPABILITY (FAR 91.33} • 720--CHANNEL RADIO (NOT REQLffRED, BUT, Z5 KHffi FREQUENCY SPACING AT OR A£'OVF 5486,. (18, O00ft) MSL NOW IN USE IN CERTAIN SECTORS_ (FAR 91.97) (POSITIVE CONTROL AREA) i i | ,m 4420m (14, 500It') MSL • TRANSPONDER (CONTINENTAL CONTROL AREA) • ALTITUDE ENCODER D i ii .............. (FAR 91.24) ABOVE 3810m (l_500ffi MSL (FAR 91.9Oh; A--10821 Figure 49. Avionics Required for Admittance to Different Types of Airspace

NASA CR-159603

WRC Report No. 78-113-15 equipment include the two-way radio and successive requirements for more transmit- ting channels (today's transceivers have 720 channels), emergency Iocator beacons, anti-collision lights, VOR receivers, DHE, transponders, encoding altimeters, and much more. Future equipment requirements could involve discrete address beacon systems, collision avoidance equipment, microwave landing aids, and navigational receivers with double the number of receiving channels. Figure 49 depicts some of the avionics required for admittance to specific airspace.

The financial burden of airplane ownership due to the cost of insurance, hangar rental, unscheduled maintenance, avionics repair, and fuel has been staggering for many years. Add to this the cost of current enroute and approach charts, and mandated annual inspections, airworthiness directive compliance, pilot physicals, biennial flight reviews, use fees, weight taxes, state registration fees, avionics upgrading, altimeter recalibrations, airspeed system leak checks, transponder checks, and Iocator beacon servicing, and the situation becomes very discouraging.

Because of an apparently unending chain of burden increases, many airplane owners have elected to sell their airplanes and seek less costly leisure-time pursuits.

Leisure-time flying has thus been effectively throttled and aviation growth limited to manageable proportions through regulatory actions of the federal government and, to a lesser extent, state and local governments. (The major state and local government influence has been with respect to sales and property taxes.)

Regulatory actions, besides adding to the cost of airplane ownership, have caused the prices of new airplanes to skyrocket. The attendant insurance and interest costs have caused a substantial increase in airplane rental charges, with a cor- responding increase in the cost of learning to fly and decrease in the number of student pilot certificates issued (down from about 160,000 in 1967 to 129,280 in 1976). To counter this decline, the General Aviation Hanufacturers Association instituted a promotional program called TAKEOFF in September 1976 which is aimed at increasing student starts, successful completions, and, over the longer term, aircraft sales.

Hany newly implemented regulations were instituted to improve safety and the pro- ductivity of the limited volume of airspace. Growth would have been ill-advised without the regulations, and it is being restrained because of them. Only tech- nology is working to reduce the cost-inflating influence of an increasingly complex regulatory environment. In this respect, the greatest contributions over the last decade have been in the avionics/electronics realm. Here advances have improved aircraft productivity, providing an offsetting influence to rising ownership costs.

The potential for similar productivity gains due to advances in powerplant tech- nology is good.

A new type of regulatory influence is emerging that has the potential for grossly altering the delicate balance between regulation and growth• This influence creates economic burdens without providing corresponding returns with respect to safety, comfort or productivity. The EPA general aviation emissions standard promulgated in 1973 for implementation in 1979 is one such example• NASA O1-159603 WRC Report No. 78-113-15 This standard offered nothing to airport/airways productivity, nothing to airplane productivity, nothing to safety improvement (perhaps to safety degradation), and nothing to ride enhancement. Furthermore, it showed little likelihood of measur- ably improving air quality.

The recent EPA proposal to drop 1979 general aviation emissions standards "because the cost of implementation simply outweighs expected benefits," if adopted, will mean the resolution of a serious problem that has been facing the industry. Little opposition to the new proposal is expected.

An increase in the number of general aviation airports and the increased produc- tivity of existing ones will go a long way toward staving off emissions problems.

This will happen through expedited traffic movements and the reduction of hold times for takeoff clearance. The Airport/Airways Trust Fund plays an important role in this regard. Also, efforts to improve engine fuel economy will exert sufficient pressure to keep general aviation engine emissions at a low level.

