General Aviation Propulsion

NASA Conference Publication 2 126

General Aviat~on Propulsion

{UASA-CP-2126) GEUEEAL A V I A T I O U PSOPOLSICI ( P A S A ) 431 E EC A 1 9 / E E A01 CSCL 2 1E

Proceedings of a conference held at

NASA L e w i s Research Center

Cleveland, Ohio

November 28-29, 1979

N A S A

NASA Conference Publication 2126

General Aviation Propulsion

Proceedings of a conference held at

NASA Lewis Research Center

Cleveland, Ohio

November 28-29, 1979

National Aeronautics and SWce Admin~stratton SckntHk a d Tschnkal lnformrtkn Branch FOREWORD The National Aeronautics and Space Adminrstration i s a c t i v e l y involved i n t h e quest f o r improved g e n e r a l a v i a t i o n a i r c r a f t .

Programs exploring and demon-trating new technologies i n general a v i a t i o n propulsion a r e being conducted a t t h e Lewis Research Canter and by i n d u s t r i a l c o n t r a c t o r s and u n i v e r s i t y grantees.

These programs a r e t h e Q u i e t , Clean, General Aviation Turbofan (QCCAT) program; t h e General Aviation Turbine Engine (GATE) study program; t h e g e n e r a l a v i a t i o n p r o p e l l e r technology program; and t h e advanced r o t a r y , d i e s e l , and reclprocating engine programs. A two-day conference was h e l d i n November of 1979 t o provide r e p r e s e n t a t i v e s from government, industry, and u n i v e r s i t i e s with t h e l a t e s t f i n d i n g s of t h e s e programs. This p u b l i c a t l o n c o n t a i n s a l l t h e papers presented a t t h a t conference.

G i l b e r t K . Sievers NASA Lewis Research Center Conference chairman CONTENTS Page F O R E W O R D . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . iii OVERVIEU OF NASA QCGAT PKOGRAH

G i l b e r t K. S i e v e r s . . . . . . . . . . . . . . . . . . . . . . . . . . 1

AIRESEARCH QCGAT ENGINE, AIRPLANE, AND NACELLE DESI;N FEATURES

R o g e r W. H e l d e n b r a n d . . . . . . . . . . . . . . . . . . . . . . . . . 11

AXRESEARCH QCGAT ENGINE PERFORMANCE Ai:D EMISSIONS TESTS

W i l l i a m M. N o r g r e n . . . . . . . . . . . . . . . . . . . . . . . . . . 45

AXRESEARCH QCGAT ENGTNE - ACOUSTIC TEST RESULTS

L a r r y S. K i s n e r . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

QCGAT XIRCRAFT/ENGINE 3 E S I S N FOR KEDUCED NOISE AND EMISSIONS L e o n a r d I ' A n s o n and K e n n e t h H . T e r r i l l . . . . . . . . . . . . . . . . 101 AVCO LYCOHI NG QCGAT PROGRAM DESIGN CYCLE, DEMONS RATED PkEFORNANCE AND EM1 SSIONS

P h i 1 Fogel and Ange lo K o s c h i e r . . . . . . . . . . . . . . . . . . . . 135

AVCO LYCOMING QUIET CLEAN GENERAL AVIATION TURBOFAN ENGINE

C r a i g A . W i l s o n . . . . . . . . . . . . . . . . . . . . . . . . . . . 155

SUMYARY OF NASA QCGAT PR0i;RAM

G i l b e r t K. S i e v e r s . . . . . . . . . . . . . . . . . . . . . . . . . . 189

NEW OPPORTUNITIES FOR FIITL'KF , S I W L , GEKERAL-AVIAT ION TURBINE ENGINES (GATE)

W i l l i a m C. S t r a c k . . . . . . . . . . . . . . . . . . . . . . . . . . 195

AN OVERVIEU O F NASA RESEARCH ON POSITIVE DISPLACEMENT GENERAL-AVIATION ENGINES

E r w i n E. K e m p k e , Jr. . . . . . . . . . . . . . . . . . . . . . . . . 2 2 1

THE SPARK-IGNITION AIRCRAFT PISTON ENGINE OF THE FUTURE

K e n n e t h J. S t u c k a s . . . . . . . . . . . . . . . . . . . . . . . . . . 2 3 1

LIGHTWEIGHT DIESEL AIRCRAFT ENGINES FOR GENERAL AVIATIOPi . . . . . . . . . . . . . . . S t e v e n G. B e r e n y i a n d A l e x P. B r o u w e r s . 2 4 7 ADVANCED ROTARY ENGINE STUDIES

C h a r l e s J o n e s . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 8 7

POSITIVE DISPLACE,YENT TYPE GENERAL-AVIAiION ENGINES : SUMMARY AND CONCLIJDING REMARKS

E r w i n E. K e m p k e , Jr. . . . . . . . . . . . . . . . . . . . . . . . . 313

= ' . s t NASA PROPELLER TECHNOLOGY PROGRAM

D a n i e l C. H i k k e l s o n . . . . . . . . . . . . . . . . . . . . . . . . . 315

W W SPEED PROPELLECS - IMPACT OF ADVANCED TECHNOLOGIES

ADVANCED TURBOPROP POTENTIAL FOR HIGH SPEED

B e r n a r d S . G a t z e n . . . . . . . . . . . . , . . . . . . . . . . . . . 345

HIGH-SPEED-PROPELLER WIND-TURBIhE AEHOACOUSTIC RESULTS R o b e r t J . J e r a c k i and J a m e s H . D i t t m a r . . . . . . . . . . . . . . . . 361 ADVANCED PROPELLER AERODYNAMIC ANALYSES L a w r e n c e J . B o b e r . . . . . . . . . . . . . . . . . . . . . . . . , . 375 PROPELLER AEROACW ST I C METHODOLOGIES K e n n e t h D . Korkan and G e r a l d N . C r e g o r e k . . . . . . . . . . . . . . . 387 NASA PROPELLER NOISE RESEARCH

George C . G r e e n e . . . . . . . . . . . . . . . . . . . . . . . . . . . 405

PROPELLER DYNAMIC AND AEKOELASTIC EFFECTS OVERVlEW OF NASA QCCAT PROGRAM Gilbert K. Sievers National Aeronautics and Space Administration Lewis Research Center Today, t h e curbofan-powered g e n e r a l a v i a t i o n f l e e t i s growing a t a g r e a t e r p e r c e n t a g e r a t e t h a n t h e rest of t h e g e n e r a l a v i a t i o n f l e e t . Jet powered g e n e r a l - a v i a t i o n a i r c r a f t n a b e r e d o v e r 2100 i n 1 9 7 8 w i t h a n n u a l sales of a b o u t 250 i n 1978. Annual s a l e s a r e e x p e c t e d t o be o v e r 400 by 1Y85.

Jet powered g e n e r a l - a v i a t i o n a i r c r a f t u t i l i z e a l l o f t h e approximately 400 commercial a i r p o r t s i n t h e United S t a t e s , p l u s a s i g n i f i c a n t number of g e n e r a l - a v i a t i o n a i r p o r t s l o c a t e d i n suburban a r e a s . There are approximately 16 100 of t h e s e suburban a i r p o r t s , most of which a r e l o c a t e d i n small coePun- ities w i t h no i n d u s t r i a l b u f f e r z o n e s and w i t h p e o p l e l i v i n g nearby. T h e r e f o r e , g e n e r a l a v i a t i o n h a s t h e p o t e n t i a l f o r g r e a t e r community reaction t o n o i s e and p o l l u t i o n t h a n commerc la1 and l a r g e t r a n s p o r t a i r c r a f t .

The QCGAT program s e e k s t o improve t h e environment a1 c h a r a c t e r i s t i c s o f c i v i l a i r c r a f t i n t h e v i c i n i t y o f a i r p o r t s . I n t h e p a s t , NASA h a s c o n c e n t r a t e d i t s e f f o r t s i n e n g i n e r e w a r c h toward t h e commercial o r l a r g e a i r c r a f t f i e l d .

Now w i t h t h e QCGAT prcgram, NASA h a s a p p l i e d t h i s l a r g e e n g i n e technology t o s m a l l e n g i n e s i n t h e g e n e r a l - a v i a t i o n o r small-engine f i e l d .

PROGRAM OBJECTIVES The program was conducted i n t w o phases, a s t u d y phase a n d a n e x p e r i u e n t a l phase. The o b j e c t i v e s f o r t h e s t u d y phase were t o examine t h e a p p l i c a b i l i t : ~ o f c u r r e n t l a r g e t u r b o f a n technology t o s m a l l e n g i n e s , t o do a p r e l i m i n a r y d e s i g n of t h e QCGAT e n g i n e , and to d e v e l o p t h e r e q u i r e m e n t s and a program p l a n f o r t h e e x p e r i m e n t a l phase.

The o b j e c t i v e of t h e e x p e r i m e c t a l phase w a s t o d e m o n s t r a t e t h a t t h e a p p l i - c a t i o n of large-turbofan-engine technology t o s m a l l , g e n e r a l - a v i a t i o n t u r b o f a n e n g i n e s c a n r e s u l t i n l e s s r a i s e , lower e m i s s i o n s , and a c c e p t a b l e f u e i consump- t ion. While low e m i s s i o n s and a c c e p t a b l e f u e l consumption a r e i m p o r t a n t , t h e program b a s p r i m a r i l y d i r e c t e d toward low n o i s e .

F i g u r e 1 shows some of t h e NASA programs t h a t hav'e c o - ~ t r i b u t e d t o t h e c u r - r e n t s t a t u s o f l a r g e t u r b o f a n e n g i n e technology. These i n c l u d e t h e WSEE pro- gram, t h e Q u i e t kngine program, t h e Quiet Nacelle prrograms, t h e Refan program, and t h e Clean Combustor program.

FROGQX APPROACH A s was mentioned p r e v i o u s l y , t n e program was conducted i n two phases. The s t u d y phase helped a e s i n e t h e e x p e r i m e n t a l phase. I t was s t a r t e d i n Ap . of 1975 and l a s t e d about b months. Three c o n t r a c t o r s were invoived i n t h e s t u d y phase: G a r r e t t AiResearch, Avco Lysoming, ana t h e General E l e c t r i c 20.

The experiuiental pbase Gas a c o v q e t l t i v e procurement. Two b i a s were re- c e i v e d , and c o n t r a c t s v e r e awarded t o both b i a a e r s , AiMesearch ana Avco Lycoming. The e x p e r i m e n t a l phase c o n s i s t e d o r a demonstration program i n which each c o n t r a c t o r was t o d e s i g n , f a b r i c a t e , and t e s t a QCGAT engine. Each e n g i n e was t h e n d e l i v e r e d t o SASA Lewis i o r f u r t h e r t e s t i n g .

The t e c h n i c a l approach f o r each c o n t r a c t o r v a s t o u s e a n e x i s t i n g modem g a s g e n e r a t o r o r e n g i n e c o r e t o save devi-i~puient time anli money. The e n g i n e was t o develop l e s s t h a n 3000 pounds of s t a t i c t h r u s t , and d i l r o t a t l n g p a r t s were t o be f l i g h t w o r t h y . A b o i l e r p l a t e r a t h e r t h a n a i l i g n t u o r t h y n a c e l l e was ac- c e p t a b l e . However, t h e i n t e r n a l aerodynamic c o n t o u r s and t h e a c o u s t i c t r e a t m e n t f o r t h e n a c e l l e had t o be o t f l i g h t design.

Goals were set f o r n o i s e , e m i s s i o n s , and i u e l consumption. Tne e m i s s i o n s g o a l s s e l e c t e d were t h e now abandonea 1979 EPA e m i s s i o n g o a l s f o r c l a s s T I en- g i n e s - NASA g e n e r a t e d i t s own n o i s e g o a l s . S i n c e e x i s t i n g g a s g e n e r a t o r s were b e i n g used, d r a s t i c reductions i n f u e l consumptiori could n o t be expected. How- e v e r , f u e l consumption s h o u l d n o t s u f f e r a t t h e expense of r e a u c l n g n o i s e and p o l l u t a n t emissions. T h e r e f o r e , a f u e l - c o n s u q t i o n g o a l e q u a l t o o r b e t t e r t h a n e x i s t i n g e n g i n e s was s e t .

S i n c e t h e f l i g h t - n o i s e c a l c u l a t i o n s r e q u i r e a f l i g h t p r o f i l e , each con- t r a c t o r was asked t o s y n t h e s i z e a n a i r c r a f t f o r t n e i r engine. A twin-engrne a i r c r a f t was s e l e c t e d f o r c s n s i s t a n c v in u ~ t i s e c a i c u i a t i o n s .

F i n a l l y , t h e e n g i n e s v e r e t o be d e l i v e r e d t o SASA Leuis f o r f u r t h e r e x p e r l - a e n t a l t e s t i n g .

PROGRAM GOALS The NASA g e n e r a t e d n o i s e g o a l s a t t h e FhR-36 measuring s t a t i o n s a r e shown i n f i g u r e s 2 t o 4. F i g u r e 2 i s f o r t a k e o f f . The g o a l and t h e F A R 36 r e q u i r e - ments a r e e x p r e s s e d i n EPNdB a s a f u n c t i o n o t a i r c r a f t t a k e o f f g r o s s weight.

The n o i s e c e r t i f i c a t i o n l e v e l s f o r f o u r twin-engine a i r c r a f t a r e a l s o shown.

These a i r c r a f t a r e c o n s l d e red t o be among t h e q u i e t e s t t~irbotan-powered a i r c r a f t i n t h e f l e e t today. A s c a n be s e e n , t h e N A S A g o a l i s 8 t o 12 E P N d B below anv g e n e r a l - a v i a t i o n a i r c r a f t f l y i n g today. I n t h e range of a i r c r a f c g r o s s w i g h t used i n t h e QCGAT program, t h e KASA g o a l i s 16 t o 19 EPhdB below t h e c u r r e n t 1977 FAA r u l e .

F i g u r e s 3 and 4 show t h e s i d e l i n s and approach goals. T h e p l o t s a r e s i m i - - l a r t o t h a t shown f o r t a k e o f f .

Again, t h e KASA g o a l s a r e w e l l below e x i s t i n g q u i e t g e n e r a l - a v i a t i o n a i r c r a f t and t h e 1977 FAA r u l e .

These g o a l s were s e t t o i n s u r e t n e i n c l u s i o n of e x i s t i n g low-noise t e c n n o l - ogy i n t h e QCGAT e n g i n e d e s i g n s . Achievetnent o f t h e s e g o a l s w i l l r e s u l t i n a i r - c r a f t n o i s e levels t h a t a r e p e r c e i v e d t o be 4 5 t o 5 5 p e r c e n t less n o i s y t h a n t h e l e v e l s o f t h e q u i e t e s t c u r r e n t b u s i n e s s jets.

Another way of i l l u s t r a t i n g t h e e f f e c t of a c h i e v i n g t h e s e g o a l s is by u s i n g n o i s e f o o t p r i n t areas. A n o i s e f o o t p r i n t i s t h e ground a r e a below t h e a i r c r a f t w h i c h is s u b j e c t t o a n o i s e l e v e l g r e a t e r t h a n a g i v e n l e v e l d u r i n g t a k e o f f and l a n d i n g . The f o o t p r i n t a r e a f o r a n a i r c r a f t u s i n g t h e WGAT e n g i n e s is p r e - A comparison d i c t e d t o be o n e - t e n t h t h a t of t h e q u i e t e s t c u r r e n t b u s i n e s s jets.

o f t h e f o o t p r i n t s i s shown i n f i g u r e 5 . S i m i l a r r e d u c t i o n s were a c n i s v e d be- tween t h e Lear 35 a n d t h e ~ i R e s e a r c h QCGAT powered a i r p l a n e and b e t u e e n t h e C i t a t i o n and t h e Avco-Beech QCGAT powered a i r p l a n e . A l s o , l i t t l e v a r i a t i o n i n p e r c e n t r e d u c t i o n of f o o t p r i n t area e x i s t s f o r l e v e l s between 70 and 90 EPNdB.

Achievement of t h e s t r i n g e n t QCGAT n o i s e g o a l s s h o u l d e l i m i n a t e n o i s e as a m a j o r c o n s t r a i n t o n t h e f u t u r e growth o f t h e turbofan-powered, g e n e r a l a v i a t i o n f l e e t .

The QCGAT e m i s s i o n s g o a l s a r e s h o r n i n t a b l e I. T h e s e g o a l s were t h e 1979 €PA e m i s s i o n g o a l s f o r c l a s s T1 e n g i n e s . &PA h a s s i n c e abandoned t h e s e g o a l s as b e i n g t o o s t r i n g e n t f o r t h i s t i m e frame. iiowever, t h e s e g o a l s were k e p t t o r t h e QCGAT program.

The QCGAT performance g o a l s are g i v e n i n t a b i e IL. The g o a l s a r e b a s e d on t h e r e s u l t s from t h e s r u d y p h a s e and a r e c o n s i d e r e d t o be a c h i e v a b l e g o a l s w i t h f u e l consumption e q u a l t o o r b e t t e r t h a n e x i s t i n g e n g i n e s .

QCGAT ENGINES ~ r t i s t s ' v e r s i o n s o f t h e QCCAT e n g i n e s a r e siioun i n f i g u r e s b and 7. D e - s i g n d e t a i l s of b o t h e n g r n e s , e x c e p t those of t h e m i x e r n o z z l e s , w i l l b e d i s - The d e t a i l s o f t h e m i x e r n o z z i e s are u n d e r t h e NASA c u s s e d i n f o l l o v i n g p a p e r s .

C o n t r a c t o r r e p o r t s on t h e m i x e r s E a r l y Domest i c iss semi cat i o n o r FEDD c l a u s e .

have b e e n d i s t r i b u t e d t o U. S. companies.

The r e a s o n s t h a t two e n g i n e s r a t h e r t h a n o n e were s e l e c t e d f o r t h e ~ C G A T program, a r e e v i d e n t i n t a b l e 11. The AiResearch e n g i n e i s a h i g h e r t h r u s t ma- c h i n e and is d e s i g n e d f o r a n a i r c r a f t t h a t c r u i s e s a t h i g h speed and a l t i t u d e and h a s a long range. The Avco e n g i n e i s a i o w - t h r u s t machine d e s i g n e d f o r a n a i r c r a f t t h a t c r u i s e s lower, s l o w e r , and h a s a n i n t e r m e d i a t e range.

QCCAT AIRCRAFT An a r t i s t ' s v e r s i o n of t h e a i r c r a f t s y n t h e s i z e d by Beech A i r c r a f t f o r t h e Avco QCGAT e n g i n e i s s h o r n i n f i g u r e 8. The AiResearch s y n t h e s i z e d a i r c r a f t is a s t r e t c h e d v e r s i o n of t h e L e a r j e t 35. A photograph of 1 - e a r j e t 35 i s shown i n f i g u r e 9. The AiResearch QCGAT e n g i n e powered v e r s l o n a p p e a r s t o be very s i m i - l a r . The major cnanges t o t n e outward appearance a r e l a r g e r n a c e l l e s ana a longer fuselage.

A comparison of t h e QCGAT a i r c r a f t v i t n s i m i l a r e x i s t i n g a i r c r a f t i s given i n t a b l e 111. The Avco QCGAT powered a L r c r a f t performs a s i m i l a r mission t o t h a t of the C i t a t i o n I, and, even through it i s a mucn l i g h t e r a i r c r a f - t , i t nas both a higher maximum payload c a p a b i l i t y and a lower f u e l consumption a t compar- The AiResearcn QCGAT a i r c r a f t nas a l a r g e r passenger or a b l e c r u i s e conditions.

payload capability tnan t h e Learjet 35 and a l s o has lower f u e l consumption a t comparable c r u i s e conditions.

Keeping in mind t h e lower n o i s e ana l o v e r p o l l u t a n t emissions o i tne QCCAT powered a i r c r a i t wnile making t hese comparisons, the QCGAT advantages a re ap- parent.

WGAT SCHEDULE A bar c h a r t snouing t h e schedule f o r t h e major items i n the QCGAT experi- mental phase a r e shown i n f i g u r e The experimental phase s t a r t e a about tne 10.

end of 1976. The e r ~ i n e design, i a b r i c a t i o n , ana t e s t i n g were accompiisnea The AiResearcn QCGAT engine was within tire time frames shown i n t h e figure.

delivered t o Lewis i n February 1979, and tire Avco engine i n October 1979. Final c o n t r a c t o r r e p o r t s covering tne deve lopmenc o i tnese engines w i 11 be avai i able and distri'outea i n t h e near future.

TABLE I. - -1 EHLSSIOYS GOALS - INSTALLED

- - h r i a r ion Coat ract goal g/M rec lbllOOO l b thrust-hr/cycle Carbon mnolride 0.266 9 . 4 Unburned hydrocarbaas .045 1.6 Oxides of nitrogen .lo5 3.7

TABLE 11. - QCCAT PULFOliMNCE COALS

[Standard day; i n s t a l l e d ] r AVCO AiResearch Sea level 17312 (3892) 1166 ( l b l l ) Tbrusc, P(lb) 0.0370 (0.363) takeoff 0.0631 (0.423) SPC, kg1ha-N (lblht-lb)

--------------

2157 (485) Design cruise n r u r t , N (Lb)

--------------

fl - 0.6

0.0610 (0.628) SFC, kglhrN ( 1 b l h r l b ) 7600 o (%5 000 f t ) I

---------- ----

Design cruise Thrust, N ( lb) 4017 (903)

-'-'-----'---' / 0.0759 (P.7U)

flr 0.8 SFC. kg/hr-N (lb/hr-lb) 12 200 m (40 000 f t )

1 i

TABLE 111. - AIitC1UFT COMPARISON

AiRestareh Citation I ~ e a r ~ e t 35 AVCO

-

8674 (19 122) 5375 ( 1 1 850) 7711 (17 000) 3538 (7800) Takeoff gross ueight, 4 ( l b ) 14 7 10 6 N u m b e r of seat8 1231 (2714) (I16 (1800) 862 (1900) 1134 (2500) Maximu pay load. kg ( lb) 3456 (1866) 1732 (935) 3926 (2120) 1408 (769) m i m u m range a t m a x i m &

pay ... ad, *I ( m i )

16.4 (4.b2) 7.56 (2.13) 14.0 ( 3.90) Paraewer lu/kg-fuel 13.8 (3.90) (passenger r i l e r l l b - f u e l ) 236 (459) 183 (355) 239 (464) 185 (359) W x b u cruise speed, m/rec (knot 8) 0.801 0.607 0.810 0.626 W x i u c r u i r e oach number 13 120 (45 000) 13 720 ( 4 5 000) 1 2 500 (41 000) Ceiling, m ( f t ) 12 340 (40 500)

SIDELINE NOISE GOAL

1969 FAR-36 REQil I REMENT

-

1977 RULE EPNdB

60 a

1 10 100 l o o o x l d

TAKEOFF GROSS WE1 GHT, kg

1 I I I

3 10 100 loooxld TAKEOFF GROSS WE1 GHT, Ib Figure 3

APPROACH NOISE GOAL

1969 FAR-36 REQUIREMENT EPNdB

60 a-

1 10 100 1000~ 1 ~ 3 TAKEOFF GROSS WE1 GHT, kg 3 10 100 1000~103 TAKEOFF GROSS WEIGHT, Ib Figure 4 AIRESEARCH QCGAT ENGINE, AIRPLANE, AND NACELLE DESIGN FEATURES Roger W. Heldenbrand AiResearch Manufacturing Company of Arizona A Dikision of The Garrett Corporation The Q u i e t , C l e a n , G e n e r a l A v i a t i o n T u r b o f a n (QCGAT) e n g i n e a n d nacelle s y s t e m w a s d e s i g n e d and tested by t h e AiResearch Uanu- f a c t u r i n g Company o f Arizona under C o n t r a c t to NASA Lewis R e s e a r c h C e n t e r . The e n g i n e u t i l i z e d t h e core of t h e AiResearch W e 1 TFE731-S e n g i n e and i n c o r p o r a t e d s e v e r a l unique n o i s e - a n d e m i s s i o n s - r e d u c t i o n f e a t u r e s , Major performance, e m i s s i o n s , and n o i s e g o a l s were d e m o n s t r a t e d , and t h e e n g i n e and n a c e l l e were d e l i v e r e d to NASA Lewis R e s e a r c h C e n t e r f o r a d d i t i o n a l t e s t i n g .

INTRODUCTION The d e s i g n f e a t u r e s o f t h e QCGAT e n g i n e , a i r p l a n e and n a c e l l e are d e s c r i b e d i n t h i s p a p e r - T e s t programs and r e s ~ l t s o f t h e e n g i n e performance, e m i s s i o n s , and n o i s e tests a r e d i s c u s s e d i n s u b s e q u e n t p a p e r s .

A n isometric cutaway o f t h e QCGAT e n g i n e i n a f l i g h t - t y p e n a c e l l e is shown i n f i g u r e 1. The e n g i n e was d e s i g n e d around t h e core o f t h e AiResearch Model TFE732-3 t u r b o f a n e n g i n e , T h i s e n g i n e is a p r o d u c t i o n u n i t used i n s e v e r a l d o m e s t i c and f o r e i g n b u s i n e s s jets. The e n g i n e c o n s i s t s o f t h e TFE731-3 h i g h - p r e s s u r e ( H P ) s p o o l and l o w - p r e s s u r e (LP) compressor, plus s e v e r a l unique and new com- por.ents i n c l u d i n g a low-speed f a n , a f a n g e a r b o x , associated d u c t s and s t r u c t u r e , a reduced-emissions combustion system, and a n LP t u r - b i n e .

An airplane design, synthesized by Garrett in order to evaluate the QCGAT Engine, was selected to be similar to business jets using Model TFE731 Engines, but somewhat larger, thus taking advantage of the higher thrust level.

Two naceiles were designed for the program: o A production flighk-weight nacelle featuring integral acoustic treatment o A 'workhorse' nacelle, fabricated especially for this test program and featuring replaceable inlets, acoustic panels, and a special mixer compound nozzle.

An overall task schedule is shown in figure 2 . The QCGAT Phase I1 experimental program was divided into ten major tasks.

These culminated with delivery of an engine, associated test sup- port equipment, and spares at the end of 25 months. As experienced with most hardware-oriented programs, difficulties and delays were experienced with design iterations and fabrication schedules. How- ever, the test program was accelerated, and the engine was shipped on schedule.

The technical goals for the program are listed in table 1.

Performance goals represented a TSFC improvement of approximately 9 percent over other turbofan engines. The noise goals were 10- to 15-EPNdB below the Federal Aviation Administration's FAR Part 36 requirements. The emissions goals were identical to the EPA 1979 standards for T-1 class engines. (The EPA subsequently determined that general aviation was not a significant source ot air polution and therefore did not impose these standards) .

aYGINE DESIGN T5e principal program objective was to demonstrate the appli- cation of large turbofan noise- and emissions-teduction techn~logy to small general (ririation turbofans. To do this, a number of unique features were incorporated in the basic design of the QCGAT engine in order to reduce the emissions and noise levels below those of the already quiet TFE731 engine. This work was initiated in 1975 during the -AT Phase I study. Twelve candidate engine configurations were screened. Uany parameters were considered, i ~ c l u d ing : Pan pressure ratios at takeoff and cruise Thrust TSFC Lapse rate Fan diameter Installed weight Noise Nacelle drag Acoustic shielding Cost .

The engine cycle selected f o r the program represented a practical engine from the standpoints of cost, weight, airplane/crcelle interference drag, and cruise propulsion efficiency. The engine also exhibited high potential for reduction of turbomachinery and jet noise, and reduction of chemical and visible exhaust emissions.

The design point for the engine (typical for most modern business jets) and principal engine cycle parameters are iisted in table 2 .

Figure 3 is a cross-section of the overall QCGAT engine design. The QCGAT engine is based on the core of TFE731-3, but i n c o r p o r a t e s a f a n used i n t h e AiResearch Model ATF3 e n g i n e . The f a n is d r i v e n by a new l o w - p r e s s u r e t u r b i n e v i a a newly d e s i g n e d f i v e s t a r - g e a r gearbox. The l o w - p r e s s u r e compressor is d r i v e n d i r e c t l y by t h e l o w - p r e s s u r e t u r b i n e . The HP spool c o n s i s t s o f a c e n t r i f u g a l compressor d r i v e n by a c o o l e d a x i a l t u r b i n e . The com- b u s t o r is an a d a p t a t i o n o f a p r o d u c t i o n TFE731 combustor t h a t w a s d e s i g n e d f o r l o w smoke. Accessories and t h e f u e l c o n t r o l are driven by t h e HP s p o o l t h r o u g h a tower s h a f t . A f i n n e d h e a t exchanqer i n t h e f a n b y p a s s d u c t cools t h e o i l f o r t h e f a n g?arbox and e n g i n e l u b r i c a t i o n system. The f lange-to-f l a n g e l e n g t h o f t h e engirie is 143.15 c m (56.36 in.) and t h e f a n d i a m e t e r is 77.47 c m (30.5 i n . ) . When f u l l y i n s t r u m e n t e d and w e t , t h e test e n g i n e

weighs a p p r o x i m a t e l y 426.38 kg, (940 l b ) . F i g u r e 4 shows t h e

e n g i n e i n t h e t e s t cell p r i o r to i n i t i a l c a l i b r a t i o n .

The major a c o u s t i c d e s i g n f e a t u r e s o f t h e QCGAT e n g i n e and n a c e l l e s y s t e m a r e shown i n f i g u r e 5 and o u t l i n e d below: o N o i n l e t g u i d e v a n e s o High i n l e t t h r o a t Mach nu,aber o L o w t i p s p e e d , s i n g l e - s t a g e f a n (36 b l a d e s ) o Phased i n l e t a c o u s t i c t r e a t i n e n t o Optimized f a n b l a d e - t o - s t a t o r vane c o u n t o 2.12 r o t o r - c h o r d , f a n - t o - s t a t o r s p a c i n g o Phased f a n b y p a s s d u c t a c o u s t i c t r e a t m e n t o Low f a n jet v e l o c i t y o Reverse -f l o w a n n u l a r c o n b u s tor o High-work, low-pressure t u r b i n e w i t h l o w c o r e - e x h a u s t v e l o c i t y o 12-lobe mixer compound n o z z l e .