Possibly an equally serious type of regulatory action as that dealing with emissions is th_ attempt to reduce the noise signature of business jets. Few in the industry will dispute the fact that jet noise reduction action is needed. This is evident from dockets of litigations over airport noise matters around the country and from community resistance to proposals for local airport expansion, a resistance that penalizes all of general aviation. The question is, how much noise reduction should be mandated.

Figure 50 suaBarizes present and proposed jet takeoff, approach, and sideline noise maximums and compares the performance of contemporary business jets with the stan- dards. Note that only the Canadair ChaUenger with its Avcc Lycoming ALF 502D turbofans meets proposed 1980 requirements. No business jet meets the 1985 re- quirements. The Cessna Citation, one of the quietest airplanes in its weight class, does not even meet the 80 FAR 36 requirements.

Figure 5] depicts proposed acoustical requirements in meaningful terms for T/F engines sized to the GATE interest. Because the 85 FAR 36 noise maximums do not vary up to the 4,536 kg (i0,000 Ibm) maximum aircraft weight limit, turbofans incorporating components designed for low noise generation producing less than 6672 N (1,500 Ibf) of thrust can conceivably be made to meet the proposed requirements when installed in a proper nacelle. There is little data base in this regard, however, and tests of existing small T/F prototypes are clearly in order to esta- blish the reasonableness of 85 FAR36.

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NASA CR-159603 WRC Report No. 78-113-15 11( 110 I " iO6 •I08 e 105 le, O_(_% TC4 IO3 t Q_04 • IO_ iIC2; r I _ I • I 100 1 @ t_o, 10( _OO 'OO 3 •99 95 ............................ , ....................... p. ................................... tP_h.....

:901 ql t92 ? t_ ---""" _ _ 8O (A') ICURRENT FAR 3b AND 7'5 FAI_ _ Jl{ SidQ|tnO lEA• ANNEX IL 75 NEW ICAO NOISE STANDARDS EFFECTIVE OCTOBER _, t977 1(,J _ 1_03 _ EPA PROPOSED FAiR 36 TOG ,oo _i •9 I 93 9a °9_9 • I 1100 I' EFFECTIVE JANUARY I. ,,9e.$_ _93 '_: 5 [1_ e _ D I • ._ .P,3 w_842 ----_" _ "" "" "" - 80 FAIR 38 Jet Takeoff -- 70 tlO 110 l t_O5 105 'tO; t10_ I II • ii II I | t 'O_ - I I00 i_oc i_ 99_ ._ _c_ I_oO /_.9,,: ._ ,...

.......... -;T.- ....... ................ .-: ....... ;.._ .................... "e" ..l).. ;;--I_- :T=._: 95 3 • 9_ • 9t3• j9_3 ........ // ._-,,.'-" _o _o i--- _j z 8O 75 ¸ FAR 36 Jet Approach ?0 7O ® _ , .., _ _ o_ a® - ,, _ ._'._ Pfe_Dusl¥ Manutac_u, ed A,rcratl Cunently Manufactured A,rcratl' Fulu_e Manutactured Acrcr8tl IN•Of RB_u_O TO MIll FAR _6_ !l_t_be¢_t TO _ _renf FAR _61 A--T0855 Figure 50. Present and Proposed Business Jet Noise Maximums REPRODUCIBILITY OF THE ORIGINAl, PAGE IS POOR I_5

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(0. 2s rim) (3. s nm_

0. 46 km G. 48 km (1.0 rim)

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110 110 ! 110_ 102 69 FAR 36 l-- 102 G9 FA R ,36

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80 79 85 R 36 77 80 36| _--73 85 FAR 36

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14 45 14 45 T4 45

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(3) (10) (3) (10) (3) (I0)

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MAXIMUM AIRCRAFT WEIGHT, t00kg A- A--10811 (I000 Ibm.)

Figure 51. Proposed Compliance Noise Levels I

Because the economic reasonableness and technical practicality of 85 FAR 36 have not been adequately assessed to date, 85 FAR 36 has not been assumed effective for !

a 1988 scenario. The proposed 80 FAR 36 regulations are expected to be adopted by the ICAO and should be in effect prior to and during 1988, however. With proper land use planning and appr( riate zoning, the noise abatement afforded by 80 FAR 36 should be adequate. Table L lists the noise maximums under this regulation that are expected to apply to turbofan engines producing less than 6672 N (1,500 lbf) of thrust.

NASA CR-159603 WRC Report No. 78-113-15 EXPECTED 1988 NOISE MAXIMUMS FOR NEWLY TABLE L.

CERTIFICATED SMALL BUSINESS JETS (THRUST PER ENGINE _ 6672 N (1,500 ibf) EPNdB Limit Measurement Point Sideline at 0.46 km (0.25 run) Takeoff at 6.48 km (3.5 nm) Approach at 1.85 km (1.0 nm) Noise regulations for propeller-driven small airplanes, and more specifically turboprop-powered small airplanes, for the 1988 timeframe are expected to corres- pond with those called out in FAR Amendment 36-4 (Appendix F, Part D, Paragraph F36.301). This amendme_ requires significant noise reductions affecting approxi- mately 20 percent of the contemporary small aircraft types. It is expected to remain in effect througl_ut the 1980's. Table LI summarizes the regulation.

Although FAR 36-4 is expected to have a negligible influence on 1988 market projections, it will influence conceptual engine designs for this market.

TABLE LI. EXPECTED 1988 NOISE REGULATIONS APPLICABLE TO NEWLY CERTIFICATED, TURBOPROP-POWERED GENERAL AVIATION AIRPLANES dB(A) Limit* Aircraft Weight Up to 599 kg (1,320 Ib) 68 plus i dB/75 kg 599 kg < weight < 1497 kg (165 Ib) (1,326 Ib) (3,300 Ib) 1497 to 5670 kg 8O (3,300 to 12,500 ib) *As obtained from horizontal test flights at rated _aximum continuous power 305 m (I,000 it) over a single noise-measuring station with the airplane in cruise configuration.

The 1988 market scenario assumes that technology will stimulate the market and regulations will moderate the stimuli. Needed regulations will become law, and potentially destructive regulations will not survive.

NASA CR-159603 WRC Report No. 78-113-15 ECONOMIC REGULATIONS There are three types of economic regulations impacting the sales of general aviation aircraft. One type stimulates sales, another discourages them, _d the third type stimulates business-use sales and discourages leisure-use sales. The type of regulations that stimulate the market include the following: • Investment tax credit (currently 10 percent) • U.S. government loan guarantee program for aircraft purchases (helpful for commuter airline purchases) • Eximbank loan program (provides 30 to 55 percent direct credit) and loan guarantee program • Business tax deductions for the use of private aircraft Sales are defeated b F punitive taxes that fatten general funds without benefiting general aviation. Several states levy this kind of tax in the form of specific ownership taxes, ad valorem sales taxes, fuel taxes, etc. Ge]Iman Research Associates of Jenkintown, Pennsylvania is studying these under the sponsorship of FAA. Canada has had a very discouraging general aviation sales tax (12 percent) and excise tax (10 percent). In mid-November 1978 the controversial 10 percent excise tax on aircraft imports was dropped and the 12 percent federal sales tax was reduced to 9 percent.

The type of economic regulation that both encourages and discourages sales involves tax collections that benefit general aviation. These include the aircraft tire/ tube excise tax, the 7-cents-per-gallon federal tax on non-commercial general aviation fuel, the $25.00 aircraft registration fee, the 2-cents-per-pound weight tax on piston-powered aircraft, and the 3.5-cents-per-pound weight tax on turbil:_- powered aircraft.

Foreign tariffs plus all of the aforementioned taxes have been assumed in the formulation of a 1988 market scenario. Additional user taxes have been excladed, because many that have been proposed would have a devastating effect on gen,ral aviation, if implemented. Examples of user taxes proposed by past and present administrations include: Administrative user charges for aircraft certification and pilot licen- sing. (On 21 September 1978 the House passed a bill which includes prohibitions against these user charges.)

• A $5.00 landing fee for landings at airports with FAA control towers.

• A $10.00 landing fee for landings at airports with FAA control towers equipped with radar.

• FAA financing of some of its operating costs with Trust Fund money.

• A 4-cents-per-gallon federal tax increase on aviation gas to 11 cents per gallon.

NASA CR-159603 WRC Report No. 78-113-15 • Graduated federal fuel taxes peaking at 35 cents per gallon in 1980.