With t h e p o s s i b l e e x c e p t i o n o f t h e r e v e r s e - f l o w combustor and t h e mixer compound n o z z l e , e a c h o f t h e s e f e a t u r e s above is b a s e d on work done w i t h l a r g e e n g i n e s and is a d i r e c t a p p l i c a t i o n o f t h a t technology.

COM WNENT DESIGNS The QCGAT f a ? ( f i g , 6 ) is a 36-blade d e s i g n d e r i v e d from t h e f a n used on t h e AiResearch Model ATF3 Turbofan e n g i n e . The p r i n c i - p a l d e s i g n f e a t u r e s a r e g i v e n on table 3 w i t h t h e d e s i g n p o i n t a t 12,192 m ( 4 0 , 0 0 0 f t ) , s t a n d a r d day a t a f l i g h t Mach number of 0.8.

T h i s f a n is a p p r o x i m a t e l y 1 0 - p e r c e n t l a r g e r i n d i a m e t e r t h a n t h e TFE731 f a n , and rotates a t 1 7 - p e r c e n t s l o w e r speed. Thus, f a n t u r - bomachinery component n o i s e l e v e l s a r e lower. The f a n - s t a g e flow p a t h ( f i g . 7 ) was d e s i g n e d to minimize t h e c o r e - f l o w Mach num- ber and t o p r e v e n t l a r g e a c c e l e r a t i o n s i n t h e s t r u t r e g i o n s . Abso- l u t e l o c a l Mach numbers, and b l a d e and vane c o u n t s a r e also shown.

The bypass s t a t o r l o c a t i o n is s l i g h t l y m r e t h a n t w o r o t o r - c h o r d l e n g t h s downstream. The vane c o u n t s of b o t h s t a t o r s were s e l e c t e d to minimize r o t o r and s t a t o r n o i s e i n t e r a c t i o n , The bypass p e r f o r - mance map ( f i g . 8 ) shows t h e e n g i n e o p e r a t i n g l i n e s f o r co- a n n u l a r n o z z l e and mixer compound e x h a u s t n o z z l e from i d l e t h r o u g h t a k e o f f . S l i g h t l y g r e a t e r s u r g e margin was a c h i e v e d w i t h t h e mixer compound n o z z l e . A f a n component r i g test was n o t c o n d u c t e d . How- e v e r , a d e q u a t e d a t a was a v a i l a b l e from t h e Model ATF3 f a n r i g tests, and a c t u a l QCGAT e n g i n e o p e r a t i o n to d e f i n e t h e QCGAT f a n f o r t h e e n g i n e performance model.

The f a n g e a r b o x ( f i g . 9 ) is s i m i l a r to t h a t of t h e TFE731.

However, t h e o v e r a l l g e a r r a t i o was changed from 0.5559 t o 0.4634 t o match lower f a n speed. R e s i l i e n t mounts were i n c o r p o r a t e d on t h e s t a r g e a r s to m a i n t a i n g e a r a l i g n m e n t d u r i n g h i g h - t o r q u e l o a d s .

The s t a r - g e a r s h a f t s were p r e c i s i o n ground to form t h e b e a r i n g i n n e r r a c e , and t h e s t a r g e a r s were c o u n t e r p h a s e d and n o n f a c t o r e d .

The g e a r r e d u c t i o n s y s t e m trs-vsmits i n e x c e s s o f t h e 2.74 MW (3675 hp) r e q u i f o r t h e QCGAT e n g i n e , and h a s been d e s i g n e d f o r l i f e g r e a t e r t h a ~ . 5000 h o u r s a t h i g h e r power.

The f a n s u p p o r t s t r u c t u r e ( f i g . 1 0 ) i n c l u d e s t h e f a n s u p p o r t h o u s i n g , i n t e r m e d i a t e case, and t h e e n g i n e s u p p o r t housing (main e n g i n e mount), a s w e l l a s t h e f a n gearbox and f a n i t s e l f . These cornponefits were d e s i g n e d to s u r v i v e a 1.8-kg (4-lb) b i r d s t r i k e a t a v e l o c i t y o f 250 k n o t s and t h e loss o f t w o a d j a c e n t f a n b l a d e s ( b u t n o t s i m u l t a n e o u s l y ) . F i n i t e - e l e m e n t s t r e s s a n a l y s e s were p e r - formed on t h e major s t r u c t u r a l p i e c e s f o r t h e l o a d s l i s t e d i n t a b l e 4. S t r e s s i s o p l e t h s and d i s p l a c e m e n t s a r e ~ i ~ o w n i n f i g - u r e 11.

The l o w - p r e s s u r e compressor, h i g h - p r e s s u r e compressor, and h i g h - p r e s s u r e t u r b i n e a r e s t a n d a r d components of t h e TFE731-3 e n g i n e and were used w i t h o u t d e s i g n changes. The d e s i g n - p o i n t c h a r a c t e r i s t i c s o f these components a r e l i s t e d g i v e n on t a b l e 5.

The LP t u r b i n e , which d r i v e s t h e f a n , and t h e l o w - p r e s s u r e compressor, is a 3 - s t a g e shrouded a x i a l d e s i g n . The QCGAT e n g i n e d e s i g n - p o i n t o p e r a t i n g c o n d i t i o n s a r e g i v e n i n t a b l e 6. S e v e r a l c r i t i c a l c o n s t r a i n t s were imposed on t h e d e s i g n o f t h e t u r b i n e .

S i n c e t h e QCGAT e n g i n e was b a s e d on t h e TFE731 c o r e , t h e o v e r r i d i n g ground r u l e was to minimize changes t o e x i s t i n g TFE731 hardware.

Because the QCGAT low-pressure turbine is larger in diameter and axially longer than that of the TFE731, it was necessary to design a gas flow path that would not cause disruption of airflow distri- bution in the combustor plenum. Location of the TFE731 aft turbine bearing was retained. The unusual shape of the third-stage disk (fig. 12) was the result of this latter constraint. Since LP spool speed is fixed ty the TFE731 LP compressor, the larger turbine represented a major design challenge from the standpoints of stress, vibration, blade flutter, life, and materials. In addi- tion, use of the 12-lobe compound mixer nozzle required low exit swirl angles. Total-to-total efficiency goal was set at 90 per- cent. As a result of these constraints, numerous compromises were necessary during design. Although it is not feasible to include the detailed results of all aerodynamic, thermodynamic, and mechanical design analyses in this report, all constraints were satisfied, including that of efficiency.

It was originally intended to use only a hydromechanical con- trol system for the QCGAT engine. However, because the hydro- mechanical unit is considered a backup system on the TFE731, it was decided to use a production TFE731 electronic control system as the primary control. The control (fig. 13) is a full-authority system providing speed control, over temperature, and overspeed protection under all operating conditions. These include start, transient, and steady state. A comparison of QCGAT engine char- acter istics and the TFE731-3 was made to determine if modifications were necessary to the existing computer. This comparison showed that the basic logic was satisfactory, and the adjustment ranges were adequate.

The QCGAT combustor (fig. 14) is a version of the TFE731 burner in production at the initiation of the program. In-house modifications for the TFE731 engine, which consisted of hole- pattern variations for smoke reduction, were incorporated in the QCGAT e n g i n e . During e n g i n e t e s t i n g , e m i s s i o n s were c o n t r o l l e d w i t h a system a d a p t e d from t h e NASA/AiResearch T1 P o l l u t i o n R e d u c t i o n Technology Program. Air was s u p p l i e d to t h e s e c o n d a r y f u e l n o z z l e s a t t h e t a x i - i d l e power s e t t i n g o n l y . T h i s a i d e d t h e f u e l atomiza- t i o n p r o c e s s (see f i g . 1 5 ) . A t a l l power s e t t i n g s e x c e p t t a x i - i d l e c o n d i t i o n , t h 2 f u e l was r e c o n n e c t e d to t h e s e c o n d a r y f u e l c i r c u i t .

An a i r - a s s i s t s y s t e m was n o t used. ( T h i s s y s t e m is d i s c u s s e d i n a s u b s e q u e n t p a p e r . ) Accessories f o r e u g i n e s l i k e QCGAT and t h e TFE731 n o r m a l l y c o n s i s t o f c u s t o m e r - f u r n i s h e d equipment. The a c c e s s o r y d r i v e g e a r - box, shown a t t h e bottom o f t h e e n g i r e i n f i g u r e 1 6 , p r o v i d e s mounting pads and d r i v e s on t h e f o r w a r d s i d e o f t h e g e a r b o x f o r a h y d r a u l i c pump or s i m i l a r equipment. These items n o t n o r m a l l y r e q u i r e d f o r a i r p l a n e s e r v i c e were n o t s u p p l i e d w i t h t h e QCGAT e n g i n e .

A starter- g e n e r a t o r was f u r n i s h e d , and a l t h o u g h n o t shown i n f i g u r e 1 6 , mounts on t h e pad o c c u p i e d by t h e l a b o r a t o r y a i r - t u r b i n e s t a r t e r .

QCGAT A 1 RPLANE DES I G N The a i r p l a n e s y n t h e s i z e d f o r t h e e n g i n e was based p r i m a r i l y on t h e L e a r j e t 35/36, a l t h o u g h it a l s o had minor f e a t u r e s found on o t h e r b v s i n e s s a i r p l a n e u s i n g TFE731 e n g i n e s . The major d i f f e r e n c e s

between t h e AiResearch QCGAT a i r p l a n e ( f i g . 1 7 ) and t h e Lear jet 35/36

are t h e e l o n g a t e d f u s e l a g e to i n c r e a s e payload ( p a s s e n g e r ) c a p a c i t y , a s l i g h t l y h i g h e r wing l o a d i n g , and t h e r e l o c a t i o n of t h e h o r i z o n t a l t a i l . The i n c r e a s e d p a y l o a d was p o s s i b l e b e c a u s e of t h e h i g h e r - t h r u s t e n g i n e s . The i n c r e a s e d wing l o a d i n g was t h e conse- quence o f t h e combined wing and f l a p c o n f i g u r a t i o n . The h o r i z o n t a l t a i l was moved t o a v o i d e n g i n e e x h a u s t . The a i r p l a n e d e f i n i t i o n had t w o p r i n c i p a l o b j e c t i v e s : F i r s t , to p r o v i d ? a n a i r p l a n e t h a t u t i l i z e d t h e i n s t a l l e d t h r u s t o f t h e QCGAT e n g i n e to p r o d u c e t a k e - o f f and a p p r o a c h f l i g h t p r o f i l e s f o r which n o i s e e s t i m a t e s c o u l d b e computed f o r s i d e l i n e , t a k e o f f , and a p p r o a c h FAR P a r t 36 measure- ment l o c a t i o n s shown i n f i g u r e 18. W i t h o u t a w e l l - d e f i n e d a i r p l a t ~ e c o n f i g u r a t i o n , it would n o t have b e e n p o s s i b l e t o make r e a l i s t i c and c o n s i s t e n t c o m p a r i s o n s o f i n - £ 1 i g h t n o i s e l e v e l s . The s e c o n d o b j e c t i v e was to r e p r e s e n t a v i a b l e a i r p l a n e w i t h respect to its a b i l i t y t o t r a n s p o r t p a s s e n g e r s and c a r g o w i t h a f u e l e f f i c i e n c y c o m p a r a b l e to c u r r e n t b u s i n e s s - jet a i r p l a n e . A t maximum t a k e o f f g r o s s w e i g h t of 8 , 6 7 4 kg ( ' 9 , 1 2 2 lb) , t h e 1 2 - p a s s e n g e r A i R e s e a r c h QCGAT a i r p l a n e t a k e s f u l l a d v a n t a g e o f t h e h i g h e r t h r u s t of t h e QCGAT e n g i n e , y e t meets t h e n o i s e g o a l s a t a l l t h r e e FAR P a r t 36 measurement l o c a t i o n s .

T a b l e 7 g i v e s t h e p r i n c i p a l a i r p l a n e d e s i g n parameters. As l i s t e d i n t h i s t a b l e , t h e wing i n c o r p o r a t e s d o u b l e - s l o t t e d f l a p s f o r good low-speed p e r f o r m a n c e . The r e l a t i v e l y h i g h wing l o a d i n g o f 354.5 kg/rn2 (72.6 l b / f t ) a s s u r e s a smooth r i d e c o m p a r a b l e to commercial jets.

The t a k e o f f p r o f i l e p r e s e n t e d i n f i g u r e 1 9 shows l i f t - o f f a f t e r a t a k e o f f r o l l of 914 m (3000 it) and, a t 6.48 km (3.5 nmi) from b r a k e r e l e a s e , a n a l t i t u d e of more t h a n 1 , 0 6 7 m (3500 f t ) w i t h t h r u s t c u t b a c k and a p p r o x i m a t e l y 1158 m (3800 f t ) w i t h f u l l t h r u s t . As i n d i c a t e d o n t h e p a y l o a d - r a n g e c h a r t , ( f i g . 2 0 ) , t h e QCGAT a i r p l a n e w i t h a maximum p a y l o a d o f 1 2 3 1 kg (2714 l b ) h a s a maximum r a n g e o f 3445 km (1860 n m i ) . T h i s would allow t h e a i r - p l a n e to f l y n o n - s t o p from P h o e n i x to N e w York C i t y a t a n a l t i t u d e o f 1524 m (5000 f t ) w i t h more t h a n 30 m i n u t e s r e s e r v e f u e l .

NACELLES D u r i n g p r e l i m i n a r y d e s i g n t a s k s , two n a c e l l e d e s i g n s w e r e selected; a f l i g h t n a c e l l e a n d a w o r k h o r s e n a c e l l e . O n l y t h e work- h o r s e n a c e l l e was c a r r i e d t h r o u g h t o d e t a i l d e s i g n a n d f a b r i c a t F o n .

The f l i g h t n a c e l l e was u s e d p r i m a r i l y t o l o o k a t a i r p l a n e i n s t e l l a - t i o n c h a r a c t e r i s t i c s a n d w e i g h t e s t i m a t e s .

The f l i g h t n a c e l l e ( f i g . 2 1 ) i n c o r p o r a t e d i n t e g r a l l y p h a s e d a c o u s t i c t r e h t m e n t i n t h e i n l e t b a r r e l , t h e i n n e r and o u t e r b y p a s s d u c t , and t h e a f t f a n d u c t . It a l - i n c o r p o r a t e d t h e e x t r a n o z z l e m i x i n g l e n o t h f o r t h e c o r e e x h a u s t n t i x e r . The w o r k h o r s e n a c e l l e e s s e n t i a l l y d u p l i c a t e d t h e i n t e r n a l a e r o d y n a m i c d e s i g n a n d a c o c - s t i c a l t r e a t m e n t o f t h e f l i g h t n a c e l l e e x c e p t f o r a s e c t i o n i n t h e a r e a i m m e d i a t e l y a f t o f t h e f a n t.hat had no a c o ~ s t i c t r e a t m e n t i n t h e f l i g h t n a c e l l e . The w e i g h t o f t h e f l i g h t n a c e l l e was e s t i m a t e d a t 1 3 4 kg (295 l b ) . The t o t a l i n s t a l l e d p r o p u l s i o n s y s t e m w e i g h t was e s t i m a t e d a t 513 kg (1130 l b ) .

A c r o s s s e c t i o n o f t h e w o r k h o r s e n a c e l l e is shown w i t h t h e e n g i n e i n f i g u r e 22. T h i s n a c e l l e was d e s i g n e d t o p r o v i d e maximum t e s t c o n f i g u r a t i o n v e r s a t i l i t y f o r t h e QCGAT e n g i n e . F i g u r e 22 a l s o s h o w s t h e b a s i c component a r r a n g e m e n t s . The p r i n c i p a l compo- n e n t s i n c l u d e t h e i n l e t b a r r e l , t h a t accommodates a f l i g h t - s i m u l a t o r l i p , a c o n v e n t i o n a l l y s h a p e d n a c e l l e l i p , t h e i n n e r a n d o u t e r b y p a s s d u c t s l o c a t e d o p p o s i t e t h e e n g i n e h o t s ~ c t i c n , t h e a f t b a r r e l , t h e c o r e m i x e r , and t h e n o z z l e .

The i n l e t b a r r e l ( f i g . 2 3 ) i n c o r p o r a t e s two 5 LS o f i n t e r - c h a n g e a b l e d u c t l i n e r s - - o n e set o f a c o u s t i c - t r e ~ ' i a r c ! i t p a n e l s a n d o n e s e t o f h a r d w a l l p a n e l s , a s well a s t n e two d i f f e r e n t i n l e t l i p s . The f l i g h t - s i m u l a t o r l i p ( f i g . 23) i s d e s i g n e d t o c o n t r o l a n d d i r e c t t h e i n l e t a i r f l o w , t h u s s i m u l a t i n g a c t u a l f l i g h t c o n d i - L i o n s . The c o n v e n t i o n a l n a c e l l e l i p is i n s t a l l e d on t h e e n g i n e a s shown i n f i g u r e 2 4 . The i n l e t b a r r e l was d e s i a n e d f o r h i g h - i n l e t r e c o v e r y a t a r e l a t i v e l y h i g h - t h r o a t Mach number o f 0.73 a t c r u i s e ( f i g . 2 5 ) . When t h e i n l e t b a r r e l is removed, a r e f e r e n c e bellmouth assembly can be i n s t a l l e d d i r a c t l y on t h e e n g i n e i n l e t f l a n g e .

D e t a i l e d performance tests were conducted w i t h t h e b e l l r r o z t h and w i l l be d i s c u s s e d l a t e r .

The i n n e r and o u t e r b y p a s s d u c t s e c t i o n ( f i g . 26) a l s o i n c o r - p o r a t e d two sets of d u c t liners--acoustical-treatnent and h a r d w a l l p a n e l s . A s i n t h e i n l e t b a r r e l , t h e s e r e p l a c e a b l e p a n e l s were i n 180-degrez s e c t i o n s and were r a d i a l l y a d j u s t a b l e so t h a t t h e flow- p a t h c o n t i n u i t y c o u l d be c o n t r o l l e d . The o u t e r bypass d u c t con- t a i n e a a f a i r e d s e r v i c e s t r u t t h a t p r o v i d e d f o r e x t e n s i v e p r e s s u r e and t e m p e r a t u r e i n s t r u m e n t a t i o n , a s w e l l 3s s u p p o r t of t h e a f t sec- t i o n of t h e e n g i n e . The a f t f l a n g e o f t h e o u t e r bypass d u c t was common to t w o n o z z l e schemes--the mixer compound n o z z l e and t h e c o a n n u l a r n o z z l e . F i g u r e 27 shows h a l f t h e o u t e r b y p a s s d u c t sec- t i o n removed. The s e r v i c e s t r u t is v i s i b l e , and t h e c o r e s e c t i o n of t h e c o a n n u l a r n o z z l e is i n s t a l l e d .

A 1 2 - l ~ b e core mixer ( f i g . 2 8 ) was d e s i g n e d f o r t h e AiReseiirch QCGAT e n g i n e to improve b o t h performance and t a k e o f f n o i s e . W i t h t h e mixer compound noz;:ie, a 1-percen t TSFC improvement i n s e a - ''.eve1 performance was denionstrated. A 3 - 2 - p e r c e n t TSFC improvement t c r u i s e was e s t i m a t e d b a s e d on mixer inodel and e n g i n e tests. A 3- t o 5-EPNdB r e d u c t i o n i n t a k e o f f n o i s e from the c o a n n u l a r c o n f i : - a - t i o n was a c h i e v e d w i t h t h e mixer compound n o z z l e . A s shown i n f i g u r e 29, smoke t r a c e s on t h e mixer c e n t e r b o d y i n d i c a t e d t h a c t h e mixer compound n o z z l e was p e r f o r m i n g a s p r e d i c t e d . S i m i i a r smoke t z a c e s were o b s e r v e d i n t h e n o z z l e s e c t i o n downstream of t h e mixer.

The f i n a l s e c t i o n s o f t h e workhorse n a c e l l e assembly ( f i g . 3 0 ) a r e t h e a f t b a r r e l , w h i c h h a s h a r d w a l l and a c c : i s t i c pane's, and t h e nozzle. These sections are used only when the mixer is installed.

They are removed when the coannular nozzle system is used.

The complete workhorse nacelle assembly is shown in figures and 32. These figures show the engine mounted on the test stand at AiResearch's remote desert test facility in the San Tan mountains, southeast of Phoenix.

The following points sulmarize the design of the AiResearch QCGAT ensine and nacelle cvstem: o An existing turbofan engine =re was utilized for an experimental demonstrator engine. This was a requirement of the original problem statement and was particularly important with respect to minimizing costs and maximizing reliability.

o Several unique components were successf ully adapted to this core: fan, gearbox, combustor, low-pressure tur- bine, and associated structure. These components formed the basis for meeting the main program objective demon- strating the application of large turbofan enqine design, emissions, and noise technology in small general aviation turbofans.

o A highly versatile workhorse nacelle incorporating interchangeable acoustic and hardwall duct liners, showed that large-engine attenuation technology could be applied to small propulsion engines. The application of the mixer compound nozzle demonstrated both performance and noise advantages on t t , e engine.

The QCGAT pyogram ~ a 3 e several significant contributions to general aviation propulsion: o Application of exhaust-emissions reduction techniques.

1 . Hydrocarbon and carbon monoxide goals were m e t .

2 . Nitroqen oxides were greatly reduced.

o With the aid of NASA, improved small engine noise- analysis techniques, including core noise and static-to- flight correlations, were developed.

o Major noise reduction, beyond that of an already quiet engine, was abtained. The AiResearch QCGAT engine is significantly quieter than any other business jet engine.

TABLE 1 . ATRESEARCH QCGAT ENGINE, TECHNICAL GOALS, Thrust TSFC 1

N kg/N . h

I A. Performance (lbf 1 (lbm/hr/lbf )

Takeoff (SLS, ISb)

I I

o Uninstalled

I

o Installed

I

Cruise

I

(12,192 m (40,000 ft), H = 0.81

I

o Uninstalled o Installed (with mixer nozzle) 4,017 0.0759 (903) (0.744) B . Noise (FAR Part 36) EPNdB Takeoff 73.3

I I

Sideline 82.1 1

I

Approach 87.3 C. Emissions (EPA 1979 Standards T-1) EPAP I

Hydrocarbon (HC) 1 . 6 1

I

Carbon Monoxide (CO) 9 . 4 1

I

Oxides of Nitrogen (NOx)

I

Smonc Number 38.0 E . Life D . Weight hr kg (lbm) 10,000 ( 8 3 5 ) TABLE 2. QCGAT CYCLE PARAMETERS.

. . . . . . . . . . . . . .

Wsign point 12,192 r (40,000 f t) .

U = 0.8, ISA

. . . . . . . . . . . . . . . . . Thrust 4,017 N (902 1bf)-

installed

I I

. . . . . . . . . . . . . . . . . . I TSFC 0.0759 kg/N.h

(0.744 lk/hr;lbf)

I I

. . . . . . . . . . . . . . / Bypass ratio 3.71

I

. . . . . . . . . . . Fan pressure ratio 1.62

I I

1 Cycle pressure ratio . - . . - . - - - . 17 -7

I

. . . . . . . Turbine inlet temperature 1,266K

(1,820.F)

i

i

. . . . . . . . . i Corrected fan aizflov 77.8 kg/s

(171.6 lb/s-ci

Ccrrected core airflow . . . . . . . . . 11.5 kg/s

(25.4 lb/stc) 1

TABLE 3 . QCGAT FAN DESIGN FEATURES.

i At Design Point--12.192 a (40.000 ft, C.8H. XSA).

!

/ Diameter . . . . . . . . . . . . 77 - 5 ca (30.5 in.)

I

i Radius ratio . . . . . . . . . . 0.46

'Znlet corrected airflow . . . . 77.8 kg/s (171.6 1b.isec)

i

/ Bypass ratia . . . . . . . . 3 . 7

I Bypass press.~re ratio . . . . . 1.62

i

Core pressure ratio . . . . . . 1 .SS

. . Inlet tip relative Pach No. 1.39

. . . Inlet corrected tip speed 6.985m/s (1375 ft/sec)

AIRESEARCH QCGAT ENGINE PERFORMANCE AND EMISSIONS TESTS William M. Norgren AiResearch Manufacturing Company of Arizona A Division of The Garrett Corporation A Q u i e t , C l e a n , G e n e r a l A v i a t i o n Turbofan (QCGAT) e n g i n e and n a c e l l e s y s t e m was d e s i g n e d and t e s t e d by t h e AiResearch Manufac- t u r i n g Company o f Arizona under C o n t r a c t to t h e NASA Lewis R e s e a r c h C e n t e r , The e n g i n e u t i l i z e d t h e core o f AiResearch Model TFE731-3 e n g i n e and i n c o r p o r a t e d numerous noise and e m i s s i o n s r e d u c t i o n f e a t u r e s . Endurance, performance, and e m i s s i o n s tests were con- d u c t e d o n t h e e n g i n e p r i o r t o t h e a c h u s t i c test sequence, T e s t r e s u l t s proved t h a t t h e e n g i n e m e t most o f t h e d e s i g n g o a l s , and a teardown i n s p e c t i o n o f t h e e n g i n e f o l l o w i n g t h e tests showed t h e u n i t to. be i n e x c e l ~ 2 n t c o n d i t i o n .

INTRODUCTION Performance and e m i s s i o n tests were conducted on a s p e c i a l l y d e s i g n e d AiResearch QCGAT e n g i n e i n t h e 17,793-N (4,000-lb) t h r u s t class. T e s t i n g i n c l u d e d aerodynamic performance, e m i s s i o n t e s t i n g , and a c o u s t i c tests. T h i s p a p e r d i s c u s s e s t h e performance and emis- s i o n s tests and i n s p e c t i o n r e s u l t s o f t h o s e tests.

Due to t h e r e q u i r e m e n t to perform a complex series o f a c o u s t i c tests, a s well a s performance and e m i s s i o n s tests, t w o s e p a r a t e test a r e a s were used. Most o f t h e f u l l y i n s t r u m e n t e d performance test- ing was conducted i n t h e Phoenix development and q u a l i f i c a t i o n test cells shown i n f i g u r e 1. Another s e r i e s o f performance c o m p a r i s o n s were run a t t h e AiResearch San Tan remote test s i t e ( f i g . 2 ) to e s t a b l i s h a b a s e l i n e f o r t h e s u b s e q u e n t a c o u s t i c tests.

The test sequence was set up to e n s u r e t h e s t r u c t u r a l i n t e g - r i t y o f t h e e n g i n e and to o b t a i n b a s e l i n e p e r f o r m a n c e i n b o t h a c o u s t i c and h a r d w a l l i n s t a l l a t i o n c o n f i g u r a t i o n s , By working around t h e clock, t h e t e s t i n g p h a s e was compressed i n t o s i x weeks.

The eng i n e was s u b s e q u e n t l y r e f u r b i s h e d , a c c e p t a n c e tested, and d e l i v e r e d on s c h e d u l e . F i g u r e 3 o u t l i n e s t h e AiResearch test s c h e d u l e . Scheduled d a t e s were met w i t h t h e c o o p e r a t i o n o f t h e w e a t h e r , b u t more s i g n i f i c a n t l y , w i t h t h e e x c e l l e n t s u p p o r t and r e s p o n s e AiResearch r e c e i v e d from t h e NASA e n g i n e e r i n g s t a f f .

The f i r s t r u n o f any new a i r p l a n e e n g i n e is r e f e r r e d to as a " g r e e n r u n m , A g r e e n r u n is a p r e l i m i n a r y test. to d e t e r m i n e how w e l l t h e u n i t r u n s , and to d e t e r m i n e p o t e n t i a l problem areas. It also e s t a b l i s h e s normal v a l u e s for v i b r a t i o n , o i l p r e s s u r e , temper- a t u r e s , etc. On c o m p l e t i o n o f t h e QCGAT g r e e n run, t h e e n g i n e w a s c o m p l e t e l y disassembled, i n s p e c t e d , r e a s s e m b l e d , and c y c l e d i n t o a 40-hour e n d u r a n c e test p r i o r to b e g i n n i n g performance and a c o u s t i c t e s t i n g .

The e n d u r a n c e c y c l e (table 1) was i n t e n d e d to d u p l i c a t e t h e c o n d i t i o n s o f a jet c y c l e w h i l e wearing i n t h e e n g i n e . Approxi- m a t e l y 40 h o u r s were r u n to wear i n t h e seals, b e a r i n g s , etc. T h i s p r o v i d e d performance and e n g i n e c o n d i t i o n s r e p r e s e n t a t i v e o f a typLca 1 eng i n e .

TEST OBJECTIVES The p r i m a r y o b j e c t i v e s o f t h e QCGAT t e s t program were to demon- strate t h e e n g i n e c a p a b i l i t i e s r e q u i r e d t o meet t h e program g o a l s , to prove t h e s t r u c t u r a l i n t e g r i t y , and to measure e n g i n e perform- ance, e m i s s i o n , and a c o u s t i c c h a r a c t e r i s t i c s . The series of tests i n c l u d e d o p e r a t i o n w i t h v a r i o u s c o m b i n a t i o n s o f i n l e t s , t h r u s t n o z z l e s , and a c o u s t i c t r e a t m e n t s . Table 2 lists t h e performance g o a l s f o r t h e QCGAT sngine.

The 1979 e m i s s i o n g o a l s s e t by t h e EPA i n 1973 f o r t h e c l a s s T1 e n g i n e s are l i s t e d i n t a b l e 3. These s t a n d a r d s have s i n c e been dropped by t h e EPA, b u t were m a i n t a i n e d a s QCGAT program g o a l s .