• A 20 percent excise tax on new general aviation (non-commercial) aircraft.

A DOT study released 14 January 1977 entitled "National Transportation Trends and Choices" forecasts reductions in gene:al aviation activity by 1990 ranging from 4 to 41 percent depending on the activity parameter used. A major reduction would resuIt from government attempLs to fully recover costs for general aviation services. The industry response to the DOT study has been to demand more effi- ciency in government and the elimination of unwanted government services. The National Pilots Association, for exampIe, has suggested the folIowing steps: • A review of FAA missions at its Atlantic City, Oklahoma City, and Cambridge, Massachusetts facilities, with an eye toward a geographic merger and the elimination of overlapping functions.

• A reevaluation of FAA pay scales.

• Personnel reductions.

• A possible transfer of air traffic control functions to a profit- oriented public utility-type company.

NEW TECRNOLOGY Twenty years ago there was little general aviation instrument flying. This was I.

because the then-existent air traffic control system and available avionics were primitive by today's s_andards. Pilot workload was high, requiring a level of proficiency possessed by few general aviation pilots. Deficiencies with respect to instrument flying, therefore, caused small airplane productivity to suffer.

T Today, general aviation instrument flying is commonplace because of the avail-

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ability and increased use of sophisticated avionics equipment. One has only to listen to one of the many air traffic control frequencies to convince himself of this fact. A 1976 FAA document (ref. 3) conservatively forecasts that the general

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aviation category of instrument operations will grow at an average annual rate of

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7 percent through FY 1988.

To date, huge sums have been spent to computerize the world's air traffic control

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systems. In airborne electronics, however, there has been only a modest beginning toward capitalizing upon the potential that the digital computer holds for helping the pilot do his job. Microprocessors and microcomputers are now slowly

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easing their way into aircraft electronics. In the decade ahead, these tiny devices will make very significant changes in the avionics world.

As the "micros" proliferate, the time of the digital integrated avionics system

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and computerized flight management system will come. Signal multiplexing with the attendant combining of functions from separate black boxes is aiready common- place. With further progress, a single digital computer couid control turbine

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NASA CR-159603 WRC Report No. 78-113-15 engine fuel flow, a stability augmentation system, an autopilot, a flight director, a collision avoidance device, and navigational system displays. To conserve pre- cious panel space, some information could be displayed on the same cathode ray tube as weather information. Triple and quaa_uple redundancy could be provided to ensure system reliability. The relief in pilot workload thus prgvided would allow non-professional general aviation pilots to operate high-perforalance airplanes with greater precision and saietv than is possible today. Small airplane utility wcild also improve to the point of being competitive with commercial airline e×perie_,ce, i.e., if the small airplane performance were to rival airliner performance.

While engines that may emerge as the result of a general aviation turbine engine activity will probably become available before digital integrated avionics and computerized flight management systems mature, they will see service in an era characterized by pushbutton piloting, coupled autopilots, navigational precision, and intermittent positive control conflict detection and alerting. There will be an accompanying demand for appropriately powered small airplanes that can take full advantage of the utility provided by the new airborne electronics, e.g., by being able to fly above the weather on a straight line course from a departure gate to an arrival gate for an approach at the destination. Enroute zigzagging for weather avoidance will be shunned. The preferred airplane will have sufficient power to provide an adequate climb rate, and adequate quantities of bleed air and electrical energy for anti-ice, deice, cabin pressurization, and cabin climate control. A quiet, vibration-free ride will be stressed.

The theme in [be automotive world today is "smaller is better," and this will carry over into the aviation world. In the late 1980's, businessmen will no longer be whisked around the country in business jets which, in one coast-to-coast trip, consume as much fuel as it takes to heat a home for one year. Small engine tech- nology will therefore be stressed so that smaller high-performance airplanes can be developed which will lessen the gap between available seats and the average passenger load. Table LII provides average passenger load data as obtained from a 1975 FAR-sp0nsored survey by Civil Air Patrol cadets.