The EPA p a r a m e t e r (EPAP) is d e t e r m i n e d from e m i s s i o n s measurements made a t f o u r power s e t t i n g s and t h e n added t o g e t h e r . T h e time weighing f a c t o r ( t a b l e 4) used i n t h i s c a l c u l a t i o n is d e r i v e d f ~ o m t h e time e s t a b l i s h e d by EPA as being t h e t y p i c a l time s p e n t i n e a c h o p e r a t i n g mode f o r an a i r p l a n e w i t h T1 Class e n g i n e s .

The smoke s t a n d a r d is e s t a b l i s h e d a s a f u n c t i o n o f r a t e d e n g i n e power and a p p r o x i m a t e l y r e p r e s e n t s t h e t h r e s h o l d f o r v i s i b l e smoke from an e n g i n e e x h a u s t . The s t a n d a r d is e x p r e s s e d as Smoke

Number ( S N ) , and is a f u n c t i o n o f t h e amount o f l i g h t r e f l e c t e d

from a sample o f p a r t i c u l a t e c o l l e c t e d on a p i e c e o f f i l t e r p a p e r exposed to t h e e n g i n e e x h a u s t . The h i g h e r t h e S N , t h e g r e a t e r t h e amount o f p a r t i c u l a t e s ; h e n c e , t h e g r e a t e r t h e smoke v i s i b i l i t y .

Smoke measurements were made a t t h e same f o u r p c r e r s e t t i n g s a s t h e g a s e o u s e m i s s i o n test. The h i g h e s t SN o f t h e f o u r power s e t t i n g s was c o n s i d e r e d t h e smoke number f o r t h e e n g i n e .

AERODYNAMIC PERFORMANCE A f u l l y i n s t r u m e n t e d e n g i n e was i n s t a l l e d i n t h e Phoenix development and q u a l i f i c a t i o n test cell. F i g u r e 4 shows t h e e n g i n e w i t h o u t t h e i n l e t a t t a c b . e d . F i g u r e 5 shows t h e e n g i n e w i t h a c a l i b r a t e d bellmouth. The f i r s t tests were r u n w i t h a c o a n n u l a r n o z z l e ( f i g . 6 ) to e s t a b l i s h b a s e l i n e performance a g a i n s t which t h e I n t o t a l , s e v e n p e r - mixer compound n o z z l e ( f i g . 7 ) c o u l d b e compared.

formance c a l i b r a t i o n s were made ( t a b l e 5 ) . A s t h e test sequence pro- g r e s s e d , t h e c o a n n u l a r n o z z l e was r e p l a c e d w i t h t h e mixer compound n o z z l e . The s u b s e q u e n t c o m b i n a t i o n s c a l i b r a t e d t h e f l i g h t simu- l a t o r l i p and n a c e l l e l i p to t h e c o a n n u l a r n o z z l e and mixer com- pound n o z z l e s , r e s p e c t i v e l y . B e f o r e f i n a l c a l i b r a t i o n , t h e e n g i n e was removed from t h e test c e l l , and t h e h a r d w a l l f a n d u c t was r e p l a c e d w i t h the a c o u s t i c f a n d u c t . S i n c e t h e f a n d u c t c o n t a i n s most o f t h e a c c e s s o r i e s and plumbing, t h i s became a r e l a t i v e l y major change. The e n g i n e was r e i n s t a l l e d and f i n a l performance c a l i b r a t i o n was run.

A c o u s t i c t e s t i n g and f i n a l a c c e p t a n c e tests were t h e n begun on t h e e n g i n e . As measured, e n g i n e performance was found to be close to what had been e x p e c t e d . W i t h t h e e x c e p t i o n o f t h e f a n , t h e new components met or exceeded t h e i r e s t i m a t e d performance. A s a n t i c i - p a t e d , t h e mixer compound n o z z l e p r o v i d e d a s i g n i f i c a n t improvement t o t h e e n g i n e . T a b l e 6 shows t h e r e s u l t s o f f o u r o f t h e c o n f i g u r a - t i o n s compared a t a c o n s t a n t l o w - p r e s s u r e r o t o r s p e e d (N1) o f 1938 r a d / s (18,510 rpm) .

Performance C a l i b r a t i o n 2 - Using t h e mixer compound n o z z l e , t h i s c a i i b r a t i a n r e s u l t e d i n a s i g n i f i c a n t i n c r e a s e i n a i r f l o w and t h r u s t a t a c o n s t a n t N1. The mixer compound n o z z l e h a s a bypass s t r e a m a r e a t h a t is e f f e c t i v e l y much l a r g e r t h a n t h e c o a n n u l a r n o z z l e . T h i s p r o v i d e d a rematch o f t h e f a n to a h i g h e r e f f i c i e n c y and flow. The core s t r e a m a r e a is e f f e c t i v e l y s m a l l e r t h a n t h e c o a n n u l a r n o z z l e and c a u s e d a g r e a t e r l o w - p r e s s u r e (LP) t u r b i n e d i s c h a r g e p r e s s u r e . The e n g i n e had a g r e a t e r h i g h - p r e s s u r e (HP) t u r b i n e d i s c h a r g e t e m p e r a t u r e because of t h e i n c r e a s e d t o t a l a i r f l o w , t h u s r e q u i r i n g more power from t h e LP t u r b i n e . T h i s i n c r e a s e d power was s u p p l i e d by i n c r e a s i n g t h e t u r b i n e - i n l e t tem- p e r a t u r e , r e s u l t i n g i n a h i g h e r HP r o t o r speed (N2) and compressor d i s c h a r g e p r e s s u r e ( P t g ) . T h e i n c r e a s e d t h r u s t r e s u l t e d p r i n c i - p a l l y from t h e i n c r e a s e d a i r f l o w .

Performance C a l i b r a t i o n 5 - Using t h e n a c e l l e - l i p i n l e t w i t h

t h e mixer compound n o z z l e , t h e e n g i n e performance i f t h r u s t , TSFC, etc.) was s i m i l a r t o performance c a l i b r a t i o n 2 , w h i c h a l s o used t h e mixer compound n o z z l e .

Performance C a l i b r a t i o n 7 - Using t h e n a c e l l e - l i p i n l e t , t h e

mixer compound n o z z l e , and f u l l a c o u s t i c t r e a t m e n t i n t h e bypass d u c t , the a c o u s t i c t r e a t m e n t had l i t t l e e f f e c t on t h e performance of the engine as compared to calibration 5. Sim'lar tests con- firmed this conclusion.

Table 7 shows two engine configurations compared with the pre- test analytical model. Thrilst, airflow, and a high-rotor speed approximated the model parameters; however, fuel flow, TSFC, and turbine discharge temperature (TtQ Z ) were 8iscrepant. Analysis of this and other data showed that at maxinum sea level static thrust, the fan was lower than predicted in efficiency and in airflow.

This characteristic is typical of most fans in this size class wherein compromises in aerodynamic configurations imposed by design for bird strike cause unfavorable airfoil loadings with consequent decrease in efficiency and airflow capacity.

COMPARISON TO AERODYNAMIC GOALS Table 8 is a comparison of the tested engine performance to the QCGAT program goals. The largest difference occurred on the uninstalled engine where the fan performance, as well as a one percent lower than estimated thrust coefficient for the coannular nozzle, resulted in a specific fuel consumption slightly over the estimate.

When the nacelle was inst- lled, including the mixer nozzle, the sea level static TSFt is seen to be 1.4 percent over the goal.

In this case, a comparison of the engine tested performance versus the analytical model showed that the mixer nozzle exceeded the estimate, while the fan performance was below the estimate.

Extrapolation of the tested data to the altitude cruise condi- tion shows that the cruise TSFC would be below the estimated level.

Since the majority of the mission fuel is consumed at cruise, it is concluded that the program fuel consumption goals were achieved and that QCGAT has demonstrated a significant advancement in engine e~ficiency.

EMISSIONS TEST Work on t h e combustion s y s t e m < ? s i g n o f t h e AiResearch QCGAT e n g i n e was conducted under s e p a r a t e contra:t f o r t h e T-1 combustor, i n i t i a l l y selected f o r t h e program, Howevx, s c h e d u l e incompat- i b i l i t i e s p r e v e n t e d d i r e c t i n c o r p o r a t i o n o f t h e T-1 combustor i n t h e program and a n i n t e r i m d e s i g n was used.

The combustor l i n e r used i n t h e QCGAT tests ( f i g . 8) was a m o d i f i c a t i o n o f t h e production TFE731 b u r n e r , These m C i f i c a t i o n s c o n s i s t e d c f sei era1 v a r i a t i o n s , and i n c l u d e d punched v e r s u s p i e r c e d h o l e s . D i f f e r e n t h o l e l o c a t i o n s and s i z e s were i n c o r p o r - a t e d f o r smoke number r e d u c t i o n . The a c t u a l b u r n e r used i n t h e test was a n e x p e r i m e n t a l i n t e r i m d e s i g n , A s a r e s u l t , t h e ternsera- t u r e p a t t e r n f a c t o r was h i g h e r t h a n d e s i r e d d u r i n g e a r l y t e s t i n g .

T h i s c o n d i t i o n was c o r r e c t e d on l a t e r b u r n e r s .

C o n t r o l of t h e g a s e o u s e m i s s i o n s a t i d l e w a s accomplished by s u p p l y i n g air to t h e s e c o n d a r y a t o m i z e r s o f ' ae f u e l n o z z l e s . T h i s a i r improved e m i s s i o n s t w o ways: I t c a u s e d a l l o f t h e f u e l to p a s s through t h e primary n o z z l e i n s t e a d o f a l l o w i n g a small p o r t i o n o f f u e l to f l o w o u t o f t h e s e c o n d a r i e s . The a i r also improved t h e v a p o r i z a t i o n o f t h e f u e l coming o u t o f t h e p r i m a r y a t o m i z e r .

F i g u r e 9 d e p i c t s t h e c o a b u s t o r l i r a s s i s t system. A i r f o r t h e a s s i s t system was p r o v i d e d f r c m a l a b o r a t o r y system t h a t a p p r o x i - mated t h e characteristics o f e n g i n e s u p p l y a i r . The a i r was pro- v i d e d a t a p r e s s u r e and t e m p e r a t u r e t h a t s i m u l a t e d compressor b l e e d a i r , and was c o o l e d w i t h a s i m p l e a i r - t o - a i r h e a t exchanger i n t h e f a n d u c t , The a i r was s u p p l i e d from a l a b o r a t o r y compressed-air s o u r c e w i t h a s u p p l y p r e s s u r e of 1 4 . 4 k?A ( 3 0 0 p s i g ) . A f t e r p a s s i n g through a 20-micron f i l t e r , t h e a i r was h e a t e d by an electric heater to between 366K (200°F) and 422K ( 3 0 0 a F ) . T h i s s i m u l a t e d a n a i r a s s i - system where t h e d i s c h a r g e t e m p e r a t u r e from t h e h e a t of c o m p r e s s i o n f o r t h e assist a i r would b e s i m i l a r to a i r e x t r a c t e d tror t h e b o o s t compressor. The air t h e n p a s s e d t h r o u g h a f l o w m e a s u r i n g s e c t i o n a n d was i n t r o d u c e d to t h e s e c o n d a r y f u e l l i n e .

F o r t h i s test, t h e l i n e was d i s c o n n e c t e d f r o m t h e f l o w d i v i d e r a n d t h e f l o w d i v i d e r p a t h c a p p e d . h s c h e m a t i c o f t h i s s y s t e m is shown i n C i g u r e 12.

E m i s s i o n s were collected f o r measurement w i t h a 24-element probe s i m i l a r to t h e o n e shown i n f i g u r e 11. Measurements w e r e t a k e n o n l y wFth t h e c o a n n u l a r n o z z l e s i n c e t h e r e was n o s t a n d a r d t e c h n i q u e of m e a s u r i n q e s t a b l i s h e d f o r t h e m i x e r compound n o z z l e .

The HC a n d CO goals were met by u s i n g a n a i r a s s i s t i n l e t p r e s s u r e o f 5.027 kPa ( 1 0 5 p s i d ) a n d a t e m p e r a t u r e o f 389K (240°F) a t t a x i idle. The r e s u l t s a r e p r e s e n t e d i n t a b l e 9. T h i s p r e s s u r e a n d t e m p e r a t u r e is r e l a t i v e l y e a s y t o o b t a i n w i t h a boost com- p r e s s o r o n a n a i r c r a f t e n q i n e . Lower a i r - a s s i s t p r e s s u r e would h a v e r e s u l t e d i n h i g h e r e m i s s i o n i n d e x v a l u e s ( i - e . , g/kg f u e l ) f o r b o t h HC a n d CO. Sicce more t h a n 90 p e r c e n t o f t h e HC a n d CO EPAP v a l u e s are c o n t r i b u t e d by t h e t a x i - i d l e t e r m s , small c h a n g e s i n H C a n d CO e m i s s i o n i n d e x v a l u e s a t t h a t power s e t t i n g r e s u l t e d i n s i g n i f i c a n t c h a n g e s i n t h e o v e r a l l EPAP v a l u e s f o r t h e t w o p o l l u t a n t s .

The CO a n d HC e m i s s i o n s m e t t h e g o a l s a n d NO, w a s s i g n i f i- c a n t l y r e d u c e d , b u t s l i g h t l y a b o v e g o a l . The smoke number was also a b o v e g o a l . However, t h e e n g i n e showed no s i g n o f v i s i b l e s m o k e w h i l e o p e r a t i n g a t t h e test p o i n t i n s e v e r a l tests.

TEARDOWN INSPECTION A f t e r c o m p l e t i o n o f a l l tests, t h e e n q i n e was c o m p l e t e l y d i s - a s s e m b l e d , i n s p e c t e d , a n d r e f u r b i s h e d p r i o r t o s h i p m e n t t o NASA.

With a l m o s t 70 a c c u m u l a t e d h o u r s of t e s t i n g i n c l u d i n g 70 s t a r t s , t h e m a j o r i t y o f arts w e r e i n e x c e l l e n t c o n d i t i o n a n d o n l y t h r e e c o m p o n e n t s showed a n y u n u s u a l s i g n s o f w e a r . A s i n g l e s u n - g e a r tooth had developed a small pit as shorn by the a r r w under mag- nifrcation in figure 12, This was later found to be the result of a flaw in the basic saterial f r o l i l which the part was constructed.

The wear pattern was judged to be good and commensurate with the time and load on the gear system.

The second discrepancy was microscopic surface cracks radiat- ing from a couple of the special instrumentation bosses (see arrow) of the turbine plenum shown in figure 13. These were the results of torch brazing the HP compressor discharge total-pressure probes into the plenum after the part had conpleted the normal stress- relieving process, This is a problem that is unique to the highly instrumented test engine and wouid not appear on production-type plenums, The third problem noted was a crack in the surface of one HP turbine cooled stator wane (figure 14). This crack resulted from a single hot streak in the engine. This was the result of using the experimental low-smoke zombustion liner that had not been suffi- ciently developed at the time this test was run. This character- istic was subsequently corrected, and later production low-smoke combustor liners did not exhibit a hot streak.

All three of the problems found during teardown inspection were determined to be the result of outside factors and not the result of design deficiency. The basic engine design fulfilled design requirements. All AiResearch QCGAT engine discrepancies were removed prior lo shipment to NASA.

TECHNICAL ACCOM PL I SHMENTS The technical accomplishments demonstrated by the AiResearch QCGAT test program are numerous, Most important is :he fact that the engine met the design goals in almost every case ( i . e . , thrust, TSFC, emissions. etc.). Performance was slightly better than Fre- dicted for the installed configuration with the m i x e r compound

n o z z l e a t t h e d e s i g n p o i n t o f 12,192 r (40,000 f t ) , 0.8 Mach

nuaber .

Performance o f t h e AiResearch QCGAT e n g i n e was e x c e l l e n t throughout a l l t e s t i n g . No s e r i o u s mechanical malf u n c t i o n s were encounterei:, and no s i g n i f i c a n t t e s t t i m e was lost due to engine- related problems. m i s s i o n s were d r a s t i c a l l y reduced o v e r similar e n g i n e s , and t h e e n g i n e e x h i b i t e d good smoke performance, The t e s t i n g o f t h e AiResearch QCGAT e n g i n e p r o v i d e d e v i d e n c e o f t h e e n g i n e r e l i a b i l i t y and performance. A f t e r 82 h o u r s and 77 s t a r t s t h e u n i t remained t r o u b l e - f r e e . The few problems encountered were r o s t l y a s s o c i a t e d with l a b o r a t o r y or cell equip- ment. Engine performance remained s a t i s f a c t o r y w i t h v e r y l i t t l e d e g r a d a t i o n as t h e u n i t accumulated t i m e - Though t h e LP t u r b i n e d i d n o t have t h e b e n e f i t o f r i g testing, it proved to meet d e s i g n g o a l s f o r t h e engine. S i m i l a r l y , t h e f u l l - scale mixer compound n o z z l e was found to perform better t h a n a n t i c i p a t e d .

CONCLUSION A s shown by t h e test program, t h e AiResearch QCGAT e n g i n e met almost a l l o f t h e program g o a l s . T h i s is g r a p h i c e v i d e n c e t h a t t h e a p p l i c a t i o n o f l a r g e engine a c o u s t i c t e c h t ~ o l o g y to s m a l l e n g i n e s a s w e l l a s t h e a p p l i c a t i o n of s p e c i a l i z e d s m a l l e n g i n e t e c h n o l o g i e s can r e s u l t i n low-noise, low-emissions, and reduced f u e l consumption g e n e r a l a v i a t i o n t u r b o f a n e n g i n e s .

TABLE 1 . QCGAT ENDURANCE TEST CYCLE, f Cycle Tire Condition bin. )

-

Start Idle Takeoff 5 Max. Continuous 10 Max. Cruise Idle 5 75% Max. Cruise 5 Idle Approach 5 Idle 5 Shutdom 15 Total 1 hr 45 min.

23 Cycles = total run time of 34.5 h : TABLE 2. ENGINE PERFORMANCE GOALS,

-

Goa 1s Thrust TSX N kg/N. h Condition (lbf) (lbm/hr/lbf Takeoff, Sea Level Static, f tandard Day: o Uninstalled 17,513 0.0426 (3,937 (0.418) o With ground test 17,312 0.0431 nacelle and acoustic (3,892) ( 0 . 4 : 3 ) treatment and mixer mapound nozzle Design Cruise, 12,192-m J40,OOO-ft) Altitude, 0 . 8 nach Number : o Uninstalled 3,954 0.0775 (889 (0.760) o with ground test 4,017 0.0759 nacelle and acoustic 1903) (0.744) treatment and mixer compound nozzle TABLE 3 . MISSIONS PROGRAM GOALS.

s EPAPS Program ,=l, kg/4448 N-h/c1 ' e (lbr/1000 lbf-hr/cycle Pollutant tmburaed I Z y ~ ~ (XC) 0.73 (1.6)

Carboa Non~xidc (a) 4.26 (9.4)

OriQs of Uitrogem (lJOX) 1.68 (3.7) SDoke 38* T A B U 4, EMISSIONS CYCLE.

Percent Time Rated Minutes power no& 19.0 ~exi-idle Taxi-out Takaof f 2 - 5 ClLbout 4.5 Approach 7.0 m x i - i d l e Taxi-in Total 3 3 . 5 i TABLE 5 . PERFORMANCE CALIBRATIONS AND ENGINE CONFIGURATIONS,

-

calibzation Descr! ptim No.

B e l l south and Coannular Nozzle .L B e l l m u t h and Mixer Caopound Nozzle Flight-Simulator Lip and Coannular Nozzle Nacelle Lip and Coannular Nozzle Nacelle Lip and Mixer Canpound Nozzle Filght-Simulator Lip and Mixer Campound Nozzle 7 Flight-Simulator Lip, Mixer Compound Nozzle and Acoustically Treated m c t s TABLE 6 . QCGAT TEST RESULTS.

It by Test Number

1 Parameter

Acoustic Treatment Pane 1

1 7

~ a c e l l e I ~iaulator Inlet Conf iguration l~cllmoutt.

Mixer Exhaust Configuration \mnnular Mixer

I

o Thrust. N(1bf) 15,413 (3.4651 o T S K . kq/N.h 0.0457

1 (

(lhihr/lbf 1 (0.448) I 3.011 o Biqh rotor speed ! N2. cad/+ (rpm) (28.760)

I i

1 a HP turbine discharge! ? ,105

teaperature Tt4-2. (1,530)

1 F [*F) I

! o n t a l airflow.

I kqis ( lbm/sec) TABLE 7 , TEST RESULTS COMPARED TO ANALYTICAL MODEL, Coannular Hirer Compound

I Nozzles Nozzle

i

1 parameter Mode 1 Test I W e 1 Test

I I

H ~ q h Rotor Speed 12, 3.024 1 3.061 1 2.970 3,024 1

:ad/s (rpia) ! 8 8 2 9 , 2 4 0 2 (28.880)

1 I i I Fan Nozzle Inlet 327 1 330 I 322 324 1

i

Teraperature Tt17, , (129.6) , (135.0) (119.6) (124.0)

K (OF) ! I I

FanNoozlcTotal 1 14.58 1 14.60 i 14.38 14.09 1

Presaure Pt17, (21.11) (21.18) / (20.85) , (20.42)

N / C r n ' (PSI ) I

I

Engine Total Airflow ' 65.6 1 65.3 62.3 63.5 , K T , k l s e j (144.6) ( 1 3 7 4 ) , (140.1b ,

, (143.9)

TABLE 8. QCGAT TEST RESULTS VERSUS PERFORMANCE GOALS.

THRUST, N TSFC, kg/N-h 8 ( l b f (lbm/hr/lbf 8 A A F l i g h t C o n d i t i o n Goal T e s t Coal Test Sea l e v e l , static, 0 17,513 17,513 0.0426 0.0459 +7.7 (0.418) (0.450) s t a n d a r d d a y , u n i n s t a l l e d (3,937) (3,937) (Bellmouth and Coannular Nozzle) Sea l e v e l . s t a t i c , s t a n d a r d 0 17,312 17,312 0.0431 0.0437 +1.4 day, i n s t a l l e d ( n a c e l l e l i p (0.423) (0.429) (3.892) (3,892) I and mixer cumpound n o z z l e ) Design c r u i s e 0 5,016 4,016 0.0759 0.0756 -0.4 ( e x t r a p o l a t e d from static (903) (903) (0.744) (0.741) d a t a ) , nach 0.8, 12.192m1

(40.030 f t ) . i n s t a l l e d

( n a c e l l e l i p and mixer compound n o z z l e ) TABLE 9. EMISSIONS TEST RESULTS VERSUS PROGRAM GOALS.

EPAPS, kg/4448 N-h/cycle (lbm/1000 l b f - h r / c y c l e ) P o l l u t a n t Program Goal T e s t R e s u l t Unburned Hydrocarbon (HC) 0.73 ( 1 . 6 ) 0.73 (1.6) Carbon Monoxide (C) ) 4.26 ( 9 . 4 ) 3.63 (8.0) O x i d e s o f N i t r o g e n (NOx) 1.68 ( 3 . 7 ) 2.09 (4.6) Smoke

38 * 4 2 *

*EPA Smoke number.

AIRESEARCH QCGAT ENGINE - ACOUSTIC TEST RESULTS h r r y S. Kisner AiResearch Manufacturing Company of Arizona A Division of The Garrett Corporation The n o i s e l e v e l s o f t h e AiResearch Q u i e t , C l e a n , G e n e r a l Avi- a t i o n Turbofan (QCGAT) e n g i n e were measured i n ground s t a t i c n o i s e tests. The static n o i s e l e v e l s were found to be markedly lover t h a n t h e d e m o n s t r a b l y q u i e t AiResearch Model TFE731 e n g i n e . The measured QCGAT n o i s e l e v e l s were c o r r e l a t e d w i t h a n a l y t i c a l n o i s e - s o u r c e p r e d i c t i o n s to d e r i v e f r e e - f i e l d component n o i s e p r e d i c - t i o n s . These component n o i s e s o u r c e s were used t o p r e d i c t t h e QCGAT f l y o v e r n o i s e l e v e l s a t FAR P a r t 36 c o n d i t i o n s . The p r e - d i c t e d f l y o v e r n o i s e l e v e l s are a b o u t 1 0 d e c i b e l s lower t h a n t h e c u r r e n t q u i e t e s t b u s i n e s s jets.

INTRODUCTION T h i s paper d e s c r i b e s t h e a c o u s t i c d e s i g n , s t a t i c n o i s e test r e s u l t s , n o i s e s o u r c e c o r r e l a t i o n a n a l y s e s , and f l y o v e r n o i s e p r e - d i c t i o n s for t h e AiResearch QCGAT engine.

NOISE GOALS NASA s p e c i f i e d g o a l s f o r t h e QCGAT e n g i n e a t t h e FAR P a r t 36 s i d e l i n e , t a k e o f f , and approach c o n d i t i o n s a s a f u n c t i o n o f maximum t a k c o f f g r o s s weight. The n o i s e g o a l s f o r t h e twin-engine a i r p l a n e p o s t u l a t e d i n t h i s program a r e shown i n f i g u r e 1. The maximum t a k e o f f g r o s s w e i g h t f o r t h e a i r p l a n e d e f i n e d by AiResearch is 8674 kg (19,122 l b ) . The s p e c i f i c n o i s e g o a l s a t t h e FAR P a r t 36 c o n d i t i o n s a r e : Takeoff (without cutback) : 73.3 EPNdB Sideline (1500 ft): 82.3 EPNdB Approach : 87 . 3 EPNdB These levels are significantly below the existing FAR Part 36 Stage 3 noise limits.

NOISE OBJETIVES To achieve the program noise objectives, large turbofan engine noise-reduction technology was applied to the smaller AiResearch general aviation engine. The objectives accomplished during the program were as follows: o The engine was defined, and the cycle cnnditions were determined to provide low noise-generation features.

o An acoustically treated nacelle was designed and fabr i- cated.

o The ground static engine noise levels were measured for several con£ ig~rations, establishing an engine baseline and demonstrating the effectiveness of the acoustical design features.

o Static noise-source correlations were developed, and component noise spectra with adjustments for flight effects were used to estimate flyover noise levels in compliance with QCGAT noise goals.

ACOUSTIC DESIGN FEATURES The acoustic design effort emphasized minimizing noise genera- tion at the source and maximizing noise reduction achieved through judicious application of nactlle acoustic treatment in the fan inlet and exhaust ducts. Acoustic: duct liner configurations were designed to balance the noise suppression at takeoff, sideline, and approach condition, providinq the broadest possible attenuation bandwidth without sacrificing sianificant attenuation from optimum at any one of the three operating conditions.

The major acoustic features of the QCGAT erlgine are illus- trated in figure 2. Noise-reduction technology was applied to the two major noise sources, the fan and the jet. The fan noise-source reduction features included the following: elimination of inlet guide vanes, low tip speed and pressure ratio, single-stage fan, a large rotor to stator spacing of 2.12 rotor chords, and c . large number of bypass and core stators to cut-off rotor-stator intcrac- tion tones. The jet noise-reduction features included low fan dis- charge and primary jet-exhaust velocities, and a 12-lobe mixer compound exhaust nozzle .

The nacelle acoustic treatment design selected for the QCGAT engine consisted of a single-cavity system used in series with different cavity depths in the axial direction and where possible equivalent depths on opposing walls. A broadband resonator was constructed from aluminu~ perforated sheet bonded to a 0.95-cm (3/8-in.) all-aluminum honeycomb backing because of its structural ruggedness, low cost, and known acoustic performance.

A schematic of the acoustic liner inst~L!ation is shown in figure 3. The inlet-wall treatments--sections A1, A2, and B--were tuned to provide primary suppression at the FAR Part 36 sideline condition with a length equal to 2.16-cm (0.85-in.) mea;! inlet diameter. The fan discharge duct treatments--sections 1, 2, ari5 3--were tuned to provide balanced attenuation between sideline and approach conditions, and have an t . f fective total length equal to 5.4 times the average duct height.

F i n a l o p t i m i z a t i o n or t h e e n g i n e a n d n a c e l l e e x h a u s t l i n e r d e s i g n w a s c o m p l e t e d u s i n g a c o m p u t e r p r o g r a m b a s e d upon t h e a x i - s y m m e t r i c mode t h e o r y of M i n n e r a n d R i c e ( r e . l T o a c h i e v e optimum a t t e n u a t i o n , t h e r e q u i r e d c a v i t y d e p t h s a n d f a c e s h e e t o p e n a r e a s w e r e c o m p u t e d .

The d e s i g n p r o c e d u r e f o r t h e i n l e t l i n e r s was b a s e d u p o n t h e r e c e n t m u l t j m o d a l d u c t t r e a t m e n t a n a l y s i s d e v e l o p e d a t NASA-Lewis by R i c e ( r e f . 2 t h r o u g h 5 ) . T h e i n l e t l i n e r s were t u n e d t o a t t e n u - a t e modes t h a t r a d i a t e e n e r a y a t l a r g e r a n g l e s f r o m t h e i n l e t a x i s , t h u ~ r e d u c i n g s i d e l i n e n c i s e r a d i a t i o n , The major d e s i g n character is ti.^^ o f t h e iclet a n d oxhaust l i n e r s a r e shown i c t,?bie 1. T h c i n l e t s e c t i o n s A A 2 , a n d B h a v e 1 ' b a c k i n g d e p t h s o r 1 . 8 3 cm ( 0 . 7 2 i n . ! , 2.8 c m (1.1 i n . ) , a n d 1 . 3 5 c m ( 0 . 5 3 i n , r e s p e c t i v e l y . Open a r e a s r a n g e f r o m 5 . 8 to 1 4 . 2 per- c e n t . The i n l e t l i n e r s a r e t u n e d for t h e s i d e l i n e c o n d i t i o n i n t h e 1 0 0 0 - to 2500-Hz r a n g e . The t o t a l l e n g t h o f t h e i n l e t t r e a t m e n t is 59.9 cin ( 2 3 . 6 i n . ) . The e x h a u s t l i n e r c o n f i g u r a t i o n is 123.7-cm ( 4 8 . 7 - i n . ) l o n g a n d was t u n e d f o r a p p r o a c h c o n d i t i o n s +ere f a n e x h a u s t n o i s e is d o m i n a n t i n t h e 2000- t o 4003-Hz f r e q a e n c y r; 'e.

ENGINE NGISE TESTS The ZCGAT e c g i n e was i n s t a l l e d a t t h e A i R e s e a r c h S a n Tan test f a c i l i t y { f i g . 3 ) f c r a c o u s t i c a l , n e a s u r e i n e n t s . Noise d a t a was t a k e n a t s p e c i f i e d e n q i n e l o a d c o n d i t i c n s f r o m g r o u n d i d l e t o t a k e - o f f power to d e t e c m i n e t h e u n t r e a t e d e n g i n e n o i s e l e v e i s , t h e n o i s e r e d u c t i o n a t t a i n e d w i t h v a r i o u s c o m b i n a t i o n s o f a c o u s e i c t r e a t - m e n t s , a n d t h e n o i s e r e d u c t i o n a c h i e v e d w i t h a m i x e r c o m ~ o u n d e x h a u s t n o z z l e . T h i s d?;d was u s e d t o d e t e r m i n e t h e s t a t i c n o i s e l e v e l s f o r use i n p r e d i c t i n g f l y o v e r n o i s e l e v e l s .

A s c h e m a t i c o f t h e a c o u s t i c test s e t u p a t San Tan is shown i n f i g u r e 5. Data was t a k e n o n a 30.4-meter ( 1 0 0 - f o o t ) r a d i u s a t e v e r y 1 0 d e g r e e s , f r o m 1 0 d e g r e e s t o 1 6 0 d e g r e e s , f o r e a c h c o n f i g - u r a t i o n a n d l o a d c o n d i t i o n . The m i c r o p h o n e s a r e B&K, 1.270-cm (0.5-in.) diameter, Type 4133, mounted f o r n o r m a l i n c i d e n c e o f t h e d i r e c t s o u n d f i e l d a n d were l o c a t e d 1 . 5 meters ( 5 f t ) above t h e g r o u n d .

I n a d d i t i o n io t h e 1 6 f a r - f i e l d m i c r o p h o n e l o c a t i o n s , 6 i n t e r - n a l n o i s e m e a s u r e m e n t s w e r e made w i t h t h r e e 0.3175-cm (0.125-in.)

c o n d e n s o r m i c r o p h o n e s a n d t h r e e 0,6350-cm (0.25-in.) c o n d e n s o r i n f i n i t e t u b e s y s t e m s ( f i g . 6). Two 0.3175-cm ( 0 . 1 2 5 - i n . ) micro- p h o n e s w e r e i n s t a l l e d f l u s h mounted w i t h t h e d u c t s u r f a c e i n t h e f a n i n l e t n a c e l l e , o n e n e a r t h e f a n t i p , t h e o t h e r n e a r t h e n a c e l l e i n l e t . A n o t h e r 0.32-cm (0.125-in.) m i c r o p h o c e was l o c a t e d i n t h e e x h a u s t d u c t n e a r t h e m i x e r e x i t p l a n e . The 0.6350-cm ( 0 . 2 5 - i n , ) i n f i n i t e t u b e s y s t e m s were located i n t h e l o w - p r e s s u r e (LP) t u r b i n e r e a r - b e a r i n g s u p p a r t a r e a a f t <if t h e LP t u r k i n e , n e a r t h e m i x e r e x i c p l a n e , a n d n e a r t h e e x h a u s t n o z z l e e x i t p l a n e . The i n t e r n a l n o i s e m e a s u r e m e n t s w e r e r e c o r d e d s i m u l t a n e o u s l y w i t h t h e f a r - f i e l d nleasilrements. T h i s d a t a was r e c o r d e d f o r 2 m i n u t e s a t e a c h con- d i t i o n to allow c o h e r e n c e a n a l y s i s betw?en t h e i n t e r n a l a n d f a r - f i e l d n o i s e .

A l l tests w e r e c o n d n c t e d w i t h i n t h e r e c o r n e n d e d e n v i r o n m e n t a l l i m i t s o f wind s p e e d , t e m p e r a t u r e , a n d r e l a t i v e h u m i d i t y . The tests w e r e c o n d u c t e d i n November 1 9 7 8 f r o m m i d n i g h t to 6 a.m. when t h e wind was c a l m an-' a m b i e n t n o i s e l e v e l s w e r e l o w . The t e m p e r a - t u r e r a n g e d from 280K (44OF) to 286K (56OF) a n d t h e r e i a t i v e humid- i t y r a n g e d f r o m 70 t o 85 p e r c z n t d u r i n g t h e t e s t s .

The key a c o u s t i c p a r a m e t e r s f o r t h e s i m u l a t e d s t z t i c t a k e ~ f f and a p p r o a c h c o n d i t i o n s a r e shown i n t a b l e 2. A t t a k e o f f , t h e e n g i n e o p e r a t e s a t 1 6 , 0 9 8 N (3619 l b ) o f t h r u s t w i t h a f a n p r e s s u r e r a t i o o f 1.41. The f a n r e l a t i v e t i p Mach number is s u p e r s o n i c a t 1 . 1 7 , a n d t h e m i x e r e x h a u s t v e l o c i t y is o n l y 258 m / s (846 f t / s e c ) .

The f a n - b l a d e p a s s i n g f r e q u e n c y is 5236 Hz, i n a l o w a n n o y a n c e r a n g e .

A t a p p r o a c h , t h e f a n o p e r a t e s s u b s o n i c a l l y a t a r e l a t i v e t i p Mach number o f C.79. The f a n p r e s s u r e r a t i o is 1-10, a n d t h e fan- b l a d e p a s s i n g f r e q u e n c y is 3638 Hz. T t e t h r u s t l e v e l a t a p p r o a c h , s t a t i c c o n d i t i o n , is 7019 N (1578 Ib) w i t h a l o w e r m i x e r e x h a u s t v e l o c i t y o f 1 6 6 m / s ( 5 4 5 f t / s e c ) .

T a b l e 3 shows t h e same key a c o u s t i c p a r a m e t e r s o f t h e FAR P a r t 36 f l i g h t c o n d i t i o n s o f t a k e o f f , s i d e l i n e , a n d a p p r o a c h . T a b l e s 2 a n d 3 show a c o m p a r i s o n b e t w e e n static a n d f l i g h t f a n r e l a t i v e t i p Mach numbers a n d jet v e l o c i t i e s . A t t a k e o f f , t h e QCGAT a i r p l a n e r e a c h e s a n a l t i t u d e o f 1 1 5 1 m (3776 f t ) a b o v e m e a s u r e m e n t l o c a t i o n .

A t t h i s a l t i t u d e , t h r u s t is a t 1 2 , 8 6 9 M ( 2 8 9 3 l b ) , w i t h a f a n pres- s u r e r a t i o o f 1.44. The f a n r e l a t i v e t i p Mach number is 1 . 2 2 , a n d b l a d e p a s s i n g f r e q u e n c y is 5495 H z . Mixer e x h a u s t v e l o c i t y is 285 m / s (936 f t / s e c ) , A t a p p r o a c h , t h e f a n r e l a t i v e t i p Mach number i r 0.78, t h e f a n p r e s s u r e r a t i o is 1.16, a n d t h e b l a d e p a s s i n g f r e q u e q c y is 3677 Hz. A t a t h r ~ i ~ t l e v e l o f 4639 N ( 1 0 4 3 l b ) , t h e m i x e r e x h a u s t v e l o c i t y is 1 7 6 m/s ( 5 7 7 f t / s e c ) . S i d e l i n e a c o u s t i c param- e t e r s a r e e s s e n t i a l l y t h e same a s t a k e o f f a c o u s t i c p a r a m e t e r s .

A c o u s t i c d a t a was t a k e n f o r t h e s e v e n t e s t c o n f i g u r a t i o n s l i s t e d i n t a ~ l e 4. The f u l l y t r e a t e d e n g i n e was t e s t e d f i r s t w i t h b o t h m i x e r compound a n d c o a n n u l a r e x h a u s t n o z z l e s y s t e m s ( c o n f i g - u r a t i o n s 1 and 2 ) . With t h e m i x e r n o z z l e i n s t a l l e d , a c o u s t i c p a n e l s w e r e s y s t e m a t i c a l l y r e p l a c e d w i t h h a r d w a l l p a n e l s i n c o n £ i g - u r a t i o n s 3 a n d 4 u n t i l t h e f u l l y h a r d w a l l c o n f i g u r a t i o n 5 was a t t a i n e d . C o n f i g u r a t i o n 6 was t h e h a r d w a l l e n g i n e w i t h t h e n a c e l l e l i p i n s t e a d o f t h e f l i g h t - s i m u l a t o r l i p . The f i n a l c o n f i g u r a t i o n , c o n £ i g u r a t i o n 7 , c o n s i s t e d o f t h e h a r d w a l l n a c e l l e , f l i g h t - s i m u l a t o r l i p , a n d c o a n n u l a r n o z z l e . C o m p a r i s o n s w e r e made b e t w e e n t h e t r e a t e d and h a r d w a l l w i t f i mixer compound n o z z l e (con- f i g u r a t i o n 1 v e r s u s s ) , t r e a t e d and h a r d w a l l w i t h c o a n n u l a r n o z z l e ( c o n f i g u r a t i c n 2 v e r s u s 7 ) , mixer compound v e r s u s c o a n n u l a r n o z z l e w i t h t r e a t e d n a c e l l e ( c o n f i g u r a t i o n 1 v e r s u s 2 ) , and mixer compound v e r s u s c o a r n u l a r n o z z l e w i t h h a r d w a l l n a c e l l e ( c o n f i g - u r a t i o n 5 v e r s u s 7 ) .

GROUND REFLECTION ANALYSIS Before t h e ground s t a t i c a c o u s t i c d a t a c a n be compared or used to p r e d i c t f l y o v e r n o i s e l e v e l s , t h e d a t a must be c o r r e c t e d f o r FAA 248K (77OF) and 7 0 - p e r c e n t r e l a t i v e humidity, znd f o r ground r e f l e c t i o n . The ground r e f l e c t i o n problem ,is i l l u s t r a t e d i n f i g - u r e 7. h wave r e f l e c t e d from t h e ground i n t e r f e r e s w i t h t h e d i r e c t sound wave a t t h e r e c e i v e r . Depending on ground . o u s t i c impediince, t h e r e f l e c t e d wave c a n d i m i n i s h or enhance t h e sound i n t e n s i t y a t t h e microphone due t o a p h a s e - a n g l e s h i f t . The t y p e o f soil a t t h e San Tan s i t e c o n s i s t s o f a random c o m b i n a t i o n o f hard- packed c l a y , s a n d , and decomposed g r a n i t e p a r t i c l e s ; no known d a t a e x i s t s on t h e impedance o f t h i s soil.

The t e r r a i n around San Tan C e l l N o . 5 s l o p e s downward from t h e e n g i n e pad so t h a t t h e ground l o c a t i o n s upon which t h e microphones were pole-mounted a r e a t an a v e r a g e e l e v a t i o n o f 1.13 m (3.7 f t ) below t h a t o f t h e e n g i n e pad. Thus, t h e QCGAT e n g i n e , which was mounted 2.29 m ( 7 . 5 - f t ) above t h e e n g i n e pad was, on t h e a v e r a g e , 3.41 m ( 1 1 . 2 - f t ) above t h e ground, r e l a t i v e to t h e microphone loca- t i o n s .

The impedance c o r r e l a t i o n p r o c e d u r e , based upon r e f e r e n c e s 6 t h r o u g h 1 3 , is o u t l i n e d a s f o l l o w s : 1. Measured d a t a a t t a k e o f f c o n d i t i o n a t a l l t h r e e micro- phone h e i g h t s and a l l 1 6 a r r a y a n g l e s were used t o o b t a i n f i n a l AiResearch San Tan s o i l impedance e s t i m a t e s .

2 . Using previously published data, an initial normalized impedance array was assumed (R/pc and X/pc versus fre- quency).

3 . The excess attenuation, A , , was computed for each micro- phone height, and cor rectcd sound pressure level ( S P L , ) spectra was determined.

4 . A 3-way difference scheme was used to calculate the dif- ferences between the three corrected spectra at each 1/3- octave band.

5 . Iterations were performed on the values of R/pc and X/pc until all differences approached zero (steps 3 and 4, above). The convergence criteria was based upon the values of average differences at each 1/3-octave band.

When reasonable values of impedance failed to provide convergence at a 1/3-octave band, the two microphone heights having a frequency furthest away from a null fre- quency were used and convergence was obtained.

6 . Inasmuch as convergence criteria was based on aver- age differences, observations of individual differences were then made, and minor adjustments to the normalized impedance were performed, thus establishing the final impedance values given in figure 8.

7, Excess attenuation 1/3-oc tave band spectra was computed for the three microphone heights, based on final ground impedance estimates.

8 . Ae spectra was then applied to the measured data for all three microphone heights. Comparison plots were prepared at representative array angles.

9. Acoustic measurements were also made at the three micro- phone heights for approach. To check the relative validity of the ground-reflection correction procedure, the Ae spectra was applied to the approach data and com- parisons of the corrected data again were made. The cor- relation of the approach corrected data was consistent with that of the takeoff corrected data.

An example of the 'as measured' spectra from each microphone is shown in figure 9. Large differences between the pole-mounted and ground microphones were observed from 200 to 4000 Hz. Fig- ure 10 shows the same data with the excess attenuation corrections applied. Overall, good agreement was obtained for all microphones and all 1/3-octave band frequencies.

An example of the final result of applying the ground correc- tion is shown in figure 11. The free-field levels were reduced in the low frequency range, and the grourid dip in the 400- through 500-Hz range was decreased, resulting in 2 7mooth spectral curve.

Little or no change occurred at the high 1 quencies.

ACOUSTIC COMPARISONS OF STATIC DATA The corrected data for each acoustic configuration tested was compared to establish trends and illustrate the level comparisons with the equivalent Model TFL731-3 takeoff and approach static data. A comparison between the hardwall coannular configuration-- the loudest QCGAT configuration--with the Model TFE731-3 at takeoff condition is shown in figure 12. The QCGAT tone-corrected per- ceived noise levels (PNLT) are considerably lower than the Model TFE731-3 primarily because of the lower exhaust velocity, even though the QCGAT engine produces 8-percent more thrust. This dif- ference is shown more vividly in the 1/3-octave spectral plot at 150 degrees from the inlet axis ( f i g . 13). This shows clearly a r e d u c t i o n i n jet n o i s e , a s w e l l a s i n t h e h i g h f r e q u e n c y f a n t o n e .

Similar r e d u c t i o n s were a c h i e v e d a t approach.

F u r t h e r r e d u c t i o n s i n n o i s e were a c h i e v e d w i t h t h e QCGAT mixer compound n o z z l e as shown i n f i g u r e 1 4 . A t t h e same 150-degree a n g l e , t h e QCGAT c o a n n u l a r and mixer compound n o z z l e c o n £ i g u r a t i o n s are compared a t t a k e o f f s t a t i c c o n d i t i o n . A t 200 Hz, t h e mixer is a b o u t 7 d B q u i e t e r t h a n t h e c o a n n u l a r n o z z l e . Note, however, t h a t t h e r e are p e a k s a t 1600 and 2500 Hz w i t h t h e mixer b e i n g 2- to 3-dB h i g h e r a t t h e s e f r e q u e n c i e s . The s o u r c e o f t h e s e t o n e s were i n v e s - t i g a t e d i n d e t a i l , i n c l u d i n g some c r o s s - c o r r e l a t i o n a n a l y s i s a t NASA. The r e s u l t s r e v e a l e d a h i g h c o r r e l a t i o n between i n t e r n a l c o r e n o i s e and t h e f a r - f i e l d n o i s e l e v e l s a t c e r t a i n d i s c r e t e f r e - q u e n c i e s , p r i m a r i l y c e n t e r e d a b o u t 200 Hz and 2500 Hz. T h i s l e d to t h e development o f a new n o i s e - s o u r c e c o r r e l a t i o n a t t r i b u t i n g t h i s e x c e s s n o i s e to c o r e n o i s e .

F i n a l r e d u c t i o n s i n n o i s e were a t t a i n e d w i t h t h e a c o u s t i c a l l y t r e a t e d mixer c o n f i g u r a t i o n a s shown i n f i g u r e 1 5 . A l s o shown i n f i g u r e 1 5 is t h e t r e a t e d v e r s u s h a r d w a l l mixer d a t a a t approach c o n d i t i o n and a t 50 d e g r e e s from t h e f a n i n l e t . A broad r a n g e of f r e q u e n c i e s from 630 Hz to 6300 Hz a r e a t t e n u a t e d due to t h e i n l e t t r e a t m e n t . The same c o n f i g u r a t i o n s a r e compared a t 1 2 0 d e g r e e s i n t h e a f t q u a d r a n t i n f i g u r e 1 6 . Here, l a r g e r a t t e n u a t i o n s approach- i n g 10 dB a r e o b s e r v e d , b u t i n a narrower f r e q u e n c y band. T h e a c t u a l a t t e n u a t i o n s i n t h e lower f r e q u e n c i e s c a n n o t be o b s e r v e d b e c a u s e o f t h e masking by jet and core n o i s e s o u r c e s .

The t r e a t m e n t was e f f ~ - t i v e i n r e d u c i n g t h e s i d e l i n e n o i s e l e v e l s a s shown i n f i g u r e 17. The t r e a t e d v e r s u s h a r d w a l l coan- n u l a r c o n f i g u r a t i o n s a t 90 d e g r e e s and a t approach c o n d i t i o n a r e shown.

The b l a d e p a s s i n g t o n e a t 4000 Hz is a t t e n u a t e d n e a r l y 7 dB.

I n s u m m a r i z i n g g r o u n d s t a t i c d a t a c o m p a r i s o n s , t h e A i R e s e a r c h QCGAT e n g i n e d e m o n s t r a t e d s i g n i f i c a n t l y q u i e t e r n o i s e l e v e l s t h a n t h e c u r r e n t l y q u i e t Model TFE731 b u s i n e s s - jet e n g i n e , a n d showed t h a t a p p l i c a t i o n o f n o i s e r e d u c t i o n t e c h n o l o g y , s u c h as a m i x e r compound e x h a u s t s y s t e m a n d a c o u s t i c a l l y t r e a t e d f a n i n l e t a n d e x h a u s t n a c e l l e s , a c h i e v e d e v e n lower n o i s e l e v e l s .

NOISE SOURCE CORRELATIONS A p r i m a r y o b j e c t i v e o f t h e QCGP.T a c o u s t i c p i o g r a m was t o d e t e r m i n e f l y o v e r n o i s e l e v e l s b a s e d o n static e n g i n e data, a n d to d e m o n s t r a t e t h a t t h e s e n o i s e l e v e l s meet t h e p r o g r a m g o a l s , which are set w e l l below c u r r e n t t e c h n o l o g y a i r p l a n e . To a c c o m p l i s h t h i s o b j e c t i v e , a m e t h o d o l o g y was d e r i v e d t o p r e d i c t t h e m a j o r component n o i s e s o u r c e s , a d j u s t t h e i n d i v i d u a l s o u r c e s f r o m s t a t i c t o f l i g h t c o n d i t i o n s , a n d p r e 6 i c t t h e n o i s e s o u r c e f l y o v e r l e v e l s .

The a n a l y t i c a l tools u s e d by A i R e s e a r c h to p r e d i c t QCGAT e n g i n e n o i s e s o u r c e s are p r e s e n t e d i n t a b l e 5.. The p r e d i c t i o n g.0- c e d u r e s f o r f a n n o i s e , jet n o i s e , and core n o i s e were b a s e d upon t h e NASA A i r c r a f t Noise P r e d i c t i o n Program (ANOPP) recommended p r o - c e d u r e s w i t h e m p i r i c a l m o d i f i c a t i o n s based upon p r e v i o u s A i R e s e a r c h e x p e r i e n c e .

A c o m p a r i s o n o f p r e d i c t e d n o i s e s o u r c e s b a s e d upon t h e s e p r e - d i c t i o n p r o c e d u r e s and m e a s u r e d d a t a is shown i n f i g u r e 1 8 . The f a n n o i s e p r e d i c t i o n a g r e e s w e l l w i t h t h e m e a s u r e d d a t a w i t h a s l i g h t o v e r p r e d i c t i o n o f t h e b l a d e p a s s i n g h a r m o n i c . However, t h e measured l o w - f r e q u e n c y n o i s e , p a r t i c u l a r l y f r o m 1 6 0 Hz t o 2500 Hz, is h i g h e r t h a n p r e d i c t e d j e t and c o r e n o i s e . I n o r d e r LO a c c o u n t f o r t h i s , it is n e c e s s a r y t o make a s s u m p t i o n s f o r t h e a p p o r t i o n m e n t o f t h e jet a n d c o r e t o t h e t o t a l n o i s e s i g n a t u r e . Two a p p r o a c h e s w e r e u s e d a n d a r e show? i n t a b 1 2 6: The f i r s t model a t t r i b u t e d t h e d i f f e r e n c e b e t w e e n p r e d i c t e d and m e a s u r e d n o i s e i n t h e 50- t o 2500-Hz f r e q u e n c y r a n g e t o t h e jet. ZL 1.oise was a d j u s t e d a c c o r d - i n g l y on an a v e r a g e d e l t a b a s i s . ' - second model assumed jet n o i s e p r e d i c t i o n s were v a l i d , and a d j u s t e d t h e core n o i s e t o e x a c t l y match t h e measured d a t a . Both models a d j u s t ? ; t h e f a n and t u r b i n e n o i s e i o e x a c t l y match t h e measured l e v e l s i n t h e a p p r o p r i a t e f r e q u e n c y range.

F i g u r e 1 9 shows a n example o f t h e jet-noise-dominated c o r r e l a - t i o n f o r t h e s o f t w a l l mixer a t 12C d e g r e e s and a t t a k e o f f c o n d i - t i o n . The a v e r a g e d i f f e r e n c e from 50 t o 2000 Hz is a p p l i e d a t e a c h f r e q u e n c y to produce a m o d i f i e d jet n o i s e p r e d i c t i o n t h a t f a i r s t h r o u g h t h e d a t a . I n t h i s model, core n o i s e is p r e d i c t e d t o be well below t h e j e t n o i s e a t t h e t a k e o f f c o n d i t i o n . Above 2000 Hz, t h e f a n n o i s e is a d j u s t e d to f i t t h e d a t a . T u r b i n e n o i s e c o n t r i b u t i o n s were u n i m p o r t a n t e x c e p t a t f r e q u e n c i e s above 12,500 H z , which was o u t o f t h e r a n g e o f i n t e r e s t f o r flyok ,r n o i s e c a l c u l a t i o n s .

S i m i l a r c o r r e l a t i o n s were m-de f o r e a c h f a r - f i e l d a n g l e from 1 0 to 160 d e g r e e s .

Tne same s e t o f a c o u s t i c d a t a is shew3 i n f i g u r e 20, w i t h the core-noise-dominated model p r e d i c t i o n s . The je2 n o i s e p r e d i c t i o n is c o n s i d e r a b l y below t h e measured d a t a . The f a n n o i s e was d e t e r - mined t o b e t h e d i f f e r e n c e between t h e measured t o t a l and t h e p r e - d i c t e d sum o f jet, c o r e , and t u r b i n e r .se i n t h e 3150- to 10,000-Hz bands. The t o t a l o f a l l t h e n o i s e s o u r c e s e x a c t l y matched t h e measured d a t a .

FLYOVER PREDICTION PROCEDURE C a l c u l a t e d f l y o v e r n o i s e l e v e l s f o r t h e QCGAT e n g i n e were based upon t h e a d j u s t e d n o i s e s o u r c e s o b t a i n e d from c o r r e l a t i n g t h e p r e d i c t e d and measured s t a t i c n o i s e d a t a . A block diagram o f t h i s p r o c e d u r e is g i v e n i n f i g u r e 21. The measured c o r r e c t e d static n o i s e d a t a and t h e p r e d i c t e d n o i s e s o u r c e s a r e f e d i n t o a program called NASADELTA. The program compares and CO4KiteS difference spectra for each noise source at each engine operating condition.

The noise source prediction program is again used to preaict the noise levels at the FAR Part 36 flyover conditions. These predic- tions are adjusted by applying the appropriate correction spectra determined from the static data.

The adjusted m ' a sources are taken to flight conditions with corrections for distance, atmospheric attenuation, jet relative velocity and dynamic amplification effects, fan inlet cleanup, dop- pler effects, wing shielding, and ground effects. The adjusted sources are 'flownm along a prescribed flight path using the GTENFLY pr og ram.

For each f lyover condition -- takeoff , sideline, and approach--

the SPL, PNL, and PNLV rere calculated for each 1/2 second of the flight trajectory. The duration time, duration correction, effeztive perceived noise level (EPNI.) for each source, and the tc Cal EPNL were calculated in accordance to FAR Part 36 procedures.

FLYOVER NOISE CALIBRATION WITH MEASURED LEARJET DATA Based upon static data comparisons, the QCGAT engine demon- strated substantial reductions in noise levels compared to the quiet AiResearcn Model TFE731-2 engine thac powers the Learjet 35/36 airplane. Tnis airplane is certified to be 5 =dB below the FAR Part 36 Stage 3 noise limits. However, the initial flyover predictions, based upon the previously describedmethodology, yielded QCGAT noise levels comparable to measured Learjet flyover noise l%vels. This methodology was thus used to predict the Learjet fly- cvcr noise levels to determ~ne its validity.

A corn~ariso~r of the predicted and measured in-flight spectra f ~ r the Learjet 35/36 3ased upon the excess jet noise model, is shown in figure 22, Although the etatic noise predictions for the TFE731 engine were adjusted to match measured static noise levels, when taken to flight, the predictions are higher than the measured flyover levels, The overprediction occurs srimarily in the low frequency, jet-dominated range. At the takeoff condition, the pre- dicted f lyover EPNL is 88.7 EPNdB compared to a measured value of 84 EPNdB. Similar differences between predicted and measured fly- over noise l e s z ' . s were observed at approach and sideline condi- t ions.

A second set of flyover predictions for the Learjet cer~ifica- tion tests were made based upon the core noise dominated nodel, as shown in figure 23. The predicted levels are even higher than those based upon the previous model, This is primarily due to the assumed dominance of core noise to which beneficial in-flight reductions are not applied.

Table 7 compares the predicted and measured total EPNL for takeoff and a~proach conditions for both predict ion mode 1s. The ind iviiual noisa sources cannot be compared directly because the flight-path position for which the maximum tone-corrected perceived noise level occurs is not the same, resulting in a different com- position of noise sources. This shift in location of the maximum PNLT prevents the use of an in-flight spectral difference array to match the measured f lyover data. The spectral correction model was abandcned in favor of a single EPNL correction delta applied to each source. An outline of this calibration procedure used to match the measured Learjet data and to predict the $GAT flyover noise levels is shown in figure 24. The final QCGAT f lyover noise levels reflect measured flyover data and are thus considered real- istic.

QCGA'I' FLYO'JER NOISE PREDICTIONS The u n a d j u s t e d and a d j u s t e d f l y o v e r n o i s e p r e d i c t i o n s f o r t h e QCGAT e n g i n e are g i v e n i n table 8. Each method, w i t h a p p r o p r i a t e a d j u s t m e n t s f o r d i f f e r e n c e s between p r e d i c t e d and measured L e a r j e t l e v e l s y i e l d e d s i m i l a r resalts, i n d i c a t i n g t h a t t h e QCGAT e n g i n e is 2.0 EPNdB below t h e s i d e l i n e n o i s e g o a l , 4.6 to 5.4 EPNdB below t h e a p p r o a c h n o i s e g o a l , and from 0.2 EPNdB below to 1.4 EPNdB above t h e t a k e o f f n o i s e g o a l .

ACOUSTIC A N A L Y S I S SUMMARY A sullnary o f t h e a c o u s t i c a n a l y s i s performed i n t h e QCGAT pro- gram is o u t l i n e d below: o A q r o u n d r e f l e c t i o r i a n a l y s i s ~ a ~ d e v e l ~ p e d t o c o r r e c t t h e measured ~ t a t i c n o i s e l e v e l s to f r e e - f i e l d .

o F l y o v e r nL ise p r e d i c t i o n s were made based upon t w o s e p a - rate n o i s e s o u r c e c o r r e l a t i o n models: One assumed jet n o i s e to be t h e dominant g e n e r a t i n g mechanism; t h e o t h e r assumed c o r e n o i s e to be r e s p o n s i b l e f o r e x c e s s n o i s e above t h e known jet n o i s e l e v e l s .

o Both p r e d i c t i o n models were a p p l i e d to t h e TFE731-2- powered L e a r 2 e t and were found to o v e r ~ r e d i c t t h e mea- s u r e d i n - f l i g h t l e v e l s , a l t h o u g h t h e ground s t a t i c d a t a was used t o c a l i b r a t e t h e p r e d i c t i o n s .

o The o v e r p r e d i c t i o n s o c c u r r e d p r i m a r i l y i n t h e l o w - f r e q u e n c y range where b o t h j e t and c o r e n o i s e a r e e x p e c t e d to be i m p o r t a n t .

o Final flyover predictions were made with adjustments far the differences between predicted and measured bearjet noise levels.

SUMMARY OF ACOUSTIC RESULTS The noise reduction technology demonstrated in the QCGAT Pro- gram is summarized below: o The QCGAT so£ twa 11 nacelle/mixer configuration demon- strated a 9.3 EPNdB reduction in flyover noise at takeoff condition, a 10.3 EPNdB reduction at approach, and a 7 . 7 EPNdB reduction at siasl-ne condition compared to the TFE731-2-powered Learjet.

o The QCGAT 'nardwall nacelle coannulat nozzle configura- tion was shown to be 4 . 2 EPNdB quieter than the Learjet at takeoff condition, although the QCGAT airplane takeoff gross weight is 963 kg (2122 lb) greater than the Lear jet.

o The -AT hardwall nacelle/mixer was 3 . 5 EPNdB quieter at takeoff and 4.3 EPNdB quieter at approach than the QCGAT hardwall nacelle/coannular nozzle, o The QCGAT softwall nacelle/mixer was quieter than the QCGAT hardwall nacelle/mixer by 2 . 6 EPNdB and 1.3 EPNdB at approach and takeoff conditions, respectively.