TABLE I.II. AVERAGE NL3IBER OF PERSONS TRAVFLING IN GENERAL ,\VIATI{_N ;\IRCRAFT (!975 ,qurvev) Average Ntunber Of Travelers Crew Aircraft Type (Including Crew) Requirements 2.1 Single-engine piston 2.9 Helicopters 3.8 Multi-engine piston 5.7 Turboprop 5.4 Turbojet/turbofan .@ NASA 624-159603 WRC Report No. 78-I13-15 Product Liability In 1971 the general aviation industry was rudely awakened to the financial reali- ties of product liability by a multi-million dollar jury award in connection with a 1968 light twin-engine airplane crash. Since that time, the major aircraft manufacturers have collectively paid out millions of dollars in settlements and judgments, and it is estimated that 2,000 product liability suits against light plane manufacturers are now in litigation. No or.e knows how many more claims are being settled out of court. Because of the large number of claims, defense costs, and settlement costs, product liability insurance rates have skyrocketed, with attendant product cost increases.

The industry-wide average product liability cost is about 7.5 percent of gross sales. About 5 percent goes directly to insurance premiums, with the rest attributable to time lost by employees responding to lawsuits and other costs.

By 1980, premiums could reach 10 to 12 percent.

Some argue that safety may in fact suffer because of product liability litiga- tion. The manufacturers are afraid to make improvements, goes the argument, becauoe any safety improvement could be considered a tacit admission that the appliance was originall7 not as safe as it could have been. Also, manufacturers have been reluctant to commit themselves to promising new production techniques and processes and potentially superior new designs because of the fear of product liability suits. Airplane manufaet'_rers, for example, have been slow to adopt bonded construction techniques because of apprehensions about bond strength 20 years hence. There is a similar reluctance to use composite airframes because of concerns over quality control inspection techniques, periodic airworthiness inspection techniques, and lightning protection.

A number of suits have involved powerplant and propeller structural failures.

Fearing litigation, engine manufacturers have been slow to undertake new engine development programs, and prospering airframe manufacturers seldom encourage such programs by planning protot.vping activities around experimental engines. Also, many of the smaller component manufacturers have gotten out of the general aviation business altogether because of liability vulnerability, and others have raised prices substantially. The FAts, which certilies engines and airplanes as safe for production, has tightened certification regulations to the point that added millions of dollars are required for engine and airplane type certifi- cation. The net result of all this has been khat the grudging progress ot light plane design has been brought almost to a scandsti!l, while new airplane costs have escalated to the point of being beyond the reach of most lndividual_ and many smaI1 businesses.

Prior to 1988, the legal pendulum governing strict liability in tort (wr,,ngful act for which a civil action may be filed) is expected to swing in., d_rect_on more favorable to manufacturers. Some probable changes wi 11 involve: • A requirement to prove negligence on the part of a manu[._, turer, n,,t just a defect NASA CR-159603 WRC Report No. 78-113-15 • A reinstatement of contributory negligence as a defense against product liability suits • Liability in accordance with the state of the art at the time of manufacture rather than at the ti_e of the accident • A disallowance of evidence that an airplane or engine was defective simply because later improvements were made The sooner the legal pendulum swings the opposite way, the sooner the flight envelope of general aviation airplanes can be expanded through technology which is now latent. Progress in the face of tort law has slowed, but by the mid- 1980ts it will resume after corrective action has been taken.

NASA CR-159603 WRC Report No. 78-113-15 APPENDIX B ENGLISH-TO-SI UNIT CONVERSION TABLE To Obtain English Pa rame t e r Unit SI Unit Multiple. By Inch (in) 2.54 x 10 ° Length Centimeter (cm) "I Feet (ft) 3.048 X I0 Meter (m) U.S. Statute Mile ".609344 X I0 ° Kilometer (km) (mi) 1.85324 x 10 ° Nautical Mile (rim) Kilometer (km) Area (Inch) 2 (in) 2 6.4516 x lO 0 Square Centimeter (Cm) 2 9.290304 x 10 -2 (Feet) 2 (ft) 2 Square Heter (m) 2 Vohme (Inch) 3 (in) 3 1.6387064 x 101 Cubic Centimeter (cm) 3 (Feet) z (ft) 3 2.8316846592 x 10 -2 Cubic Heter (m) 3 3.785411184 x 10 ° Liter (1) Gallon (gal) (U.S. liquid) "1 3.048 x 10 Feet/sec (ft/s) Velocity Meter/sec (m/s) "1 3.048 x lO Feet/sin Meter/min (m/m) (ft/min) Statute mile/hour 1.609344 x 10 ° Kllometer/hr (mi/h) (k m/h) 1.85200 x 10 ° Knot Kllometer/hr (k_lh) Force 4.44822161 x I0 ° Pound force (lbf) Newton (N) 4.5359237 x 10 -t Mass Pound mass (ibm) Kilogram (kg) 6.8947572 x 10 ° Pressure Pound per square inch Kilopastal (kPa) (lbf/in 2 ) NOTE: One N/m 2 = 1 pascal One atm = 101.325 kPa KSI 6.8947572 x I0 ° Stress Hegapas(al (HPa) Ibm/in 3 2.76799 x 101 Gram/Cent imeter 3 Density (g/cm 3 ) Flowrate Pounds/hour (Ibm/h) 4.5359237 X I0 "I Kilogram/hr (kg/h} l_____ _ 4.90355 x lo -2 Flow Parameter s kPa s psia