The final QCGAT flyover noise levels based upon the excess core noise model are shown in figures 25 through 27, compared with the FAR Part 36 noise limits, the QCGAT 1,oise goals, and the mea- sured Leacjet flyover levels.

CONC WSIONS The measured static noise levels of the AiResearch QCSAT engine were markedly lower than the demonstrably quiet TPE731 eng j ae. The following conclusions were made: o Based on the excess jet noise correlation model, the QCGAT engine met or bettered the program noise goals.

o Based on t h e excess core noise correlation, t h e -AT

engine met o r bettered the program noise goals both for

hardwall and softwall nacelle configurations at sideline and approach conditions, and was slightly above t h e take- off noise goal.

The AiResearch QCGAT program has demnstrated that it is pos- sible to design quiet engines for general aviation aircraft, Minner, G.L. and Rice, E . J . , 'Ccmputer Method for Design of Acoustic Liners for Turbofan Engines," NASA TM X-3317, 1976.

Rice, E . J . , "Acoustic Liner Optimum Impedance for Spinning Modes with Mode Cut-Off Ratio as the Design Criterion,' AIAA

-

Paper 76-516, Palo Alto, C A . , 1976.

Rice, E. J . , "Inlet Noise Suppressor Design Method Based Upon the Distribution of Acoustic Power with Mode Cut-Off Ratio,' Paper Presented at the Thirteenth Annual Meeting of the Soci- ety of Engineering Science, Inc., George Washington Uni- versity, 1976.

Rice, X,J., "Multimodal Far-Field Acoustic Radiation Pat- tern -AF. Approximate Equation,' AIAA Paper 77-1281, presented in Atlanta Georgia, October, 1977.

Rice, E . J . , "Attenuation of Sound in Ducts with Acoustic

Treatment - A Generalized Approximate Eqiiation," NASA TMX-

71830.

Pao, b . ~ - . , A.R. Wenzel, and P.B. Oncely, "Prediction of Ground Effects on Aircraft Noise," NASA-TP-1104, 1978.

Zorumski, W . E . , "Prediction of Aircraft Sideline Xoise Atte~u- ation," NASA-TM-78717, 1978.

Chessell, C . I . . "Propagation of Noise Along a Finite Impedance Boundary,' J. Acoust. SQC. America, Vol. 62, No. 4, October 1977, pp. 825-834.

Chessell, C. I . , "Meteorological and Grcand Ef fzcts on the Propagation of Aircraft Noise Close to the Earth's Surface," , : 1 . of Sound and Vib., Vol. 60, No. 2, 1978, pp. 251-266.

Embleton, T.F.W., J.E. Piercy, and N . Olson, "Outdoor Sound Propagation Over Ground of Finite Impedance," J. Acoust. Soc.

Am.. Vol. 59, No. 2, February 1976, pp. 267-277.

Daigle, G.A., Effects of Atmospheric Turbulence on the Inter- face of Sound Waves Above a Finite Impedance Boundary," J.

Acoust. Soc. Am., Vol. 65, No. 2 . January 1979, pp. 45-49.

Yutnam, T.W., "Review of Aircraft Noise Propagation," NASA-TMX-56033, September, 1975.

Oncely, P.B., "Propagation of Jrt Engine Ncise Near a Porovs Surf~ce," J. Sound Vib., Vol. 13, No. 1, 1970, pp. 27-35.

I TABLE 1- Q C G A T NACEUE ATTENUAmR DESIGN CHARACTERISTICS.

I I Face

I

Cavity Honeycomb Sheet Liner mseC Operating Liner Open Depth Cell Size, Length, Condition t

Section Area, cB (in.) + cm (in.) cm (in.) 112 Tuned for

5 . 8 18.83 (0.72) 1.905 (0.75) 17.5 (6.9) 1000 Sideline " 1 8 -6 2.79(1.10) 1.905(0.75) 27.2(10.71 1000 Sideline " 2 1.32(0.52) 0.953(0.375) 15.2(6.0) 1 14.2 B 2500 Sideline

32.0 (12.6) I 7.0

1 1-42(0-56) 2500 Approach 0.953(0.375)

2 0.9lt0.36) 0.953(0.375) 1000 Approach 3 8 . 5 ! 8 . 6

3 1.96(0.77) 1.905(0.75) . 2000 Approach 52.8 (20.8) 6 . 8 TABLE 2 . QCGAT ENGINE KEY ACOUSTIC PARAMETERS FOR SIMULATED STATIC TEST CONDITION.

Simulated Static Test Condition, 282K ( 4 8 . F ) Engine Parameter Takeoff Approach

-

I 7,019 N (1.578 lbf)

Engine net thrust f 16,098 N (3,619 lbf)

I

i 634.8 rad/s (6,063 rpm)

Fan rotor speed 1 913 rad/s (8,726 rpin)

1 Pan pressure ratio, tip I 1.41 I 1.18

I

) Fan tip relative Mach No. 1.17

I

Fan blade passing frequency 15,236 Hz

I

Pan airflow Core airflw

I Mixer exhaust velocity 1 257.9 m/s (846 ft/sec) 1 166.1 m/s (545 ftlsec)

Mixer exhaust total

temperature 406.8R (732.3OR) 1 365.91 (658.L0R)

i !

I LP turbine rotor s ~ z e d 1 1,941.5 rad/s (18,543 rpml 1.348.9 rad/s (12,883 rpm)

I

Turbine last stage relative tip Mach No. 0.472

I

Turbine last stage pressure ratio (total

to static) 1.61 - - - . - I 1.22

TABLE 3 . QCGAT ENGINE K E Y ACWSTIC PARAUBTERS POR FLYOVER NOISE CONDITION.

FAR PART 36 CIRTIIICATIOei IITIOr

~ a t e o f f I sideline Approach

wine nt thrust 12.869 1 (2,893 lbf 4,639 N (1,043 lbf) Fan rotor speed 958.9 rad/s (9,159 rpml Pan pressure ratio, t i p 1.44 Fan t i p relative Uach no.

Fan blade passing f rcqutmy Fan airflow Core a i r f l a Uixer exhaust u l o c i t y Mixer exhaust t o t a l telpcrature LP turbine rotor speed Turbine l a s t s t w e relative t i p llllch no.

Turbine l a s t s t q e pressure r a t i o ( t o t a l t o s t a t i c ) TABLE 4. ACOUSTIC TEST CONFIGURATIONS.

Acocstic con fig^ ration Number Description Fully treated engine with mixer compound nozzl~ Fully treated engine with coannular nozzle

1 3 6 4 Partially treated w ~ t h mixer compound nozzle

I

5 Hardwall engine with mixer compound nozzle

I

I Hardwall engine with nacelle lip, mixer nozzle I 6 Hardwall engine with coannular nozzle I L- I TABLE 5 . QCGAT ENGINE NOISE PREDICTION PROCEDURE.

najor Caponent Noise Prediction nethod Sources Predicted

Fan inlet noise -

Discrete, broadband NASA MX-71763.

FAA-RD-11-73, Buzz saw

Fan discharge noise -

Discrete, broadband NASA 'RIX-71763.

NASA 'RIX-73552 Jet noise NASA R(S-71627 Combustion noise Turbine noise A I M 75-449 Total noise Sum of individual component noise levels NOTE: One-third octave spectra fraa 50 to 16,000 Hertz directivity angles from 0.17 to 2 . 7 9 radians (10 to 160 degrees) from inlet centerline.

*Modified by AiResearch i TABLE 6 . NOISE PREDICTION METHODOLOGY COMPARISON.

Excess Jet nodel Excess Core Model Jet noise based on NASA 1 Jet noise based on NASA rethod method adjusted to fair through low-frequency data Core noise based on N A : A Core noise defined as difference met hod 1 between mrasured and predZcted I sum of jet, fan, turbine in 50-250 Hz frequency bands r c discrete, broadband, Fan inlet and fan dischrrge and buzz saw adjcsted to defined as differc.~ce between

I :31 data measured and pre*icted sum of

I

iet, :ore. turbine in 3150- i - 0 , 0 0 3 Hz bands

I

/ OE turbine noise mett 3 1 GL turbine noise method I

I TABLE 7. TFE731-2/LEAR 36 PLYOVER NOISE COMPARISON.

EPNL, EPNdB Pred icted Excess Jet Excess Core Measured W e 1 node 1 Takeoff 88.7 (+4.7) 90.8 (+6.8) 84.0 Approach 92.2 95.9 (+3.7) 98.6 (+6.4) TABLE 8, QCGAT FLYOVER NOISE SUMMARY.

EPNL, EPNdB Excess Jet Excess Core -AT Configuration Prediction Method Goa 1 Prediction Method Unadjusted With Lear.\ Unadjusted With Leari\ Hardwall mixer takeoff 79.3 74.6 83.1 76.0 77.8 Softwall mixer takeoff 73.3 73.1(-0.2)' 81.7 4 . + 1 4 Hardwall mixer approach 88.2 84.5 91.0 84.5 Softwall mixer approach 86.4 87.3 82.7 (-4.6). 88.5 8 1 9 - 5 4 )

nardwall mixer sideline 1 85.7 1 81.7

:::: 81.7 1 1 Softwall mixer sideline 84.3 80.3(-2.0). 80.3 (-2.0)*

*Indicates difference between goal and predicted EPNL.

I

SIDELINE W I S E POI- 823 EPNd0 .

%IDELIN€ NOISE LEVEL BASED ON HIGHEST LEVEL .

OCCURING AT THE THREE SlDELlNE PO1NTS.

F i g u r e 1. QCGAT A i r p l a n e Noise Goals.

FAN EXHAUST DUCT ACOUSTICAL

VANES TREATMENT (LIH = 5.4)

\ , LOW FAN JET VELOCITY F i g u r e 2 . QCGAT A c o u s t i c D e s i g n F e a t u r e s .

F i g u r e 5 . A c o u s t i c T e s t S e t u p .

@ DENOTES 0.3175-crn (1 18-IN.) FfllCPOPHONE

@ DENOTES 0.635-crn (114-IN.) SEMI-INFINITE TUB2

F i g u r e 6 . I n t e r n a l A c o u s t i c I n s t r u m e n t a t i o n .

SOURCE

R1

\ RECEIVER -\

---

hs

r

\ \ \ \ \ \ ' * \ \ \ \ \

m r *

F i g u r e 7 . Gr0ur.d R e f l . e c t i o n C o r r e c t i o n Model.

ONE-THIRD OCTAVE BAND CENTER FREQllENClES - HZ

Figure 8 . Correlated Acoustic l r n p e d ? ~ c e f o r D e s e r ? Soil a t San Tan ( A i R e s e a r ~ h ) .

10' 2 4 68102 2 4 6 8 1 0 3 2 4 6 8 1 0 1 2 ONE-THIRD OCTAVE BAND CENTER FREQUENCY, HZ Figure 9 . Measured Data for Acoustic Configuration No. 2 at 2.62 Radian (150-Degree) Position.

10' 2 4 6 8 1 0 2 2 4 6 8 1 0 3 2 4 6 8 1 d l 2 ONE-THIRD OCTAVE BAND CENTER FREQUENCY, HZ Figure 10. Corrected Data for Acoustic Configuration No. 2 at 2.62 Radian (150-Degree) Position.

ONE-THIRD OCTAVE BAND CENTER FREQUENCY. HZ

30.48 m (100 FT) RADIUS - RAW DATA

MIKE HEIGHT: 1.524 rn (5 FT)

-- --

CORRECTED DATA AT 2.62 rad (1500) Figure 11. Raw Data versus Ground Corrected Data.

DEGREES TONE CORRECTED PERCEIVED NOISE LEVEL, PNd% 0 TFE731-3 RUN 9 A QCGAT CF708C Static Comparison at Takeoff of Model TFE731-3 Figure 12.

and QCGAT Coannular Nozzle Configurations.

-TFE731-3 PRODUCTION ENGINE ---GCGAT COANNULAR NOLZLE WlTH HARDWALL NACELLE Figure 13. Noise Level Comparison at Takeoff of Model TFE731 and QCGAT Coannular Nozzle Configurations.

---- COANNULAR NOZZLE

WlTH TREATED NACELLE I 1 1 1 I I 1 1 10' 2 4 6 8 1 d 2 4 6 8 1 0 3 2 4 6 8 1 0 ~ 2 ONE-THIRD OCTAVE BAND CENTER FREQUENCY, HZ Figure 14. Noise Level Comparison at Takeoff of QCGAT Mixer Compound and Coannular Nozzle Configurations.

C.1'7 rad (50") FROM INLET Figure 15. Approach Noise Level Comparison o f Hardwall versus Treated Acoustic Panels at 0.87 rad ( S O a ) from Engine Inlet Centerline for the Mixer Compound Nozzle Configuration.

2.09 rad ( 1 2 0 ~ ) FROM INLET Figure 16. Approach Noise Level Comparison of Hardwall versus Treated Acoustic Panels at 2.09 rad (120°) from Engine Inlet Centerline for the Mixer Compound Nozzle Configuration.

IW 1 1.57 rad 190') FROM INLET 1

I I .I I

---- TREATED NACELLE

- HARDWALL NACELLE

S o ! I I I I I .

ONE-THIRD OCTAVE BAND CENTER FREQUENCY, HZ Fiqure 17. Takeoff Noise Level Comparison of Hardwall versus Treated Acoustic Panels at 1.57 rad ( 9 0 ° ) from Engine Inlet Center line for the Coannular Nozzle Configuration.

10' 2 4 6 8 1 0 2 2 4 6 8 1 0 3 2 4 6 8 1 0 4 2 ONE-THIRD OCTAVE BAND CENTER FREQUENCY, HZ Figure 18. Comparison of Measured and Predicted Noise Levels Before Correlation.

ONE-THIRD OCTAVE BAND CENTER f REQUENCY, HZ

Figure 19. Data Correlation, Excess Jet Model, Softwall Mixer, 2.09-Radian (120-Degree) Position.

ONE-

Figure 20. Data Correlation, Excess Core Model, Softwall Mixer, 2.09-Radian (120-Degree) Position.

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I PREDICTED E P N L 88.7 1

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ONE-THIRD OCTAVE BAND CENTER FREQUENCY IN HZ Figure 22. Measured Learjet 35/36 versus Predicted Excess Jet Model.

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MEASURED 84.0 EPWdB ONE-THIRD OCTAVE BAND CENTER FREQUENCY, HZ Figure 23. Measured Learjet 35/36 versus Predicted Excess Core Model.

SlATlC NOISE DATA PREDICTED T FE731-3

NOISE SOURCE LEVELS NOISE 30URCE LEVELS 1

PREDICT OCGAT FLYOVER LEVELS APPLY TFE73llLEARJET MEASURED LEARKT 35/36 CALIBRATION DELTAS FLYOVER LEVELS AT TO OCGAT ENGINE FAR PART 36 CONDITIONS A

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1969 F A R 3 6 REQUIREMENT 1 I I 1977 RULE

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QCGAT GOAL W AIRESEARCH I 70 1 - 1 1 . 4 4 . 5 22.7 45.4 9d.7 (X 1000 kg) 3 1 0 50 100 200 (X l~JOO LBM) TAKEOFF GROSS WEIGHT Figure 2 6 . S i d e l i n e Noise Summary.

1977 RULE 0 FALCON 10 I LEARJET 36 0 C!TATION o QCGAT GOAL m

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1.4 4.5 22.7 45.4 90.7 (X 1OOO k ~ ) - 10 50 100 200 tX 1OOO LBH) TAKEOFF GROSS WEIGHT F i g u r e 2 7 . Approach Noise Summary.

QCCAT AIRCR\FT/ENGINE DESIGN FOR REDUCED NOISE AND EMISSIONS Leonard I'Anson and Kenneth M. Terrill Avco Lycoming Division INTRODUCTION The multi-engine general aviation fleet size is expected to increase by 70 percent in the decade of the 1980's according to the General Aviation Manufacturers Association (GAMA). These general aviation aircraft typically use suburban airports that are unprotected by commercial buffer zones. Con- sequently, there is the potential for general aviation to create nore wide- spread adverse community reaction to noise and ?ollution than that experi- enced with conrmercial air carrier aircraft. Recognizing this, NASA let con- tracts to apply large engine quieting and emissions reduction technology to smaller engines and to develop new and more suitable technology where re- quired. These resulting "Quiet, Clean, General Aviation Turbofan" (QCGAT) contracts required delivery of a turbofan engine and nacelle demonstrator, as well as preliminary defintion of an appropriate general aviatim aircraft system, that could use the engine as propulsion.

This paper describes the resulting aircraft/nacelle/engine designs created under the Avco Lycoming contract. These designs reflect the technical expertise of the following subcontractors:

Aircraft Design - Beech Aircraft Corporation

Nacelle Mechanical Design - Avco Aerostructures

Nacelle Acoustic Treatment - Lockheed Aircraft Corporation

AIRCRAFT PRELIMINARY DESIGN To guide the aircraft system design, fivo primary objectives were estab- lished: 1. Practical, direct application of technology without significant scaling was very important, This required selection of aircraft and engine sizes which trould be appropriate for an appreciable eegment of genercl aviation.

2. The aircraft must also offer attractive range, fuel economy, and flight speed. A target of 2593 kilometers (1400 nautical miles) was established. This exceeds the range of most current small business aircraft. It also provides non-stop travel between opposite extremes of high density traffic areas.

3. A cruise Mach Number of 0.62 was chosen as an optimum compromise be- tweeil time and fuel economy. It provides 40 percent higher cruise speed chan a turboprop, with a 30 percent improvement ic fuel econ- omy over operation at 0.8 Mach Number.

4. A balanced field length of 762 meters (2500 feet) was desired be- cause it permits safe operation from 70 percent of all U.S. airports which are open to the public, including airports with sod runways.

5. Ecological characteristics of an aircraft system are likely to be- come primary competitive parameters for general viation in the 1980 decade. Therefore, they deserve close attention in design selecrims.

The initial step in aircraft preliminary design was the selection of ap- propriate size and design. The vast majority of general aviation aircraft operating from airiields located in suburban communities are in the size class below 5433 kg (12,000 lb) qross weight. In the lower extremity of the gross weight spectrum, small private aircraft in the range below 1814 kg (4,000 lb) are generally powered by single-piston engines. It is expected that marker constraints for very low-cost aircraft in this class will dictate continued usage of piston engines for the foreseeable future. It, therefore, follows that the greatest public ecological benefits can be realized by in-

troduction of a quiet, clean "ircraft system in the 1811 - 5433 kg (4,000 -

12,000 lb) gross weight class. Figure 1 shoes the projected market volume for various sizes of general aviation aircraft.

As with the passenger c i i r trend towards smaller and more sophisticated cars to perform the same function, it is expected that the decade of 1980's will see a similar general aviation trend towards reduced aircraft weight and smaller engine size for the same mission. Because noise, emissions, and fuel consumption reduce with engine size, subsequent iaprovement in ecological characteristics can be anticipated. Utilizing technologies such as turbofan propulsion, high aspect ratio super critical wing and lightweight composite structures, it is expected that a new class of small general aviation 3ir- craft will emerge in the eighties. A target of 30 percent weight reduction was considered achievable.

For aircraft size selection, our target was the largest segment of gen- eral aviation oircra : where cost of turbofan propulsio~ does not preclude its introduction.

Figure 1 presents a composite plot of aircraft gross weight versus both "The Number of New Aircraft co be Built" and "The Current Estimated Nominal Aircraft Cost". The number of aircraft is based on General Aviation Manufac- turers Association data. The expected trend toward lighter weight and higher cost for the same mission has not been reflected to ensure conservative en-

gine sizing. The range cf 3175 - 4536 kg (7,OOC to 10,000 lb) gross weight

appeared attractive, with 3629 kg (8,CSO lb) selected as our goal.

With the defined aircraft golls and Lycoming estimates for engine per- formance, Beech Aircraft Corporat i on conducted parametric studies co optimize the aircraft preliminary design.

The aircraft which evolved is depicted in figure 2 . It is a sleek, ad- vanced design, six-place aircraft with 3538. kg (7,800 lb) maximum gross eight. l c offers a 2776 kilometer (1500 nautical mile) range with cruise speed of 0 . 5 Mach Number and will take-off and land on the vast majority of general aviation airfields. Advanced features include broad application of composite materials and a ~upercritical wing design with winglets. Full-span fowler flaps have been introduced to improve landing capability. Engines are fuselr;o-mounted with inlets over the wing to provide shielding of fan noise by the wing surfaces.

The high bypass ratio QCGAT engine plays an important role in shaping the aircraft design. It offers a dramatic reduction in specific fuel consumption c q a r e d with current pure jets a n d lorto-moderate bypass ratio turbofans.

Figure 3 provides this comparison, reflecting a 22 percent improvement in fuel economy.

This loti~r fuel consumption may be used in either of two ways or in corn- binat ion: It can substantially reduce aircraft gross weight for the same range. The reduced weight provides compound interest on the fuel economy. It also requires lower thrust favoring reduction of noise and emissions.

If preferred, the lover fuel consumption can be translated into longer range for the original gross weight.

We chose to reduce gross weight and favor ecological characteristics.

Composite structures have been used extensively in the aircraft prelimin- ary design to further reduce gross weight. Areas selected by Beech for the applicatian of composite materials are shown in figure 4 . Kevlar graphite composites were used for aircraft weight estimates. Further -potential for weight savings exists in the engine nacelles. Conventional design vas used to reflect the lorrisk, lorcost test nacelle. Critical load carrying members such as the wing spar are conventional aluminum construction.

Approximately 40 percent of the structure is fiber epoxy or honeycmb- bonded structure. The use of composite structure in aircraft design provides a decreasing rate of ber;-.fit as the application of composites becomes more widespread in the design. Initial selection of applications is in noncritical aieas. As the stress in selected areas increases, the design safety fdctor also increases to compensete for uncertainties resulting from the youth of the composite application. Beech cautions that, while these composite appli- cations are technicaliy feasible, development beyond the scope normally undertaken by industry would be required to assure success.

Structura; design is ic accordance with Federal Aviation Regulation Part 23 airwcrthiness standards for normal category airplanes.

A 17 percent thickness-to-cl~or~i rat io supercri t ical wing shape was selec- ted because it offers a i~umber of advantages over the conventional 1 2 percent NACA shape. These advstltages are summarized in figure 5.

From thc cross sectional comparison shown here, it can 5e concluded that the supercritical wing provides lsrger volume far fuel storage for the samr ci1or.i vidth. The thickness increase has the supplemen- tary benefit of higher section modu7.as, permitting lighter construc- tion for cquivalcnt bending loads.

The two shapes Lave comparable drag characterist ics in the cruise mode. Increase in the NACA airfoil thickness in an attempt tt.

achieve similar v.jlume is impractical, because it results in a sig- nificant reduction in useful flight speed combined with an overall drag incrcast* at lower speeds.

Iterative design studies show a 25 percent increase in fuel capacity combined with s 3 percent decrease in aircraft gross weight. These savings arc for anl,-quivalent aspect ratio of 1 0 and a desigc wing loading of 2250 N!m- (37 l b ! s q ft) of wing area.

Priol test data havt- shown an appreciable increase in lift capa- bilitv as depicted in this comF;arison. This promises a more forgiv- ing sircraft for vat-istinns in angle of attack. enhancing safety.

For equivalent sophistication of flap systems, reduced landing speeds arc ac;~icvable resulting in shorter landing field iength capability.

The airfoil selcct~d by Beech i s simi lar to the NASA CA (W) - 1 airfoil, but is tailored sp~cificallp for the high-speed, high fue! vo!ume and the high-lif t req~irem~nts of tl~t- QCCAT conf iguration. The pressure distribution used to guide the design taii,*ring wolrld identifv i t ss a BAC Sonic Plateau airfoil wit11 a 17 prrctwt tiiickncss rstio.

Full-span fowlt-r flap:: nnd spoilers have hctn introduced to achieve the desireu 761 m i . t t . r (25OC) feet) take-oft field length and landing distances with reduced ~ i n c 31.C3. Wi~iglt-ts I I ~ V C nlso hctn addcd to reduce actual span and wing strirct~r31 wt*it~l1t, whilt- maintaining high effective aspect ratio.

Major i i f t pnrnmrters nrc summarizcJ hclow. Establ ishing optimum flap settings was bcvond t h i . scopr of this study. Howevcr, rsp-rience indicates that a full flap dt.flectio:l of 40 dc.gt-t-rs for landing and take-off flap setting of $0 percent of full deflection art* a p ~ -opriate for fowler flap de- sign. These valucs of C rcprcqent avni lahl? statr-of-the-art uith ad- lm3 x vanccd airfoils.

L Flap Posit ions

e - 0

Down 2.13 .088 3.45 Nany drag influencing design details of the QCGAT airplane a n aot ertab- lished at this time, because the airplane is as yet a preliminary deriga study. For drag analysis, ambitious estimates were u d e for the various items. Achievement of total airplane drag coef ficienta vill require s u c t i r y effort in the practi~al development of the airplane. The r e s u l t i w QCGAT air- craft drag compares *ith that of : h e tearjet W e 1 26, which is 8 1 3 e x t r m l y clean airplane. A?lwances have been made for differences in wing thicknear, cmponent sizes, etc. Drag coefficients used are 8-rized below: Flaps and gear up .02661, .02534 cruise Total CD

P '

Incre=nral CD for landing gear .0164 P

. N O 6 6

i-cremental CD for full flap P .0163 Incremental CD for T . O . flap P Incremental CD for one engine out .01209 P Four majar airplane variables were considered in the parametric study to optimize the wing configuration. They are: 1. Wing area 2 . Wing aspect ratio 3. Fuel 8-pipht 4. Take-a~f weight In the study, for each performance goal, the limiting aspect ratio verrus wing area is plotted for several take-off weights, including the eEfectr of wing geometry on wing weight. These limits for each of the performance go818 are then sirmmarized on a graph so that the best compromise can be aelccted. A design ?oint of 15.33 square meter (165 square feet) wins area and an effec- tive aspect ratio of 10 were selected.

Table 1 summarizes the expected weights for fuel, structure-plus-propul- sion. and complete aircraft with payload for both conventional and QCCAT rir- craft designs.

The first line represents a hypothetical aircraft of current vintage de- sign with low bvpass turbofan propulsion. Introduction of a QCCAT high byparr turbofan reduces fuei consumption by 22 percent. When this savings ia iter- ated through the aircraft design. structure and gross weight reduce, provid- ing an additionai 5 . 5 percent in fuel economy. Similar iterations with weight savings f r m composite materials and supercritical wing result in an addi- tional 4.4% savings in fuel. The combination sf engine and aircraft changes providc 32L better fuel etonumy. The 221 reduction in gross weight permits.

the use of a smaller engine with 22% lover thrust and, therefore, lower abso- lute emissians and noise.

% t aircraft study projected the m a x i m ranges shown in figure 6 for various payloads. While 11% kg (2530 lb) is depicted as maximum payload for the aircraft, only 7 5 3 kg (1660 ~ b ) is required to accommodate six people with their baggage. At this payload, the achievable range is in excess of 2963 kilometere ( 1 6 0 0 nautic miles). Flight coitditions are 10058 meters (33,000 feet) and an average flight speed of approximately 0.5 Mach Number.

In 9ti.r QCGAT aircraft study. landing distance, rather than take-off capa- bility, set the minimnu usable air£ ield length. Introduction of full-span fowler flaps vith moderate wing loading results in a ver j low "landing con- figuration" stall speed. The 32 metersjscc (62 knots) stall speed compares

vith 4 1 - $6 metersjsec (SO - 90 knots) for current typical jet and turbofan

aircraft. Since landing distance is proportional to stall speed squared, this low landing speed provides an attractive sea levei FAR landing field length of 811 meters (2603 feet).

Figure 7 shims a representative sample of general aviation airfields plotted on cocrdinates of field elevation and field length. The Beech QCGAT aircraft with full useful pavloads has a landing capability consistent with the majority of these fields.

The expected stall spezds promise a very forgiving airplane in the take- off and landing mode where most accidents occur.

The aircraft preliminary design was conducted to establish realistic cri- teria for noise measurement. Figure 8 depicts the iocations for noise aea- sulement, as well as the aircraft and engine conditions at the point of measurement.

Approach noise is measured directly below the flight path 1552 meters (one nautical mile) prior to the beginning of the runway. Approach aircraft glide slope is fixed at three degrees. Take-off sideline cc-sists ? £ multiple measurements 4 6 3 meters (0.25 nautical miies) to ,one side of the take-off flight path. Take-off flyover condition is measured 6 4 8 2 meters ( 3 . 5 nautical miles) from brake release, directly below the flight path.

Looking at the tabulation of aircraft and engine conditions, the approach conditions are quoted for 40-degree wing flap angle. This gives the shortest landing distance and the highest noise level. Where increase runway length is available, 1 6 degrees flaps could be ilsed. Velocity would increase to 55.5 meters/sec (108 knots) and thrust would rcduce to 818.4? n (184 1b)jenp;ine providing further reduction in noise.

Where take-off sideline noise is measured at multiple locations, the al- titude of 262 m (860 ft) produces the highest estimated noise. Conditions are sumarized for this altitude.

Climb rate of the QCGAT aircraft approximates current aircraft of similar mission. It attains an altitude of 106 m (3630 ft) at the take-off flyover measurement point.

ENGINE DES IGN This portion of the paper will touch on design objectives, noise, and emission considerations, engine cycle and engine description, and conclude with specific design featurec.

Sefore proceeding into the details of the engine design, a brief reviev of the design objectives is in order.

The ecological characteristics of an aircraft system are a direct reflection of the engine design. Careful attention to engine design details which impact noise and emissions is required to produce an engine that will became a welcome resident in a suburban comwmity.

Appeal of turbofans is indisputable. They swept virtually the entire corrrmercial carrier market in a period of twenty years. The same trend has started in general aviation with the larger size aircraft.

The rate of turbofan penetration into the smaller general aviation aircraft size is a function of the engine cost. This cost generally equates to simplicity of engine configuration. The result is basic; to be successful, it must bc simple in configuration. Take-off thrust should be sufficient to permit operation from the majority of general aviation airfields.

The mechanical design life goal should reflect the anticipated air- craft mission. Beech projected a useful aircraft life of 12,000 hours with an average flight cycle lasting 90 minutes, Our design goal wss to match this life without replacement of major parts.

Despite the best intentions of the designer, parts do break and it is desirable to be able to replace them conveniently. Modular ensice construction achieves this goal.

A 12.191-meter (40,000-foot) flight envelopc i s attractive for avoidance of traffic and bad weather.

The need for fuel economy goes without saying.

The larger engines for comsrc ial carrier aircraft have demonstrated sub- stantial advances in the technology of noise reduction. They have provided the recipe for quiet engine design which was used for QCGAT and is sumarized in figure 9.

Blend low fan blade tip speed and low fan pressure ratio with high fan bypass ratio.

The fan scator should t e set at least two fan blade chord lengths aft from the blade trailing edge. The quantity of fan stator vanes should exceed two times the number of fan blades to avoid inter- action of fan blade wakes with the stator vanes. Canted stator vanes are preferred.

Exhaust noise reduces with exhaust velocity, and turbine blade-pass f requency should exceed the audible range.

During the iterations which optimize an engine performance cycle, contin- uous attention is required to avoid adverse impact on emissions characteris- tics. Table I1 summarizes the primary causes for emissions along with the en- gine parameters which have a beneficial influence on emissions.

Unburner! hydrocarbons and carbon monoxide emissions are primarily a reflection of poor combustor efficiency at idle. Low combustor inlet temperature at idle aggravates the carbon monoxide emissions. To re- duce these two constituents, one would strive for very high c a b u s - tor efficiency at idle combined with elevated combustor inlet tem- perature. To achieve the higher inlet temperature, a campressor with poor efficiency at low speed if desired. This cumpressor should then be run as fast as idle thrust constraints will permit, and then bleed air to achieve even higher speed for the same thrust.

Whereas idle conditions have the primary influence on URC and CO, take-off conditions predominate in the creation of %. Generally, the higher the combustor inlet temperature at take-off, the more difficult the problem with NOX. Another important a x i a , !KIX and CO can usually be traded through conbustor design modification.

Either emission can be improved at the expense of the other to achieve the desired combination.

Comments, so far. have ignored engine bypass ratio. Emissions are produced exclusively in the core engine. The higher the bypass ratio, the lower the emissions for a given thrust rating.

A high-pressure compressor pressur? ratio of approximately 1011 under cruise conditions was selected as being achievable without compromise in con- figuration simplicity. Demonstrated component technology indicated two tran- sonic axial stages combined with a single centrifugal stage would be suffi- cient. Modest work input requirements for this compressor permit selection of a single-stage air-cooled twbine drive.

A n N4[ emission goal was considered the most difficult to achieve, with high-pressure ratio engines requiring a complex combustor configuration. Be-

cause y emissions increase with pressure ratio, this 1011 pressure ratio

selection also favored ccmbustor configuration simplicity.

Figure 10 shows the results of one of many parametric performance studies conducted during the cycle selection phase. This particular study was conduc- ted for 7620 meters (25,000 feet) cruise at 0 . 6 Mach Number. Engine perform- ance is plotted two w a ) s for comparison. The chart on the left provides bare engine performance as it would be measured in an altitude test chamber. The righthand chart modified the SFC coordinate to reflect installed specific fuel consumption. Here, losses associated with nacelle drag and weight are factored in as the nacelle size varies with engine bypass ratio.

Performance for a variety of fan bypass ratios are plotted on coordinates of specific fuel consumption and fan pressure ratio. In both figures, speci- fic fuel consumption is seen to reduce with increasing fan bypass ratio up to a ratio of 1011. Optimum fan pressure ratio decreases with increasing bypass ratio.

A fan bypass ratio of 9.6/1 at cruise was selected to limit required in- put work to the capability of a single-stage fan drive turbine. The corres- ponding fan pressure ratio was set at 1.35.

The engine was sized to produce in excess of 7117 N (1600 lb) of thrust under sea level static operating conditions. Sea level and altitude perform- ance are suunnarized in table 111. This perforntnce compares favorably vith even larger, more scphisticated engines currently in use.

Figure 11 schematicallv shows the engine configuration we selected. The gas generator section was not funded by the NASA QCGAT Program. R e engine has only six rotating cascades in total, and only two in the hot section where maintenance costs normally accrue. As such, it achieves the simplicity necessary to penetrate the medium aiicraft size general aviation cost barrier.

The configuration is a high-bypass turbofan composed of a single-stage fan, a gas generator section, and a single-stage, low-pressure, fan-drive turbine.

Initial compression is provided by the fan stage with the major<ty of the air bypassing the gas generator section to produce thrust directly, much as a small propeller would. Air flowing through the hub of the fan enter* the gas generator and is further compressed by the high-pressure compressor.

A reverse-flow annular atomizing combustor accommodated fuel burning and energy release. Hot gases take a second 180 degree turn before floving through the high-pressure compressor-drive turbine.

These gases then continue axially aft through the fan-drive turbine.

Power from this turbine is transmitted forward by a shaft that is concentric within the hollow gas generator shaft. Rotational speed is reduced by a re- duction gear to match the optimum fan engine.

The advantages of the selected configuration are numerous. Initial stu- dies projected attractive specific fuel consumption while maintaining the desired simplicity of only six rotating cascades. Compliance vith the recipe for low noise and emissions has been achieved. The reverse-flow canbustor permits packaging the gas generator turbine inside the combustor to achieve a short coupled engine, thus avoiding difficulties of casing deformation and shaft natural frequencies. The resulting engine center-of-gravity is close to the axial plane of the main engine mounts, simplifying installation require- ment s .

Use of the reduction gear permits individual speed optimization to achieve the best efficiency for the fan and the single-stage low-pressure turbine. Also, the low-pressure turbine may then operate at higher rotational speeds where blade pass frequencies, a common noise source, are outside the audible tone spectrum even vnder low-speed aircraft approach conditions.

Overall, engine configuration is shown cross sectionally in figure 1 2 .

External dimensions are approximately 610 m m X 910 nan ( 2 feet X 3 feet), not including the accessory gearbox.

~ e l i c a l reduction gearing introduces an axial mechanical load which op- poses aerodynamic loads on both the fan and the low-pressure turbine. This provides a significant reduction in thrust loads on the ball bearings in both the fan and low-pressure, turbine rotor systems.

The accessory gearbox is chin-mounted at the b o t t m of the main frame for esse of maintenance without core cowl removal. Accessory drive is provided from the high-pressure rotor spool via a conventional bevel gear mesh and through-shaf t with intermediate bearing support.

Figure 13 shows the modular maintenance features of the engine. The en- gine disassembles into four basic modules, as shown. The fan module includes fan, stators, reduction gear, and main engine frame. The core module contains the high-pressure compressor, its drive turbine, and the combustor.

Separation of the gas generator and low-pressure turbine modules allows visual inspection of all the hot-section components. Full disassembly of both to NASA representatives during the short period of a modules was demonstrated coffee break at one of our coordination meetings.

On the wing maintenance is virtually unlimited by engine configuration.

Hot-sect ion inspect ion, fan blade accessory replacements, gas generators and low-pressure turbine module exchanges are but a few of the options available to the operator.

The QCGAT engine fan module is depicted in figure 14. The fan blade tip diameter is approximately 559 mm ( 2 2 in.) with a modest tip speed of 335 meter/second ( 1 100 feet/second! at take-of f conditions. The ratio of stator vanes to fan blades is 2 . 4 6 for acoustic considerations. The distance between the fan blade trailing edge and the fan stator vane leading edge has been maintained at 2.1 fan blade chord lengths to minimize noise from rotating blade wakes. The fan stator is canted aft to maximize this distance for a given engine length.

The reduction gear permits high turbine rctat i onal speed producing a blade-pass frequency which is above the audible range, even under reduced power approach conditions.

Fan blade containment capability has been provided in the fan shroud design. Imbalance i l-ting from blade loss, has been a design criteria for the supporting strc -.

Hot oil sprayed into the hub of the spinner provides continuous anti- icing and additional oil cooling.

Figure 15 shows an assembled fan vheel. 15ere are 24 rugged long-chord fan blades which are designed to withstand bird impact without the support of a mid-span shroud. Fewer, long-chord blades were selected as being preferable to a higher quantity of short-chord blades incorporating midspan dampers.

This reduces wheel cost and avoids the performance penalties associated with midspan shrouds.

Figure 1 6 depicts the fan bypass stator assembly. In this design, the stator vanes are manually inserted into potted boots retained in the inner and outer shrouds. This feature permits individual vane replacement rather than returning the entire assembly for overhaul repair in the event of for- eign object damage.

The QCGAT mairr -.-ttl~ctural frame is shown in figure 17. The frame is inte- grally cast of aluminuni. Four engine mounting bosses are provided to permit selection by the airframe designer for top, side. or bottom engine mounting.

A cross-section of the QCGAT core engine is shown in figure 1 8 . The com- pressor and turbine stages are mounted on a hallow shaft which acts as a throughbolt furnishing the necessary clamping force for the rotor system.

The compressor rotor thrust load is carried by a ball bearing at the front end. A spring-loaded ball bearing at the end of the shaft permits ex- pansion while maintaining radial posit ion. Accessory drive is taken from this rotor by means of a bevel gear drive.

' I t r o ball bearings, supporting the fan-drive turbine rotor, are contained in the same h0usi.n~ which supports the aft high-pressure rotor bearing. This avoids the cost of iubricating and sezling individual packages. A concentric drive shaft through the high-pressure rotor delivers the fan-dr ive turbine power to the fan module.

All rotating cascades for both the high-compressor and low-pressure tur- bine rotors are integrally cast to reduce cost. Air cooling is confined to the high-pressure turbine.

Blade loss containment is provided throughout.

The combustor is a folded annular atomizing burner.

In a conventional atomizing combustor, an axial vortex is generated around each atomizer by swirling the air with vanes. In the QCGCT size com- bustor, sixteen atomizers and swirlers would have been required to attain even, circumferential, temperature distribution.

Figure 19 shows the "circumf erentially stirred" con£ iguration which was selected in preference to the conventional combustor. Primary air is admitted through slots in the liner header producing flow circulation in a circumfer- entially oriented vortex. Secozdary air jets, called "folding jets", enter the inner wall directly downstream of each atomizer and force the circumfer- ential vortex into a horseshoe shape as it flows downstream. In this manner, two downstream vortexes are created for each atomizer, and the required num- ber of atomizers is cut by one-half, to eight.

Prior testing has shown this configuration to be superior to conventional combustors in emissions characteristics. It also demonstrated significant margin in UHC and CO but was initially somewhat above QCGAT 134( goals. This permitted the trade-off of CO for NOX mentioned previously to assure achievement of NOX goa 1 s .

As a result of this intensive design effort, the first QCGAT engine was assembled in Gctober 1978. In figure 20, the core engine module is shown being connected to the fan module. Figure 21 is a front 3/4 view of the basic engine fully assembled. The addition of the test inlet bellmouth, plus the core engine cowling, is depicted in figure 22.

Figure 23 shows the birth of a new engine model installed in the test cell just prior to its initial test run. Fizlre 24 is an enlarged view of the engine installed. The first engine run was in October 1978. A 30-hour mech- anical verification test was cmpleted in April 1979. May and June were devoted to damping an undesirable resonance in the ring gear. Emission tests were conducted in July.

Figure 25 shows the engine at the acoustic test site during the acoustic testing phase of the program, which was completed in August. The demonstrator engine was then inspected, acceptance tested, and delivered to NASA in Octo- ber 1979.

NACELLE DESIGN A preliminary design in the flight nacelle was defined to establish a realistic baseline from which a ground test nacelle could duplicate the im- portant features at reduced program cost. Only the ground Lest nacelle was fabricated.

An artist's conception of thz flight n~celle is shown in figure 26. The nacelle is composed of the following sections: 1. An inlet duct to provide uniform flow into the engine 2 . A fan outer duct and core cowl to guide the bypass air around the engine 3 . A mixer assembly to force the mixing of hot, higher velocity core engine exhaust with the cooler, low velocity fan stream 4 .

A confluent mixing chamber preceding the final nacelle exit nozzle 5. An aerodynamically shaped outer skin, designed to minimize drag at the higher flight speeds.

A mixed-flow confluent exhaust system was selected because it reduces the peak axit velocity, improves propulsive efficiency, and reduces jet noise.

Noise attenuation treatment in the f o m of perforated acoustic panels has been introduced in the air intake section and in the fan duct outer wall.

Figure 27 shows the enginelnacelle mounting and maintenance access panels. The engine is designed to carry the nacelle aerodynamic and "G" loads. An airframe or nacelle yoke attaches to two points on the engine main frame plus an aft steady link. The entire nacelle is then carried by the appropriate engine flanges. Four access panels are provided for ease of main- tenance.

The nacelle aerodynamic contours, summarized in figure 28, are optimized for. low drag at 0.65 Mach Number, 10688-meter (35,000-foot) cruise condi- tions. The intake is designed for high cruise efficiency with modest compro- mises for static 20.6 meters/sec (40-knot) crosswind tolerance and pressure recovery at low-speed take-off conditions. A NASA/McDonnell-Douglas computer program for three dimensional flow calculations has been used for predicting inlet cowl flow conditions. This program computes flow over axisymmetric bod- ies at various flow angles of attack. Inlet and fan duct flow velocities are generally below 0.4 Mach Number.

The nacelle is circular in cross-section except for the bottom portion which expands into an elliptical section to house the engine accessories.

Boat tail angles vary from 1 4 to 18 degrees.

The QCGAT ground test nacelle which was used to explore noise and emis- sions reduction is shown in figure 29. The internal flow lines, flight inlet lip, and the exhaust nozzle are identical to the flight nacelle. External skin was eliminated to reduce program cost.

Flight-worthy hardwall and noise attenuation panels can be readily ex- changed as desired.

Three inlet lip configurations (figure 30) which were tested with the QCGAT engine are as follows: 1 . An inlet bellmouth for 1.oss-free baseline calibration 2. An exact replica of the flight nacelle lip 3. A lip designed to simulate landing approach inflow conditions.

The flight lip and the inlet bellmouth are compared in figure 31.

Figure 32 shows two views of the mixer nozzle. Studies showed seven lobes to be the optimum for our engine. We selected six, with a very minor perform- ance penalty, to avoid any possible seventh-order upstream excitation of tur- bine blading. Both shaker tests and engine strain gage testing showed satis- factory dynamic characteristics.

CONCLUSIONS Challenging objectives were set for the QCGAT aircraft preliminary design to respond to our assessment of general aviatioil needs for the 1980 decade.

The aircraft design achieves these objectives to provide six-place, long-dis- tance flight which will be at~ractive to both the user and the suburban com- munity .

Flight characteristics of this aircraft have been computed to define realistic criteria for measurement of ecological characteristics.

Reflecting on the engine and nacelle designs, the primary objectives of the QCGAT Program have been fulfilled. Large engine noise reduction technol- ogy has been successfully employed to the general aviation size engine. The QCGAT Program culminated in demonstration of QCGAT acoustic goals with margin.

A simplified approach to emissions was conceived in response to this pro- gram, and the QCGAT goals for emissions were very nearly achieved. Consider- able margin was demonstrated for both CO and YHC emissions and NOX was within 1 percent of the goal.

QCGAT has given birth to a new engine which is designed to serve the needs of general aviation in the 1980's. While still in its infancy, it has demonstrated attractive performance by current standards. Further development tuning will be required to achieve its full potential which is reflected in the QCCAT performance goals. Component tests have verified the long term ob- jectives. However, turbine rematch is required .to recoup the inherent config- uration perf ormarice.

Versatile ground test nacelles were created to investigate ecological desig parameters. Acoustic panels versus hardwall and mixed exhaust versus split streams were tested under this program.

TABLE I BENEFITS FROM ADVANCED AIRCRAFT DESIGN

FUEL I S T R U C I U R I Y D I QROIS 1

/ mcYn

PROPULSION WEtoWr.

CONFI~UR*TION r::fZj wsanr, ra grq ( L m )

' 1.368 1 2.518 4,548

Current

Aircraft 1 (3.016) (5.551) 1 (9.8601 1

I

2 390 4,013 :8.848) r Use Composites 3 . 5 3 (2: 53) (4.3GO) (7.800) Supercritical Wing SAVINGS 22% 32% 23% L -.

TABLE I1

DESIGN CONSIDERATIONS FOR EMISSIONS

EMISSION CAUSE Unburned Hydrocarbons Combuslion inefficiency Carbon Monoxide Inadequate: Residence Time.

Temperature, Ef!iciency NO:: High Residence T~me/lemperature Smoke Local Rich Zones 8 Benef~cial Engine Characteristics: High Combustion Efflc~ency at ldle H:gh Combustor Inlet Temperature at ldle (High Speed. Low Compressor Eff~ciency. Bleed!

TAB= 111 PROPOSED QCGAT ENGiNE PERFORMANCE FLIGHT CONDITION . -. -. . - - Y A W CRUISE TAKEOFF 0.6u.nis N PIRAYElER sea LW Static ( 2 5 . ~ FT) Thrust specific 0.0367 Kg/N/hr 0.063 (Kgm hr) Fuel (0.360 lbllb hr) (0.626 IWlb hr) Consumption

NEW GENERAL AVIATION AIRCRAFT

FOR THE 1980 DECADE

NUMBER N O M I N A L OF AIRCRAFT AIRCRAFT

COST, $ x lo6

A

- 1

I 1 I 1 4,000 6,000 8.000 10,i)Cil 12,000 L I I i I I J 2.000 3,000 4,000 5.000 Kg AIRCRAFT GROSb WEIGHT Figure l

QCGAT R E STATOR

Figure 16

QCGAT ENGINE MAIM fRONT FRAME

AVCO LYCOMINC QCCAT PROGRAM DESfGN CYCLE, DEMONSTRATED PERFORMANCE AND EMISSIONS Phil Fogel a d Angelo K d i e r Avco Lycoming Divisiar SUMMARY Lycoming was awarded a NASA contract to design and build a quiet, clean, general aviation turbofan (QCGAT) using existing techno- logy for noise and emissions reduction, In addition, to the noise and emissions considerations, the Lycoming QCGAT engine was designed to provide both minimum fuel consumption in cruise and maximum take- off capability. The engine, which was built and tested a t Lycoming, has met and, in some cases, surpassed the design goals for emissions.

The engine program has also demonstrated that emissions and noise re- duction technology can be effectively applied to small turbofan engines without significant performance penalties.

This paper describes the basis for the cycle and coxr~ponent selec-

tion, for the Avco Lycoming - NASA QCGAT engine, and the resulting

demonstrated performance and emissions of the complete engine. An artist's conception of a cut-away view of the propulsion system is shown in figure 1.

The Avco Lycoming QCGAT engine is a high bypass ratio, twin spool turbofan engine of modular design. It incorporates a front fan module driven by the Ll SlOl core engine modified, a s required, to achieve the QCGAT goals. The engine i s housed in a nacelle incorporat- ing full length fan ducting with sound treatment in both the inlet and fan discharge flow paths.

Design goals of components developed under this contract and results of component tests a r e presented, herein, together with full engine test results.

In the emissions portion oi this paper, the rational behind the combustor design selected for the Avco Lycoming QCGAT engine is presented a s well a s the test results. Total system (engine and nacelle) test results are also presented.

Lycomingl s goal under this contract r r s not only to demoartrate the transfer of state-of-the-art acoustics and emissions technology currently used on large engines to smrlf engines, but t o build this around a high performance engine aad airframe system attractive for the 1980's and beyond, It is clear that a high performance fan engine integrated with an advanced airframe design concept is advantageous primarily for high performance t w i n engine aircraft currently propelled by piston o r small turboprop engines in the 373 (500) t o 746 (1000) kilcwatt (shaft horsepower) class. This segment of the market which has recently shown a strong growth, is expected to continue, especially with the introduction of a quiet, clean high performance aircraft rhich offers the highest benef<t, in terms of noise and pollution reduction, for those communities living at airport boundaries The engine installed in the aircraft must offer modern high perfor- mance, economical cruise speeds beyond the reach of present turboprop applications a d a range over 2224 kilometers (1200) (nautical miles).

Prime cruise altitude was targeted for 7620 m (25,000 ft ) a t Mach 0.6.

with a potential to climb and cruise a t l2,192 m (40,000 £to ) -

These targets were based on data received from aircraft operators PERFORMANCE GY CLE ANALYSIS Design and trade-off studies were performed to define the optimum cycle in terms of noise, emissions and performance. The rational used t o select the overall engine characteristics and the fan configura- tion is exemplified in figures 2 and 3. The optimization study assumed component efficiencie s expected at the critical operating conditions: sea level, static take-off 7620 m (25,030 f t ) Mach 0.6 cruise

Z524 m (5.000 f t ) , hot day single engine climbout

Figure 2 shows engine specific fuel consumption (SFC) versus fan pressure ratio for selected values of bypass ratio. As shown, there i s a point of minimum specific fuel consumption for each fan bypass ratio. Higher bypass ratios coupled with lower fan pressure ratios results in lower specific fuel consumption. This, however, has to be moderated because of two factors: installation losses and mechani- cal complexity. An increase in engine bypass ratio results in increased engine-nacelle drag and weight, which in turn causes an increase in air- frame weight or reduction in payload. Also, further increase of the by- pass ratio would require a variable geometry exhaust nozzle to prevent excessive fan unloading with resulting loss in fan cruise efficiencies.

The effect of an increase in cycle pressure ratio on SFC is shown in figure 3 . Although, increasing cycle pressure ratio d e c r e a s * ~ SFC, any increases in high compressor pressure ratio beyond approximately 10.2 would require the added complexity of an additional low pressure turbine stage.

As a result of the design study, an initial design bypass ratio of 9.6 and high compressor pressure ratio of 10.2 were selected. Installa- tion weight and nacelle drag effects were considered.

The iz-pact of the selected cruise design point on the maximum thrust, at the critical single engine climbout condition 3 0 8 ~ ~ (555%)

ambient day a t 1524 m (5,000 f t ) , 69.5 m/ sec (135 knots) was exam-

This flight requirement was used to size the engine.

ined.

It was found that the 7620 m (2 5,000 f t . ) Mach 0.6 design point.

when lapsed to 1524 m (5,000 fb 1 , produces a maximum thrust for the selected bypass ratio.

The selected design cycle is presented in table L The changes in the engine parameters, shown in the table 1, from initial perforxllance analysis were caused by detail component design and final cycle optimi- zation for maximum thrust at the single engine climbout condition.

T h QCCXT engine installed performance goals for the two prime flight conditions are shown in table 2. This installed performance is with the nacelle system including the flight lip, mixer nozzle and acoustic treatment The sea level static take-off thrust i s 7166 N (1611 1bf) and specific fuel consumption i s 0.037 kg/ h r / N (0,363 lbm/ h r / For the 7620 m (25,000 ft. ) Mach 0.6 cruise, the thrust is 2157 N lbf).

(485 lbf) and specific fuel consumption i s 0.064 kg/ h r / N (0,628 lbm/ h r / lbf) A mixer nozzle, reference 1, was chosen for the engine configura- tion because of acoustic and performance reasons.

Figure 4 presents the estimated variations of specific fuel consumption, along an engine operating line, with total net thrust at the selected cruise condition, for the split and forced mixer exhaust systems. As shown, a potential performance gain, at the cruise thrust, of approximately 3.0 percent could be realized with a mixer.

COMPONENT DEVELOPMENT AND TEST Core Engine Definition The Avco Lycomiag LTSlOl turboshaft engine w a s selected a s the basic core for QCGAT engine. Core component modifications required, t o meet QCGAT design goals, were Lycoming funded.

Component Development The major components developed, under the NASA contract, were thc fan module, reduction gearing and the nacelle system which includes the forced mixer nozzle. The fan and M C ~ & were designed with tow noise as a primary criteria.

In addition, combustor system modifications were made, as re- quired, to meet the emissions goals.

Core Compressor The core compressor w a s tested to establish mechanical and aerodynamic performance with the turbofan inlet duct The compressor performance and surge characteristics with pressure distortion a s measured during the fan component testing were also established.

The rig test results showed that the compressor efficiency was within L 0 percent of the design goal The compressor showed high tole r a c e to pressure distortion produced by the fan.

Also, the turbofan inlet duct caused a reduction in airflow to the compressor of L 0 percent at the QCGAT operating conditions- Gas P r d u c e r Turbine Rig tests on the initial gas producer turbine hardware confirmed that the design efficiency cr this stage was met within L 0 percent.

However, the nozzles were substantially larger in flow area than design.

An attempt was made to correct for flow size, by reducing the annulus area formed by the inner and outer wall contour. This corrected the flow area prcblem but caused cascade losses which reduced stage performance by approximately 3 points.

In addition, the interturbine duct pressure 10s se s increased because of a resulting change in the turbine exit swirl angle.

A redesign of the nozzle and rotor, to recover gas producer efficiency, was completed in July 1979 and the revised hardware is being procured Fan Component A a experimental evaluation of the QCGAT fan module has shown that the bypass performance has exceeded design goals. At the design pressure ratio (L 38) and speed (U, 200 RPM), stage polytropic efficiency of 0.875 was demonstrated. This exceeded the design goal efficiency of 0.870. Bypass airflow at this point was 33.7 kg/ sec ( 7 4 3 lbm/ sec) compared with a goal of 33.6 kg/ sec ( 7 8 0 lbm/ sec).

Limited distortion testing was done to insure mtisfactory engine operatioh The response of a turbofan to inlet distortion is of prime importance from the viewpoint of aerodynamic performance and mechan- ical integrity of the blader Significant distortions occur in aircraft installations as a result of intake flow separation induced either by crosswinds o r high angles of attack.

The Lycoming QCGAT fan rotor demonstrated very good aerodyna- mic and mechanical performance under inlet distortion conditions which are representative, o r in excess, of those found in typical turbofan installations.

Low Pressure Turbine The low pressure turbine, which was not rig tested, appeared to perform as anticipated based on measured engine data.

COMPONENT STATUS SUMMARY Engine performance estimates obtained from math model simula- tions, based upon component test re sults, showed that further component development of the core, which was initiated in the spring of 1979, was required to achieve performance goals.

However, a s a result of the analysis, it was concluded that the Lycoming QCGAT engine was a viable vehicle for demonstrating noise, emissions and specific fuel consumption improvements which were the program* s objective s.

FULL ENGINE TESTS Referee Configuration Following the component rig tests, the full engine and nacelle system tests were conducted. Two engine configurations have been te sted. The referee configuration consists of a calibrated bellxxmuth followed by a straight inlet duct to the fan shroud as shown in figure 5.

In the exhaust system, the bypass and core flows are physically separated (see figure 6). Separate exhaust nozzles permit individual change of fan pressure ratio and variation of the power split between the fan and ccre.

Test Nacelle Configuration The QCGAT test nacelle configuration is shown, in figure 'I, with the flight inlet lip and diffusing duct which is mounted to the fan mhroud.

The flight lip can be readily interchanged with the bellmouth or the approach simulator inlets Details of the test nacelle are shown in figure 8 . The diffusing duct following the inlet contains interchangeable hardwall o r acou~tically treated softwall liners. The nacelle rear section consists of a core cowl covering the core engine while providing a smooth aerodynamic inner wall contour for the : i n flow surrounding the core. The common mired exhaust nozzle clamps to the rear face of the fan frame and contains the removable duct portion of either hardwall or softwall panela Engine Test Plan Various combinations of the two basic engine configurations, the referee and te st nacelle, were tested during the performance calibration sequence.

Table 3 shows an overview of the 7 prime engine configurations which were tested in order to determine the performance characteris- tics of the engine and nacelle system components. Prior to these tests, a baseline engine configuration was tested with a calibrated bellmouth coupled to a constant area duct and split exhaust The first three configurations, listed in table 3, with the split, or referee exhaust system, were tested with the diffusing flight inlet duct and the various interchangeable inlet l i p s A l l tests with the split exhaust were performed without fie The referee configuration with a bellmouth inlet was acoustic panels.

also used for the emissions sampling.

The test nacelle configuration with the mixed exhaust was initially tested, for performance purposes, only with the bellmouth inlet. First, tests were conducted with hardwall panels in the inlet and fan bypass Then acoustic panels were placed in the inlet only. Finally, exhaust.

the engine was tested with acoustic panels in both the inlet and fan by- pass exhaust The installed performance demonstration was with the flight nacelle inlet, mixer nozzle and full acoustic treatment.

Referee Engine Tests The purpose of the initial tests with the referee configuration w a s to evalclate mechanical engine operation and stress levels on fan and gear components Subsequent tests using the referee system, were conducted to eva- luate overall engine and component performance prior to evaluating losses associated with acoustically treated nacelle system. Variations in performance attributed to the mixer system was also to be determined The purpose of these tests were twofold: first, to establish a base calibration for determining component performance. Secondly, to eva- luate inlet pressure losses associated with the diffusing duct coupled to the various inlet lips. As previously stated, emissions sampling was also conducted using the split exhaust configuration.

Detailed analysis of test data has indicated that the diffusing duct and various inlets had a negligible impact on the overall engine perfor- mance. The engine test results with the referee configuration confirmed the predicted engine performance.

Nacelle Engine Tests Following the referee system performance and emissions tests, the installed nacelle test sequence was conducted. The purpose of these tests was twofold: first, to establish engine performance with a mixer nozzle; second, to evaluate the impact of the inlet and fan bypass exhaust acoustical panels on engine performance. After the performance evalua- tion tests, the engine was transferred to the acoustic test site for noise evaluation.

Table 4 shows a comparison between the demonstrated installed engine thrust and specific fuel consumption with the design goals.