NASA CR-159603

WRC Report No. 78-113-15 4.892467 x 10 °t Kilometer/liter Nautical mile/gallon Fuel Efficiency (kin/l) (n=/gal) in-lbf 1.129848 x lO -I Watt (W) Power s ft-tbf 1.3558179 x I0 ° Watt (W) s 7.4569987 x I0 -I KilowaLt (kW) Horsepower (hp) hp. s 1.6439869 x I0 ° kW . s Specific Output Power Ibm kg 9.80665 x I0 ° N-s/kg (lbf-s/lbm) Specific Thrust 1.01972 x I0 "l kg/N-h TSFC (lbm/lbf-h) Specific Fuel Consumption (Thrust) 6.082774 x I0 -l kg/kW-h BSFC (Ibm/hp-h) Specific Fuel Consmaption (Power) 3eg. Fahrenheit (OF) Temperature T k = _ (T F + 459.67) Kelvin (K) Beg. Rankine (OR) T k = _ TR Kelvin (K) + 273.15 NOTE: Tk = Tcelsiu s 1.055056 x I03 Joule (J) Btu Energy 2.3260 x I0 ° (8tu/Ibm) Kilojoule/Kilogram Enthalpy (kJ/kg) Radians (rad) Degrees (o) 1.745329 x 10 -2 Unit Angle Acceleration of Gravity = 32.1725 ft/s 2 = 9.80621 m/s 2

NASA CR-159603

WRC Report No. 78-113-15 REFERENCES I. Anon. : The General Aviation Aircraft and Associated Avionics and Instrumentation Market t 1974-1984). Frost and Sullivan, Inc. , New York, N.Y., 1975.

Built Air- 2. Anon. : TLe Genera I Aviation Market through 1982 (U.S.

craft). Forecast Associates, Inc., Ridgefield, CT, 1977.

3. AVP-120: Aviation Forecasts - Fiscal Years 1977-1988. FAA Aviation Forecast Branch Technical Report FAA-AVP-76-17, September 1976.

3. Ziff-Davis 4. Anon.: The Weekly ol Business Aviation, Volume 24, No.

Publishing Co., 17 January 1977.

2, Ziff-Davis 5. Anon.: The Weekly of Business Aviation, Volume 22, No.

Publishing Co., New York, N.Y., 12 January 1976.

Ziff-Davis 6. Anon. : Flying Annual and Buyer's Guide, 1977 Edition.

Publishing Co., New York, N.Y., 1977.

7. Anon. : The AOPA Pi lot. Aircraft Owners and Pi lots Association, Washington, D.C., March 1917.

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Aerospace Facts 10. Aerospace Industries Assoctation of America, Inc.: 1976/77 and Figures, HcGraw Hill, New York, N.Y.

11. Peter Garrison" "Turboprop Singles: Around the Corner--or out of Sight?," Flying Magazine, Ziff-Davis Publishing Co., Net, York, N.Y., December 1977.

Av i a t i on 12. Betty V. Cayce : Census of U.S. Civil Aircraft. Vedera 1 Administration, 31 De,ember 1915.

NASA CR-159603 NRC Report No. 78-113-15 BI BLIO(]PO_>h_ Policy Development Division of the FAA Office of Aviation Policy: FAA Aviation Forecasts, Fiscal Years 1979-1990, September 1978.

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