The measured static thrust and specific fuel consumption are 6485 N (1458 lbf) and 0.0400 k g / h r / N (0.392 lbm/hr/lbf). The cruise per- formance was estimated based upon engine static test data and com- ponent rig test results.

PERFORMANCE SUMMARY Engine test results indicated that the acoustic panels, used for noise reduction, had a negligible influence OF, the overall engine per- formance. The estimated cruise performance of the Avco Lycoming QCGAT engine, in terms of specific fuel consumption, is approxirnate- ly a 10.0 percent improvement over currently available small turbofan engines in the 13,344 N (3000 lbf) or less thrust class.

Also, although the program performaxe goals were not achieved, the loss in engine performance has been identified a s deficiencies in the turbine section of the core engine. The performance of the fan, which was developed under the NASA contract, exceeded the design goal s .

A redesign of the affected hardware has been completea under a separate Lycoming funded program. Rig tests are scheduled to be conducted to evaluate the redesign a s soon a s the hardware i s available.

EMISSIONS Emission Standards In 1970, Congress passed the Clean Air Act. This Act, which was to be effective in 197.. , directed the Environmental Protection Agency to e stablish emissions standards applicable to aircraft, These standards, reference 2, for small turbofan aircraft, which have now been abandoned by the EPA, were kept as NASA goals for the QCGAT engine program.

To achieve these emissions limits, the basic combustor design used in the LTSlOl engine, references 3 and 4, was selected.

Combustor Design This design, which i s a circumferentially stirred combustor, is shown in figure 9. In principle, the primary air is admitted through slots in the liner header producing flow circulation about a circumferen- tial mean line. Air jets, called "folding jets" entering through the inner wall reinforce the primary zone recirculation, and the vortex fills the full annular height of the liner.

The vortex spreads circurnferentially in both directions and i s forced to turn in the axial direction on either side of the folding jets and the mean path of the combustion zone flow vortex takes the shape of a horseshoe. The number of fuel injectors i s thereby reduced by one half, compared with normal practice, because of L !s unique combustor primary zone aerodynamic concept.

Emissions Projections Emission measurements, for this type of comtustor, attained from the LTSlOl engine were available for use in predicting emissions for the QCGAT performance cycle. Table 5 shows the estimated emissions value s, for the QCGAT cycle, with the production LTSlOl conlbustor.

These EPA parameters were generated for a take-off and landing cycle for class T1 aircraft (reference 2).

These emissions projections indicated that further development of the LTSlOl c o m b ~ s t o r was required to reduce smoke. The hot end durability was in question because of the more severe operating condi- tions of the QCGAT engine.

Combustor Modifications Airblast injectors, which replaced the dual orifice injectors, were selected to reduce smoke. The introduction of the airblast injec- tors also increased combustor efficiency and oxides of nitrogen (NOx) a t idle.

Increasing the ccmbustor pressure drop for temperature distribu- tion control, also increased NOx and combustor efficiency while appre- ciably decreasing carbon monoxide and unburned hydrocarbons. This i s typical of the impraved primary zone mixing, which results from the higher pre s sure drop.

Air partition modifications were then made, a s required, to meet the design goals for NOx.

Figure 10 presents the effect of a j r partition modifications on NOx.

Unburned hydrocarbons and carbon monoxide we r e within the goals in all tests. Initially, the NOx slope for the LTSlOl combustor was a s predicted, and met the goal. However, a s the combustor pressure drop was increased to reduce smoke, NOx increased.

Air partition xnodifications, a s previously stated, were then made to meet the NOx emissions goal.

The final selected QCGAT liner, which met the goal, has a slightly steeper slope than the initial configuration.

The Lipfert correlation, reference 5, for conventional combustors i s shown for comparison.

Emis sions Sampling Development and initial emissions testing of the combustor was cond~lcte? in the laboratory. After the laboratory tests, the QCGAT liner was transferred to the engine for demonstrated emissions sampling.

The e.missions test probes were installed a s shown in figure 1 1 .

The probes, which a r e cruciform-shaped, were set at two angular One probe measured along the horizontal and vertical axes.

positions.

The othtr probe was rotated 45 degrees.

Table 6 i s a comparison of the =missions test results with the NASA goals. Measurements from the engine test showed that the unburned hydrocarbons were 60 percent lower than required.

The carbon monoxide was 30 percent lower, oxides of nitrogen L 0 percent higher and the smoke number 50 percent lower than the goal.

EMISSIONS SUMMARY The emissions requirements of the QCGAT engine have been met and, in most cases, surpassed. The QCGAT com5ustor ~ 5 3 v i d e s sub- stantial margin for carbon mc.noxi~ e and unburned hydrocarbons emi s sions while meeting the goal for NOx within the scope of the program.

The com5ustor system modifications required to meet the emissions goals had a negligible effect on engine performance.

CONCLUSION The QCGAT developmc rlt program was designed to demonstrate, as well a s advance, state-of-the-art technology with regard to noise, emissions and fuel economy of small turbofan engines used in general aviation- type aircraft.

The program objectives, in terms of ern is sic..^ and fuel consumption, were m e t With the knowledge and experience gained through the NASA- Avco Lycomin8 engine program, the thrust and SFC goals, although not demonstrated within the time period of the program, a r e achievable with additional component development.

REFERENCES 1. 3. F . Hurley, L. I'Anson, and C. Wilson, 'T?csign of an Exhaust Mixer Nozzle for the Avco- Lycoming Quiet, Clean General

Aviation Turbofan (QCGAT), I' Avco- Lycoming Div. , Avco

Corp., Proj. FEDD, NASA CR-159426, 19'18, 2 . Emissions Standards and Test Procedures, Title 40, CFR Part 87, published in the Federal Register, July 17, 1973, 3, U. S. Patent 3.67 1,17 1, "Annular Combustors", Brian W, Doyle.

4. U. S. Patent 3,645,095, "Annular Combustorn, Jerry 0 . Melconi~n Lipert, F. W. , "Correlation of Gas Turbine Emissions Data", 5.

ASME Paper 72-GT-60, RESULTS Of DESIGN STUDY A L r m r o E = ~ 2 S , Q b O r r ) . Y A C H = a $ Fm pressure RUio 1.36 Cycb Pressure Ratio 13.7 Cote Compessx Pressure Ratio 10.3 Tht'uWlota A i W . N/kglt4c(lbfflbi%x) 113.7(11.6) 0 y p . s ~ Ratio 9.4 Table 1, AVCO QCGAT PERFORMANCE GOALS (STANDARD DAY. I N S T U D ) SEA L E M

STATIC - - Y ( I ~ ---

Rating Takeoff Cruise Thrust. N(lb0 7166(1611) 21576485) SFC. kg/hrlN(lbm/hrilbf) 0.0370(0.363) 0 W ( O . 6 2 8 ) Table 2.

ENGINE CONFIGURATIONS TESTED (P€RFORYAWCE TESTS) Hardrvall HardwaU Softwall Hardwdl Softwall Sonwdl Softwall SottrrJl Table 3.

AVCO QCGAT PERFORMANCE (STANDARD DAY. INSTAUED) ~W SEA LtvEL TIIEOCF Thrust. N(1bf) 7166(1611) 6485(1rSB) SFC. kgArlNlllbihrrlbbf) 0.0370(0.363) O.M00(0.392) D L W l CIIwsc v 2 5 , O o (I) YIeI 0.e Thrust. N(1bf) 2197(485) 1850(416)' SFC. kg:hr/N(lbm/hr/lbl) 0 O M ( O 628) 0.074(0.723)' 'Estimated trom Static Data Table 4.

INITIAL ESTIMATED =GAT EMISSIONS Lls 101 NASA G & ' 0.015 0.266 O . l O S 45.0 ( 1 . 6 1 ( 9 . 4 ) ( 3 . 7 ) *g/kNs (lW1000 ibt thrust hrcyde) Table 5 , QCGAT EMISSIONS RESULTS Goal* Engine T e s t ' 0.017 0.193 0.106 24 (0.6) (6.8) (3.75) Engine Test/Goal 0.4 0 . 7 1 . 0 1 0 . 5 'glkNs (!bm/1000 I b f thrust hr-cycle) Table 6 .

AVCO LYCOMING QUIET CLEAN GENERAL AVIATION TURBOFAN ENGINE Craig A . Wilson Avco Lycoming Division SU MMAR Y Avco Lycorning participated i n the NASA QCGAT program b.9 developing a fan module using an existing turboshaft. engine. The fan was designed using the latest i n large engine noise control technology. A mixer was added to reduce the already low exhaust gas velocity. A nacelle incorporating sound treatment was provided for the test engine. The noise prediction model was used through the design process to evaluate the various design alternatives. Acoustic tests \-?ere then made to verify the prediction and identify the noise characteristics of the fan, core, jet, and eound treatment. Anrzlysis of the recorded data yielded close agreement

, . I the expected results. Core noise, a s was expected, was the predom-

in& ..c source of noise for the QCGAT engine. Flyover noise predictions were made which indicated that the Avco Lycoming QCGAT engine would meet the goals set for the QCGAT program.

INTRODUCTION The Avco Lycoming Quiet Clean General Aviation Turbofan ennine program was designed to demci~strate the latest gas turbine engine noise control A considerable amount of technology in a general aviation size engine.

work has been done to identify the design features that offset the generation of noise. And this work is still in progress a s can be witnessed by the complexity of the facilities at the Lewis Research Center and elsewhere.

The majority of this work, however, has been directed toward the commercial transport class of engines. The QCGAT program was designed to broaden the scope of this effort to include the general aviation size The significant features of the QCGAT design a r e the low exhaust engine.

velocity achieved by a high bypass fan design, the use of a mixer, no inlet guide vanes, subsonic fan blade design, large blade to vane spacing, a high vane to blade ratio, the acoilstical lining of the inlet and discharge The nacelle and aircraft play fan ducts and the use of a long inlet duct.

an important roll in incorporating these features in the overall acoustic design. F o r example, the mixer i s enclosed in a shroud formed by the nacelle. The fact that forward airspeed mitigates the amount of jet noise generated has also been factored into the design. These features were optimized for the QCGAT aircraft based upon the results of our prediction of the acoustical performance of the engine aircraft system and the impact of each component on the overall design. The QCGAT noise goals were selected by NASA to force a design that included the latest noise control technology. We responded with an engine design that consisted of adding a new fan design module that incorporated the latest noise techniques of one our turboshaft engines. Our original estimates of the engine noise emissions, based upon that design, are shown in figure 1. Our analysis indicated that the takeoff noise levels would be 4 EPNdB below the goal, the sideline 6 EPNdB below, and the approach 9 EPNdB. This analysis indicated that the core would be the dominant noise source at each measure- ment position, with the fan contributing to the approach noise and the jet contributing to the takeoff noise levels. Note that the goals a r e given in terms of aircraft flyover noise. The takeoff measurement point lies 3.5 nautical miles down range from brake release with the aircraft flying directly overhead, The sideline measurement point also lies down range on a takeoff but is displaced 1/4 of a nautical mile to the side and consists of a series of points in order to determine maximum noise level. The approach measurement point i s located under the landing flight path at a point 1 nautical mile from the runway threshold. As the approach gli :,?

slope i s defined as 3 O the altitude of the aircraft over the measuremt .

point i s 270 feet. Thus we had to consider aircraft performance in the engine design. For this we worked with the Beech Aircraft Co. to define the characteristics of a twin engine QCGAT powered aircraft. With respect to noise, the rate of climb at takeoff, the power required at approach and the geometry of the wing were determined. Airframe noise, however, was not included in our noise estimates, The design and performance of this aircraft plays an important part in the As has already been discussed, noise emissions of the QCGAT engines.

the approach speed and takeoff performance can vary to meet the market requirements ot the aircraft. For example, a lower approach power could have been used that would have resulted in lower approach noise levels.

As the approach noise levels were predicted to be low, we felt that a small penalty was acceptable to reduce field length requirements. This ail1 allow the aircraft to be certified for use at a large majority of the existing air fields in the United states.

Gas turbine engine noise source identification and control, figure 2, starts with the engine and its geometric and performance characteristics from which prediction of its noise emissions can be made. The engine noise is subdivided into five distinct noise generating mechanisms. They are the fan, compressor, combustionpprocess, power turbines, and the tur- bulent mixing of the exhaust jet with the ambient air. The majority of the work done to advance the state-of-the-art gas turbine and aircraft noise identification and prediction was, and is, being carried out by NASA as part of their Aircraft Noise Prediction Procedures (ANOPP), References 1 thru 5. This work has formed the basis of our noise prediction efforts.

Of course, we have made certain modifications in order to more accurately reflect cur experiences.

These prediction procedures a r e continuously updatea to more accurately predict the engine noise levels.

Given the aircraft performance and applying flight effects aircraft flyover noise can be calculated.

ENGINE DESIGN AND NOISE PREDICTION The first task was to design a fan module for the engine. This involved several iterations to access the design alternatives. Fan noise reduction was achieved through the use of a ldw pressure ratio fan to reduce blade loading and noise generation. This has to be part of the fan design f r , ' its conception. Other fan design features a s shown on figure 3, have been shown to recult in quieter fan designs for the large turbofan el?. nes Specifical!y, the fan blade tip speed should be designed to be subsoni Thus multiple pure tones, o r 'Buzz Saw Ncise" a r e eliminated altogethe-.

The design relative tip mach number for the QCGAT engine i s 1.13 which yields a subsonic value at all sea level operating points. The distance separating the blades from the exit guide vanes should be large when corn- pared to the blade width to reduce rotor stator interaction noise that i s expressed in fan broadband noise. We used a value of 2.3 for this ratio.

The ratio of fan vanes to blades was optimized at a value of 2.5. This wzs to elir~inate what are known a s spinning modes that p r o ~ a g a t e at the blade passing frequency fundamental. In addition, inlet guide vanes were not used in our fa= design. To h r t h e r insure that :nlet tcrbulence was reduced, a long inlet duct was included in the nacelle design. These features were accounted for in our prediction of the fan noise levels. Our prediction indicated that the fan would be a contribu~or along with the core only to the approach power levels. By identifying the effect G the various alter- natives with the aid of our prediction procedures we were able to maintain this balance to achieve a low- noise sigzature at approach.

An aircraft engine operates differently in flight than it does tied down to a test stand. Its noise characteristics also change. In flight, the a i r inflow is streamlined due to both flight cleanup affects of the forward air speed and the absence of ground turbulence that influence the generation of fan noise, particularly the tone at the b?ade passing frequency. Forward flight however. has its greatest impact on the generation of jet noise (see figure 4). It acts to reduce the relative velocity between the exhaust and the ambient air, This can play an important part in the overall design of the engine aircraft system. For example, the airspeed at takeoff is in part determined by the length of runway availability. A longer takeoff roll rould permit a higher takeoff speed. Consequently, the same jet noise level and relative velocity could have been achieved using a higher exhaust velocity.

k s tk- aircraft flies past the observer, the soume -garic; in both time and spectral content. Dynamic amplification acts to increase the noise level as the ;kcraft approaches, and reduce the ncrise levels as it recedes. Then there is the doppler effect that imparts a frequency shift to ths noise spectrum as the aircraft flies past, These phenomena must be accounted for to accu- rately predict the perceived noise of the aircraft.

It is the reduction in jet nsise that has the greatest potential for noise reduction, Jet noise i s thus the second major element in the QCGAT engine design A high bypass fan design i s used to reduce the exhaust velocity and therefore reduce the noise generated by :he turbulent mixing of a high vc-ocity jet, The jet noise predictions indicated the jet would contribute to the takeoff noise and possibly cause the aircraft engine combination to exceed the limits set by the QCGAT gods at the reduced thrust and alti- tude condition. The calculations showed that the differences between the core engine and the fan exhaust gas velocities would contribute to t h s turbulent mixin;, noise (see figure 5).

A six clement mixer was then designed to mix the core engine and fan exhaust gas co yield a single low velocity exhaust jet, The mixer, however, is not entirely free of side effects. Pre and post mixing turbulence can be an additional source of noise that has to be dealt with, These noise sources can be reduced by the addition of a shroud. In our design we provided that shroud by extending the nacelle considerably past tF.e mixer to affect a Setter mix.

The high bypass fan and mixer were designed to reduce the jet noibe component to a noise level below that of the core when forward flight effects cause reduction to occur in the jet noise. That leaves the core noise component. Core noise means the noise generated by the combustion process. The ergine compressor and turbine noise were predicted to be above the audible range. Their xroi-se sources do not contribute to the perczj -ed noise of the QCGAT eng:.u.: and were not considered in the design.

Core noise models have, for the most part, been emperically derived.

the ANOPP routine was found to be adequate for our turboshaft engines.

This predicticn model uses combustor mass flow, temperature rise, and pressure drop a s the basis for predicting core noise (see figure 6).

Emperical data also suggest a 7 to 10 dB reduction for the turbof= version of this model. Core noise is now recognizcd a s a major source in turbofan engine noise and is the focus of much research. We a r e working on this both in-housc and with NASA. However, we have not included any new coxe noise control features, Some of the design modifications for emissions may have contributed to higher core noise levels. As our prediction showed from the beginning, the core was going to be a significant contributor to the noise characteristics of the aircraft. Consequently, we felt that further fan and j et noi sc reduct ion would have be en unwarranted.

It has been long recognized that the fan inlet and discharge ducts of the engine nacelle (see figure 7) offer ideal locations for the installation of sound treat- Absorptive material a r e ment to absorb the noise generated by the fan, particularly efficient in absorbing sound energy in the high frequency region where niuch of the acoustical power radiated 'by the fan is concentrated.

In addition, the sound treatment can be constructed of flight worthy mate rials that add little weight to t h e aircraft. Finally, the theory and experience of designing sound treatment panels a r e sufficiently sophisticated to accurately Consequently, predict the results that will be achieved by a particular design.

sound t reat ment panels we re en~ployed for the QcGAT engine nacelle to determine the benefits that would be derived from their incorporation in an aircraft design, The sound attenuaticn requirements we re Jete rmina-d by comparing the The approach predicted noise levels with the QCGAT program goals.

position represented the only point where the fan noise was predicted to In addition, the frequency of the contribute to the aircraft noise levels.

blade passing t o w at approdch is located in the more heavily weighted part of the audible spectrum. Consequently, the appioach power point was At the other positions the selected for the design of the sound treatment.

fan does not contribute tothe aircraft noise levels. The Lockhecd Cali- fornia Company was contracted to design the sound treatment for the nacelle. Given the dimensional limitations the nacellc imposed upon the placement of thc sound t reatmcnt and th2 cnginc operating parameters at approach, Lock. :w-ed pcncratcd a set of dcsign curvcs from which the Thcsc curves uvcrc based upon an analyti- sounc~ treatment -**as designed.

cal and emperically derived solution tc what are known a s the convected wave cquations Thc.sc. cquations dcscribc thc- sound gcncratcd by the fan This acoustic a s modcs of acoustic cncrgy rotating with and against the fan.

Thr* physical cncrgy can only propagate undcr certain boundary conditions.

characti-ristics and operating parameters form these boundary conditions and determine which modes will propagate. Lockheed performed this analysis and recommended a design. We then took this design to our Nacelle contractor, Avco Ae rost ructure s, for fabrication.

The Lockheed design recommendations are shown in figure 8 . Their design was for a single degree of freedom panel for both the inlet and dischargc ducts. This design consists of a solid backing plate held 16 mm ( 5 1 8 in) off an inner plate perforated to a 57~ open area. A honeycomb cell struc- ture material separates the inner and outer plates. The inlet panel is 330 mrn (13 in. ) long to f i l l the available space in the inlet duct. The dis- charge sound treatment consists of a 45.7 mrn (18 in, ) long panel on the outer duct wall. The inner duct wall formed by the core cowl was not treated. The disch-rrge panel was terminated before the start of the mixer to simplify the design Other wise the radiant heat frcm the mixer would have necessitated the selection of more expensive materials.

The predicted insertion loss for t k fan inlet sound treatment panel at the approach and takeoff points a r e shown in figure 9. The sound treatment a s mentioned earlier was designed fcr the approach condition At this power setting the peak attenuation i s made to coincide with the blade pass- ing frequency. The insertion loss is higher at the takeoff condition due to the increase in air flow through the engine. The blade passing frequency at takeoff is also higher. The result is an attenuation approximately the same as that for the approach condition.

The predicted attenuation for the fan discharge treatment are in figure 10.

The duct width between the inner and outer wall makes th2 treatment more effective even though the inner wall is not treated.

The test nacelle and sound treatment panels were fabricated by Avco Aerost mctures in Nashville, T e n n The test nacelle was designed without the outer skin and to take insert panels in the fan inlet and discharge ducts where ordinarily the sound treatment would have been placed. Two sets of inserts were fabricated. Each was derigncd to be of one piece to ease removal and installation during testing and to be rigid enovgh to maintain the dzsired wall contours. The panels were of sandwich construction with a honeycomb structure separating the inner and outer plates. The thickness of the honeycomb was determined by the Lockheed sound atten-lation require- ments. One set was fabricated with a solid inner plate, and one set (see figure 11) was fabricated with an inner plate perforated to achieve a Sff0 open area. This way we could test the engine with and uithout sound treatment in the nacelle.

The small .ddius of the inlet and dischargc ducts limited to the dcpth of honeycomb that could be used without warping the cell structure walls, Plugging the holes was also considered during design, The honeycomb mate rial selected used a small cell pattern in order to be flexible enough to take the curvature. This meant that there would be fewer holes per cell and more holes blocked by the cell walls as the honeycomb was laid over the perforated plate. Fortunately, we we re able to use an adhesive that migrated up the cell walls and did not plug holes. The perforated plake was punched to a 6% open are a . When the honeycomb was then bonded to it, the open area was reduced to 5%.

The program gods are given in t c r c s of aircraft flyover noise parameters Experience has shown that when the e ~ g i n e is placed above the wing, the wing serves as a barrier. A barrier attenuation routine was included in the aircraft model to account for this affect.

As shown in figure 12, the wing creates a shadow zone that moves along with the aircraft. As only a small fraction of the noise i s refracted around the leading and trailing edges of the wing, the forward radiated f a n noise will not reach the ground as the shadow zone passes by.

ACOUSTIC TEST PHASE The test phase took most of the month of August to complete, The gods of the test program are shown in figure 13. They were to verify the noise predict;ons through comparison with measured data, determine the noise reduction of the mixer, and determine the effectiveness of the sound ttcat- ment panels, A test plan was prepared to accomplish these goals. The normal method of recording the noise emitted by the engine i s to record the sound pressure levels at nineteen locations on an arc 100 feet from the engine. With the microphones located cverv 10 degrees, a full set of data over an arc of 180 degrees can be obtained, Four power settings corresponding to the operating envelope of the engine were used. In addition to the far field microphenes, acoustic probes were placed on the engine to aid in identifying core and mixer components and the noise reduction of the sound treatment, A barrier was also used during part of the testing to aid in isolating the fan inlet and discharge component sound levels.

Three soparate engine configurations were used during the acoustic testing of the QCGAT engine. They arc a split flow exhaust nozzle configuration called the referee system, the hardwall nacelle in which the test nacelle, mixer, and hardwall fan inlet and discharge panels were used and the soft- wall nacelle in which the hardwall panels were replaced with the sound treatment panels. Each configuration was tested to record the effect on engine noise at four power settings. T h e QCGAT engine was mounted in a test frame and after a series of tc-sts in our test cells, it was moved to our free field test site, This site is located close to the plant in an area free of most noise intrusions and where testing will not intrude into the local connnunity.

The engine in tLe nacelle aad test frame were installed on a rotating test stand. This stand is capable of rotating a f u l l 360 degrees, The normal method of testing is to record the noise of the engine on an arc 100 feet from the engine by five microphones placed 10 degree^ apart as shown on figure 14. By rotating the engine and repeatiag the test p3iryts, a full 180 degrees of mise can be oblzined with some overlapping points, The microphones at the 170 and 180 degree points were in exhaust stream and were not used.

Positions 5 and b indicate the orientation of the engine inlet during the barrier test, One-half inch condenser microphones fitted ai th wind screens were placed on the ground as used and recommended by NASA, This allows for a simple 6 dB correction to be used when correcting the measured data to free field conditions for comparison with the predicted noise levels. The microphone array is shown in figure 15. The signal conditioning instrumentation are located in the acoustic data acquisition trailer where the data is recorded on magnetic tape for later analysis, A sample of the engine performance data is given inTable L At each test point, a complete set of engine performance data was recorded for use in predicting the engine static noise levels for comparison with the measured sound levels.

The ambient pressure, temperature, and relative humidity we re also recorded.

Fan noise is composed of tones that are easy to identify near the axes of the engine, krt they blend together at the 90 degree locations, The purpose of the barrier then was to isolate the fan inlet noise from the fan discharge noise by physically placing a barrier between them. This was accomplished at the free field test site with the barrier shown on figure 1 6 , The barrier was constructed o f a fixed Fartition 14 feet high by 20 feet long and a movable partition through which the engine inlet protruded. This effectively removed the fan discharge noise from the measurements of the fan inlet noise.

By rotating the engine 40 degrees between measurements, data was recorded over an arc of 80 degrees. The movable partition was then pulled out and the engine rotated 180 degrees so that the exhaust protruded through the barrier when it was moved back into position. The fan discharge noise was then recorded without fan inlet noise contributions. Both of these tests were run at the same four power setting with the hardwall a d the softwall nacelles installed on the engine.

The locations of the engine mounted probes are shown on figure 17. Half- inch condenser microphones were located upstream and downstream of the s c a d treatment to measure tbe noise reduction across the inlet sound treatment panels, Semi-infinite wave guide probes supplied by NACA were used to sample the acoustic pressure levels in the primary engine exhaust and at the mixer exhaust plane, These probes consisted of 114 inch con- denser microphone s in a sealed tube. A low volume flow of nitrogen at a pressure just above that in the duct provided a gas seal to prevent hot exhaust gas from entering the tube where it could damage the microphone.

These probes were designed to record the acoustic pressure levels at the indicated probe locations. The recorded data will also be used in corerence analysis to determine what part of the noise in the engine i s in fact radiGd out to the different far field measurement locations.

The split £low nozzle configuration with the semi -infinite wave guide probes installed in the primary exhaust nozzle are shown on f i g u ~ ~ 18.

This con- figuration was used to obtain baseline d ata for comparison with rnixer noise levels, DATA ANALYSIS The data recorded on magnetic tkpe %-as then analyzed. Reducing, organ- izing, and cataloging all this data was a time consuming task.

The audysis was straight forward. During the individual tcst runs, the engine perfor- mance was monitored and the relevant ambient and operating parameters recorded, Using these data and the appropristc cycle sheet data, we could predict the expected sound pressure levels.

These were then col.:pared point by ?aint, frequency by frequency, and angle by angle with the measured sound pressure levels. In this manner, we estimated the contribution of each component to the overall noise lcvels at each power setting. The predictions were then adjust-d to reflect this conlparison, and the correla- tion was run again. We also evalcated the insertion loss due to the sound treatment and determined the mixer noise reduction.

With the appropriate flight corrections and aircraft performance estimates, we were ready to estimate the flyover noise levels. The individual com- poaent contirbution to the overall noise levels were determined on a spectrum basis as shown on fi-wre 19. This plot consists of the one-third octave band sound pressure lcvels over a frequency range from 25 Hz to 20,000 Hz. The procedure for dcriving the flyover noise levels only The prc- considers the sound pressure levels from 50 Hz to 10,000 Hz.

dicted f a n noise contribution was overlaid, The calculations correctly locsted the bla-!e passing tone, its harrilonics and the broadband component.

The magnitude of the blade passing tone fundamental however was under- predicted. Next, the predicted jet noise component was added as shown on figure 20.

As was expected, the jet component does not contribute directly to t5e noise levels at the low power setting when the predicted core noise component is added to the noise spectrum as shown on fwre 21.

The predicted spectra matches the measured spectral shape. The agreement however is only fair in the mid-frequency region at the blade passing tone fundamental.

This same analy sis was carried out for the softwall and split flow configuration. The analysis was also carried out at each power setting.

The high power setting is shown on figure 22. Note that the agreement is only fair across the mid and high frequency regions of the spectrum, The Here low frequency part of the spectra appear to be in close agreement.

the jet noise corr.ponent is predicted to be the predominant source. Based upon this comparison and similar ones at other power seeings and con- figurations, u-e concluded that the jet noise prediction routine is adequate for the QCGAT program. Consequently, the predicted jet noise levels could be analytically removed from the measured data. The remaining noise levels would then be that composed of the core and fan components. Once the jet component had been removed thd sound power levels attributed to the core were then compared with the predicted core sound power levels as shown in figure 23. Also plotted a r e the sound power levels derived from ihe acoustic probes located in the primary exhaust. The probe data are shoWz more a s a confirmation of the slope rather thanthe sound power levels correctly calculated. These data indicate that the core noise model This underprediction underpredicts the core noise level by roughly 3 dB.

appears to be indepelldent of the power setting of the engine. A simple 3 dB correction factor was the ref0 re applied to the core noise prediction procedures. After making this refinement to the core noise model, the predicted-to-measured correlation was then rerun. Figure 24 shows that comparison. The spectral agreement between the measured and predicted data is good over the frequency range of interest. Note that the sound levels in the band containing the tone at blade passing are also in good agreement. This indicates that the core noise contributes across the spectrum. The dominance of the core noiso can be seen in figure 25. The noise levels in the discharge quadrant are dominated Ey the core component The core to the extent that the fan component is almost entirely masked.

noise component i s present in the forward quadrant. The reduction in the fan noise levels by the sound treatment was hard to discern for this reason.

When the core noise component is removed from the one-third octave band containing the blade passing tone, and the resulting blade passing tone is plotted against the ar.gle from the inlet, a s shown on figure 26, a fan tone directivity plot is formed. The predicted sound pressure levels at the peak angles are also shown for the inlet and discharge quadrants. The expected results with the barrier in place come from the prediction procedures, Only when the barrier is in place will the measured data approach these lines which it does as can be seen by the dotted lines. This plot shows h~ the fan noise contributes to the foxward and aft radiated engine noise levels, If an observer were to move past this plot as indicated, the noise levels experienced would first rise and then fall off as the observer moved past. Once past the engine, the noise levels would then rise again as the discharge fan noise reached the observer. This is roughly how the static data was converted to observed flight sound levels. At the high power setting (figure 27) the core noise obscures the aft fan tone from the analysis.

A small adjustment was made to the fan noise model from which these data were derived. This adjustment had to do with the effect of relative tip design mach number. With this adjustment, we concluded from the agreement shown here and on the previous figure that the fan noise model is accurately com- puting the fan noise levels. The sharp dip at the 60 degree point is due to the fact that the data from 0 to 40 degrees were recorded at slightly different power settings than the data from 50 to 90 degrees, The predicted data shows this same dip, We feel this is an artifact of the data acquisition process and i s not a characteristic of the fan noise, The individual com- ponent contributions appear to be adequakly predicted once the noted corrections have been made. Figure 28 shows a final comparison of the measured and predicted overall sound power levels. This plot was generated to verify the accuracy of the prediction techniques for the static case before proceeding to the flyover analysis. The agreement shown here indicates to us that the updated noise prediction model accurately reflects the static noise emissions of the QCGAT engine, As noted earlier, it was difficult to discern the noise reduction of the sound treatment panels from the far field data. Figure 29 shows the ose-third octave band sound pressure levels at the upstream and down- stream micropho.?e locations in the inlet. Here the acoustic energy is propagating against the air flow in the inlet duct. The upstream microphone then recorded the inlet noise after it had passed through the treated part of the inlet duct. Figure 30 shows that the expected insertion loss and the insertion loss derived from the test data, These are the values that will be used inthe flyover noise estimates. Figure 31 shows that the expected and estimated insertion loss for the fan discharge duct sound treatment panels. The discharge panels had no provision for microphones and were unable to discern a noise reduction from the far field data due to the presence of the core noise. We have assumed that the treatment is The estimated values for the discharge sound treat- functioning properly.

ment panels are shown here.

The jet noise Icvels were predicted to be low due to the use of a high by- pass ratio fan. Figure 32 shows the difference between the noise spectra of such an engine fitted with the split flow nozzle configuration and with the mixer nacelle configuration. The shaded area represents the static noise reduction of the mixer. Above 250 hertz, the core noise source starts to mask the jet noise and above 1000 hertz, the fan i s dominant.

When flight effects are *dded, both the mixed and split flow jet components will drop leaving the mixed flow jet noise levels below the core noise levels.

The split flow noise levels would drop and be roughly equal in magnitude to static jet noise levels.

The procedures employed (figure 33) in the QCGAT program to assess the nois* emissions of a QCGAT powered aircraft are the Federal Aviation Administration's certification procedures for turbojet powered aircraft (Reference 6 ) . This is a very rigorous method. Basically, the FAA requirements call for measuring the aircraft noise every hslf second as the aircraft flies over the measurement point. For this analysis, pre- dicted data was substituted for the actual ne asurements. Thq demon- stration engine performance a d the Beech aircra£t design were us- ' to compute the i~dividual test point performances. These data were ,- a entered into the prediction procedures. The appropriate flight and wing shielding effects were then applied to the individual component noise predictions. The aircraft noise signature was then derived by combining these into a table of aircraft noise. Then by analytically moving the air- craft noise table past the measurement point, the time history of the flyover could be constructed for each half-second interval. These sound levzls were then used to compute the tone corrected perceived noise levels for the flyover event. The maximum tone-corrected perceived noise levels was then feud along with the time the aircraft noise is within 10 PNdB of the maximum. From these data, the effective perceived noise level is calculated.

Figure 34 shows the tone corrected perceived noise levels versus time for the approach flyaver. The maximum tone-corrected perceived noise level, labeled PNLTM occurs after the aircraft ha.s passed directly overhead.

The time the PNLT was within 10 PNdB of the value is 8.5 seconds. This plot also shows that the fan inlet ard discharge noise are heard at separate The valley between the peaks is caused by the lower sound le~.rels times.

gene rated at the sideline positions. Wing snieldinb, the shaded portion, acts to cut the inlet peak off early and makes this valley deeper. The core noise component i s heard after the aircraft is past as most of the core noise is in the aft quadrant of the engine. Because of the duration correc- tion, the fan component noise levels are higher and contribute more to the effective perceived noise levels. Figure 35 is the same type of plot show- int the takeoff flyover tone-corrected perceived noise level time history.

Here the time the noise is within 10 PNdB of the max is much longer. At F o r the take- the approach condition, the altitude at flyover is 370 feet.

o f f condition, it is 2600 feet. Consequently, the time will be considerably longer. The maximum tone-corrected perceived noise level also occurs much later a s the sound requires longer to I aach the observer and because the dominant noise sources a r e the core and jet. These components radiate most of their acoustic energy in t). rear quadrants and, a s such, a r e not heard until the aircraft is past tht? observer. Also shown here a r e the higher noise levels of a split flow nozzle configured aircraft.

Here the jet component contributes more the aircraft noise levels both in magnitude and duration, The dura..ion i s increased because the jet noise peaks farther aft than does the core noise. This means that the peak noise occurs later in the flyover. Thus, the addition of the mixer not only reduces the aircraft flyover noise levels, the aircraft noise does not linger as long.

CONCLUSION For an aircraft powered by two Avco Lycoming QCGAT engines installed in a nacelle that includes a mixer and fan inlet and discharge sound treat- ment panels and mounted over the wings, the effective perceived noise levels for the takeoff, sideline, and approach conditions will be 68.4, 7.06, and 77.3 EPNdB, respectively, These noise levels shown in figure 36, are below the limits set by the QCGAT program goals. In the analysis, the effect of several alternative engine configurations on the aircraft noise was assessed. For example, removal of the sound treat- ment panels would add 2 EPNdB to the approach noise levels and still be below the QCGAT goals. The other positions would not he affected.

The noise levels shown here are for the engine that was tested kid

delivered to NASA . When the iterations are completed for this engine

design, the increased thrust of the engine will mean that the aircraft will achieve an altitude of 3600 feet over the takeoff point versus the present 2600 feet. This upill result in a 3 EPNdB reduction in the takeoff noise levels and a 1 EPNdB reduction in the sideline noise levels. In this case, the split flour exhaust nozzle configuration would be within 1 EPNdB of the QCGAT goals, Figure 37 shows the Avco Lycoming QCGAT engine effective perceived noise levels plotted against thAc Federal Aviation Admiriistration's Stage III noise standards and the high technology that used by NASA for the QCGAT program goals.

This demonstrates that the technology that has worked for the large engine can be transferred to the general aviation size engine. Consequently, turbofan engine noise emissions should not be a constraint to the growth of the general aviation market.

In summary, (see figure 38) large turbofan noise control technology was successfully applied to a general aviation size engine.

The stringent program goals set by NASA forced a design that required the use of a design that required the use of a quiet fan and integration of the nacelle and aircraft in the engine design. This demonstrates that the QCGAT program goals can be met with the latest noise control techniques with- out incurring a pe rforrnance penalty.

REFERENCES 1. Stone, J. R. : Interim Prediction Method for Jet Noise. NASA TM X- 71618, 1974.

2. H u f f , R. G. ; Clark, B , 3. ; and Dorsch, R. G. : Interim Prediction NASA TM X-71627, Method for Low Frequency Core Engine Noise.

19 7 4.

3. Heidman, M. F. : Interim Prediction Method for Fan and Compressor Source Noise, NASA TM X-71763, 1975.

4. Dunn, D. G. : Aircraft Noise Source and Contour Estimation. NASA CR- 114649, 1973.

5. Gillian, R. E . ; e t aL : ANOPP Users Manual. July 1978.

6. Federal Aviation Administration Noise STandards, Title 14, Code of Federal Regulation, Chapter I, P a r t 36.

Table I TYPICAL ENGINE PARAMETERS RECORDED DURING NOISE T ESTS TEST CONDITIONS LOW POWER HIGH POWER ENGINE PARAMETER SETTING SETTING 5376 9584 Fan Rotor Speed. rpm Fan Blade Pass~ng Frequency, Hz 2150 3673 Fan Relative Tip Mach Number ,509 .89 14 a(32.6) 26.3(57.9) Fan Airflow. k g / ~ e c ( l b / ~ e ~ ) 6.7(12) 15(27) Fan Temperature Rise. "C("F) Combustor Airflow. kg/sec(lb/sec) 1.33(2.95) 2.37(5.22) Combustor Temperature Rise. " C(" F) 630(1135) 878( 1580) Jet Exit Velocity. kg/sec(ft/sec) 106(350) 194(636) Jet Exit Temperature. "K(OR) 358(645) 389(697)

QCGAT POWERED AIRCRAFT NOlSE GOALS

APPROACH FLYOVER MEASUREMENT POINT 3.5 - N M - / / TAKEOFF FLYOVER .25 NM MEASUREMENT

-+d-++- POINT

MEASUREMENT POINTS TO DETERMINE MAXIMUM TAKEOFF SIDELINE NOISE QCGAT ACOUSTIC PERFORMANCE EPNL GOAL PREDICTED ENGINE CONDITION EPNdB EPNL, EPNdB Takeoff Flyover 69.4 64.8 Takeoff Sideline 78.4 71.7 4pproach Flyover 83.4 73.8 Figure 1

AVCO LYCOMING AIRCRAFT ENGINE

NOlSE PREDICTION PROCEDURES STATIC ENGINE NOlSE I ENGINE CYCLE DATA I MEASUREMENTS t STATIC NOlSE PREDICTION FOR:

I

ATMOSPHERIC COMPARISONS MODEL FOR TO IMPROVE NOlSE PROPAGATION JET

I

I I

I

NOISE LEVELS I

AIRCRAFT 1

FLIGHT I

- FLYOVER h IISE PREDICTIONS

CONFIGURATION EFFECTS] AND PERFORMANCE I

NOISE LEVELS 1

Figure 2

FAN NOISE REDUCTION TECHPIQUES

PARAMETER TECHNIQUE Blade Loading Low Pressure Ratio Subsonic Blade Tip Speed Greater Than 2 Blade Widths Blade to Vane Spacing Greater Than 2 Vane to Blade Ratio lnlet No lnlet Guide Vane Low lnlet Turbulence Figure 3

FORWARD FLIGHT EFFECTS

Reduce lnlet Turbulence Reduce Jet Noise Dynamic Amplification Doppler Shift SMOOTH INFLOW REDUCED FLIGHT

-----------

7 HIGH

RELATIVE VELOCITY TURBULENT

INFLOW I

INLET

EFFECT OF NOZZLE CONFIGUR..' "ION

ON JET NOISE

WITH HIGH BYPASS FAN TWO TURBULEi'l I INTERF9CES LOWER EXlT VELOCITY

CORE NOISE MODEL

(GOOD AGREEMENT FOR TURBOSHAFT ENGINES) COMBUSTOR INLET . - . . - - . , ; J ~ L NOZZLE COMBUSTOR EXlT Noise a Function of: Mass Flow Temperature Rise Pressure Drop Figure 6

QCGAT FLIGHT NACELLE

with

SOUND TREATED PANELS INSTALLED

PANU Figure 7

LOCKHEED DESIGN RECOMMENDATIONS

FOR SOUND TREATMENT PANELS OPEN THIUWESS LENGTH AREA

-

Fan Inlet 16mm(0.63 in.) 330mm(13 in.) 5% 16mm(0.63 in.) 460mm(18 in.) 5% Fan Discharge SOLID BACKING P A W L HONEYCOMB FACE SHEET

PREDICT ED FAN INLET ATENUATION

r TAKEOFF FREQUENCY. Hz Figure 9

PREDICTED FAN DISCHARGE ATTENUATION

VEWICATION OF PREDICTION TECHNIQUES

FREQUENCY - HZ Figure 19

METHOD OF ANALYSIS

PREDICT STATIC NOlSE EMISSIONS AT TEST CONDITIONS COMPARE WITH MEASURED DATA DETERMINT COMPONENT CONTRIBUTIONS REFINE CALCULATIONS TO REFLECT TEST EXPERIENCE DETERMINE SOUND TREATMENT NOISE REDUCTION DETERMINE MIXER NOISE REDUCTION APPLY FLIGHT CORRECTIONS PREDICT FLYOVER NOISE L EVELS Figure 20

VERlFlCATlON OF PREDlCTlON TECHNIQUES

I LOW POWER SC b-;1NC- - H A R D W A U CONFIGURATION 1

FREQUENCY - HZ Figure 21

VERIFICATION OF PREDICTION TECHNIQUES

I LOW POWER SETTING - HARDWALL CONFIGURATION 1

MEASURED PREDICTED FREQUENCY - HZ Figure 22

CORE NOISE PREDICT ION

/ / 100% 'CORRELATION

//

/ 0 1 / /

y ,/

/ .) TAILPIPE UIC

/ 0 /

+ FARFIELO DATA

PREDICTED' 'Sound Power Level, dB RE 10-12 Watts Figure 23

RESULTS WlTH UPDATED CORE NOISE MOOEL

HIGH POWER SETTING - HARDWAU CONflGURATlON

r V) I- PREDICTED t - a $ 1 2 0 - $ U.

a : m - Q 110 i W > W

E 100-

rJ ?

rJ r J I FAN

z -1

U) I I 1 I 1 1 1 I I 31.5 63 125 250 500 1030 ?Oo0q00C B O O O 1 ~

FREQUENCY - HZ

Figure 24

TYPICAL INLET AND EXHAUST NOISE SPECTRA

Fan Tone Masked By Core Noise EXHAUST QUADRANT SPECTRUM 15.

70.

rJ 1

65. -

I

-

L, 60.

I iNLFT QUADRANT

55. - SPECTRUM

50. -

1 1 I I 1 I 4 I I I I 45.

31.5 63 125 250 500 10oo-q000-16000 FREQUENCY - HZ Figure 25

FAN TONE DlRECTlVlTlES

WITHOUT BARRIER ---- EXPECTED RESULTS FROM BARRIER TEST I .. . - . . - - ..

BARRIER TEST DlRECTlVlTlES 1

- PREDICTED PEAK SOUND LEVELS

ANGLE FROM INLET, DEGREES Figure 2 6

FAN TONE DlRECTlVlTlES

EXPECTED RESULTS FROM BARRIER TEST

. .. ... BARRIER TEST DlRECTIVlTlES

- PREDICTED PEAK SOUND LEVELS

ANGLE FROM INLET, DEGREES Figure Z l

COMPARISON BETWEEN MEASURED -

AND PREDICTED ENGINE NOISE

cn

14Or I HARDWALL NACELLE

115 120 125 130 135 140 PREDICTED SOUND POWER LEVEL, dB RE 10-'2 WATTS Figure 28

INLET TREATMENT NOISE REDUCTION

FAN BLADE PASSING 150. r TONE J V ) UI -J > a W U - 1 Y) w e a n $0 a 2 no a?

Z W 3 C r :

2%

I I I I 1 1 1 105.1 250 1000 *O0O 4000 16000 FREQUENCY - Hz Figure 29

FAN INLET SOUND TREATMENT

Treatment Estimated to Meet Design Specifications Analysis Limited By Low Fan Sound Levels MEASURED APPROACH TAKEOFF r

'REOUENCY - Hz

Figure 3 0

FAN 7ISCHARGE SOUND TREATMENT

Treatment Estimated to Meet Design Specification Analysis Limited By Low Fan Sound Levels

- GOAL

ESTIMATED APPROACH TAKEOFF FREQUENCY - Hz Figure 31

NOISE REDUCTION DUE TO MIXER

Greater Than 5 dB Coincides with Predicted Results FAN COMPONE 31.5 63 125 250 'O0 1000 2000 4Wl *OoO 16000

FREQUENCY - Hz

Figure 32

QCGAT FLYOVER NOISE

CALCULATION PROCEDURE PERFORMANCE ENGINE PERFORMANCE NOISE PREDICITON

I FOR FLYO' .7 CONCllTlON W PROCEDURES I

- FLIGHT EFrECTS 10.5 SECOND TlME HISTORIES I

I TONE CORRECTED PERCEIVED

NOISE LEVELS IPNLTI AT EACH 0.5 SECOED t MAXIMUM PNLT AND LENGTH OF TIheE BETWEEN 10 PNdB DOWN PCINTS EFFECTIVE PERCEIVED

NOISE LEVELS (EPNLI I

I

Figure 33

COMPONEhT CONTRIBUTION

TO APPROACH NOISE

MAXIMUM PNLT

PNLT -10 dB '), P L T - l o d B

4 8.5 SEC

L EPNL. EPNdB =

72.2 - CORE

I 73.9 - FAN

I I I I

1 I I I ' --

-10 -5 0 5 10 15 TlME FROM FLYOVER. SECONDS

QCGAT vs FAR PART 36 LIMITS

SIDELINE TAKEOFF t i ; m !

z Q W i APPROACH FAR-36 LIMIT QCGAT GOAL AVCO LYCOMING I I I 1 1.3 4.5 45 450 Kg GROSS TAKEOFF WEIGHT x loo0 Figure 37

SUMMARY

Successful Application of Large Turbofan Noise Control Technology in a General Aviation Size Turbofan Engine All QCGAT Noise Goals Demonstrated (Takeoff. Sideline and Approach) Used Available Noise Control Techniques to Meet Stringent Noise Goals Without a Performance Penalty Noise N?ed Not Be a Constraint to General Aviation Growth Figure 3 3 SUMMARY OF NASA QCGAT PROGRAM Gilbert K. Sievers National Aeronautics atid Space Administration Lewis Research Center As was s t a t e d i n t h e program overview, t h e QCGAT program o b j e c t i v e s were t o d e l a n s t r a t e t h a t t h e a p p l i c a t i o n o f l a r g e t u r b o f a n e n g i n e technology t o soall g e n e r a l a v i a t i o n t u r b o f a n e n g i n e s c a n a c h i e v e low n o i s e , low emissions, and ac- c e p t a b l e f u e l consumption.

NOISE GOALS F i g u r e 1 shows t h e t a k e o f f n o i s e g o a l . T h i s is t h e same f i g u r e t h a t was shown i n t h e overview, e x c e p t t h a t t h e r e s u l t s p r e d i c t e d f r o a ground static t e s t i n g are shown. Avco i s q u i e t e r t h a n t h e g o a l w h i l e AiResearch is meting t h e goal. For t h e s i d e l i n e and approach c o n d i t i o n s ( f i g s . 2 and 3 ) , b o t h en- g i n e s are q u i e t e r t h a n t h e g o a l s . By meeting o r b e t t e r i n g t h e s e s t r i n g e n t N A S A n o i s e g o a l s , we b e l i e v e t h a t WGAT h a s demonstrated t h a t n o i s e need n o t be a major c o n s t r a i n t on t h e f u t u r e growth of t h e g e n e r a l a v i a t i o n t u r b o f a n f l e e t .

EMISSION GOALS The measured e m i s s i o n r e s u l t s f o r t h e AiBesearch QCGAT e n g i n e a r e shown i n f i g u r e 4. For each p o l l u t a n t , t h e b a r a t t h e l e f t shows t h e p u b l i s h e d l e v e l o f The b a r i n t h e c e n t e r shows t h e QCGAT g o a l a n d t h e p r o d u c t i o n TFE-731-2 engine.

t h e b a r a t t h e r i g h t , t h e measured QCGAT r e s u l t s . A s c a n be seen, t h e c a r b o n monoxide e m i s s i o n s were lower t h a n t h e g o a l ; t n e g o a l f o r t h e unburned h y d r o c a r boas was m e t ; and, w h i l e t h e N O , g o a l was n o t m e t , QCGAT NOx e m i s s i o n s are Also, e n g i n e smoke u a s n o t visi- lower t h a n t h o s e of t h e p r o d u c t i o n TFE-731-2.

ble.

The measured emission r e s u l t s f o r t h e AVOO QCGAT e n g i n e a r e shown i n f i g u r e No comparisons w i t b p r o d u ~ ~ i o n e n g i n e s a r e made because t h e r are no c o m p a r 5.

However, i t c a n be s e e n t h a t measured e m i s s i o n s f o r a b l e p r o d u c t i o n engines.

carbon monoxide and unburned hydrocarbons a r e lower t h a n t h e s t r i n g e n t g o a l s and Again, t h i s e n g i n e d o e s n o t produce v i s i b l e t h a t N O , is r i g h t a t t h e goal.

s m k e .

PERFOWNCE GOALS A comparison of t h e AiResearch QCGAT e n g i n e measured performance v i t h t h e performance g o a l s is g i v e n i n t a b l e I. The AiResearch QCGAT engine met its The s e a leuel t h r u s t g o a l s b o t h a t s e a l e v e l t a k e o f f and a t d e s i g n c r u i s e .

t a k e o f f SFC i s about 2 p e r c e n t h i g h e r t h a n t h e g o a l , b u t a t d e s i g n c r u i s e , v h e r e it r e a l l y c o u n t s , t h e SPC i s l o v e r t h a n t h e g o a l and i s approximately 9 o r 10 p e r c e n t l o v e r t h a n t h e c u r r e n t p r o d u c t i o n TPE-731-3.

A comparison of t h e A V C O QCGAT e n g i n e areasured performance w i t h t h e g o a l s is g i v e n i n t a b l e 11. The AVCO QCCAT e n g i n e d i d n o t meet t h e sea l e v e l t a k e o f f o r d e s i g n c r u i s e g o a l s f o r e i t h e r t h r u s t o r SFC. However, t h e measured numbers a r e q u i t e r e s p e c t a b l e . ,It must be remembered t h a t t h i s e n g i n e d i d n o t e v o l v e It is be- from a mature productioi?engine as d i d t h e AiResearch QWAT engine.

l i e v e d t h a t t h e AVU) QCCAT e n g i n e is o n e i t e r a t i o n away from meeting t h e goals.

T h : major g o a l s f o r t h e QCGAT p r o j e c t were m e t and t h e p r o j e c t was com- W e c o n s i d e r t h i s t o have been a p l e t e d on s c h e d u l e and w i t h i n t h e NASA budget.

very s u c c e s s f u l NASA j o i n t e f f o r t w i t h i n d u s t r y .

TABLE 1. - AIRESEARLH WGAT PERFOHMNCE

{ S t a n d a r d d a y ; i n s t a l l e d 1 Goa 1 U e a s u t e d

S e a l e v e l / T h ~ s t , N(1b) 17312 (3892)

17312 (38921 takeotf S F C , kg/hr-N ( lb/hr-lb) 0.0431 (0.423) 0.0440 (0.4311

I I

Design crulse ? h r u s t , N( lb) 4017 (903) 4017 (903) h = 0.8 12 200 km ( 6 0 000 f t ) SFC, kg/hr-N ( I b / h r l b ) 0.0759 (0.746) 0.0756 (0.741)

-

TABLE 11. - AVCU QCGAT PERFORMNCE

( S t a n d a m day; i n s t a l led]

i - < l SFC, kg/hr-N ( lbt hr-lb) 0.0370 (0.363) 0.040 (0.392)

a ;

Design c r u i s e I T h r u s t , N(lb; 2157 (485) 1850 (41b)

n 0.6

7600 km ( 2 5 000 f t ) i S F C , kg/hr-N i lb/hr-1b) 0.0640 ( 0 . 6 2 8 ) 0.0737 to. 723)

i

TAKEOFF NOISE LEVELS

EPNdS

60 I I 1

1 10 l m loahtld

TAKEWF GROSS WE1 GHT. kg

I I I I

3 10 1 0 mxld

TAKEWF GROSS WEIGHT, Ib Figure 1

SIDELINE NOISE LEVELS

1969 FAR-36 REQUI REMENT 1977 RULE EPNdB 80 AIRESEARCH

60 a

1 10 100 1000~ 103 TAKEOFF GROSS WE! GHT, kg

I-

3 10 100 1000~ 103 TAKEOFF GROSS WEIGHT, Ib Figure 2

APPROACH NOISE LEVELS

A 3 0 0 0 7 1 1 0 1969 FAR-36 REQUl REhWT 1 9 7 7 RULE QCGAT GOAL LEARJET 36 0 0 AIRESEARCH AVCO

I I I

1 1 0 1 0 0 loooxlo3

TAKEOFF GROSS WE1 GHT, kg

1 I I I

1 0 0 l o o o x l d

3 1 0 TAKEOFF GROSS WEIGHT, Ib Figure 3

AIRESEARCH QCGAT EMISSIONS

T R 731-2

14 0 QCGAT GOAL

QCGAT ENGINE (MEASURED)

4 z 10

8 " L TFE 731-2 PRODUCTION ENGINE VALUES1

3 S F : NASA TM 79009. OCT 1978 ~ 0 . 2 c (

- Q . 1

= 2

0 n " CO UHC OX C $ - 7 0 - 2 7 0 ] Figure 4

AVCo QCGAT EMISSIONS

0 OCGAT GOAL QCGAT ENGINE MEASURED) Figure 5 NEW OPPORTUNITIES FOR FUTURE, SMALL, GENERAL- AVIATION TURBINE ENGINES (GATE) William C . Strack National Aeronautics and Space Administration Lewis Research Center The r e s u l t s of f o u r independent contracted s t u d i e s t o e x p l o r e t h e opportun- ities f o r f u t u r e small t u r b i n e engines a r e s u m a r i z e d i n a composite overview.

Candidate advanced technologies a r e screened, v a r i o u s c y c l e s and s t a g i n g ar- rangements are parametrically evaluated, and optimum conceptual engines a r e i d e n t i f i e d f o r a range of 300 t o 600 hp applications.

Engine improvements of 20 percent i n SFC and 40 percent i n engine c o s t were f o r e c a s t using high-risk tech- nologies t h a t could be t e c h n i c a l l y demonstrated by 1988. The ensuing economic b e n e f i t s a r e i n t h e neighborhood of 20 t o 30 percent f o r twin-e?.gine a i r c r a f t c u r r e n t l y powered by p i s t o n engines.

INTRODUCTION The preceding p o r t i o n of t h i s conference was devoted e n t i r e l y t o business jet e x t e r n a l n o i s e and p o l l u t i o n , primarily, and f u e l economy, secondarily.

This turbofan-powered segment of general a v i a t i o n r e p r e s e n t s about one-fourth of U.S. engine n e t f a c t o r y b i l l i n g s even though only 2 percent of t h e general- The remainder of t h e a v i a t i o n a i r c r a f t engines s o l d a r e of t h i s t y p e ( f i g . 1).

conference w i l l address t h e concerns of t h e o t h e r 98 percent. A t t h e lower c o s t end a r e t h e single-engine a i r p l a n e s powered by 100- t o 300-hp p i s t o n engines.

About 23 000 of t h e s e were produced l a s t year. The twin-piston powered a i r planes u t i l i z e 200- t o 400-hp engines of which about 8000 a r e produced annually.

The turboprop a i r c r a f t a r e mostly twins t h a t r e q u i r e 500- t o 1000-hp engines.

The U . S . produced about 750 of t h e s e l a s t year, and Canada produced approxi- mately t h e same number. C o l l e c t i v e l y , t h e s e t h r e e c a t e g o r i e s represent 3/4 of t h e general-aviation engine net b i l l i n g s . Last year t h e t o t a l general-aviation engine b i l l i n g s were about 70 percent a s l a r g e a s those f o r t h e l a r g e connnercial t r a n s p o r t turbofans.

The p r i n c i p a l problems facing t h e s e t h r e e c a t e g o r i e s a r e n o t s o much envi- ronmental a s they a r e economic, f u e l , and s a f e t y r e l a t e d ( f i g . 2). Perhaps of g r e a t e s t concern i s t h e c o s t and a v a i l a b i l i t y of a v i a t i o n fuels. Continued s t e e p p r i c e hikes and t h e v u l n e r a b i l i t y of a v i a t i o n g a s t o severe production cutbacks, o r o u t r i g h t elimination, propels o u r q u e s t f o r t r u e m u l t i f u e l p o w e r plants. The s a f e t y of t h i s c l a s s of a i r c r a f t continues t o be questioned - with t h e s p o t l i g h t a l t e r n a t i n g between a i r p l a n e and automotive s a f e t y record campar- isons and t h e c o n t r o v e r s i a l one-engine-out twin problem. Engine dependability i s e s p e c i a l l y important i n t h e s e s m a l l e r a i r c r a f t . P a s s e n g e r comfort l e v e l s a r e f a r l e s s t h a n t h o s e o f l a r g e r t u r b o f a n a i r c r a f t , and t h e powerplant is a major c a u s e o f t h e discomfort. O t h e r c o n c e r n s i n v o l v e p r o p u l s i o n - r e l a t e d a c q u i s i t i o n and maintenance c o s t s - e s p e c i a l l y f o r t u r b o p r o p e n g i n e s ( f i g . 3). NASA'S i n - volvement i n a d d r e s s i n g t h e s e c o n c e r n s f o r t h e s m a l l e r g e n e r a l - a v i a t i o n p o w e r p l a n t s is r e c e n t . Two y e a r s ago w e recognized t h a t t h e t i n y amount of R&T e f - f o r t devoted t o small g e n e r a l - a v i a t i o n t u r b i n e s was n o t i n p r o p o r t i o n t o t h e i r a c t u a l importance. A s t h e f i r s t s t e p i n r e c t i f y i n g t h a t s i t u a t i o n , we i n i t i a t e d a series o f a n a l y t i c s t u d i e s - known as t h e GATE s t u d i r s - t o e x p l o r e s m a l l t a r b i n e technology o p p o r t u n i t i e s . The q u e s t i o n was: I f w e h y p o t h e s i z e a brand new, s m a l l t u r b i n e e n g i n e t h a t i n c o r p o r a t e s , s a y , 1088 l e v e l technology, what s i z e should i t be; how s h o u l d i t b e c o n f i g u r e d ; and what b e n e f i t s would i t be- stow upon us? The purpose, t h e n , became o n e o f p r o v i d i n g i n f o r m a t i o n t o a s s i s t u s i n p l a n n i n g f u t u r e r e s e a r c h . W e wanted t o emphasize t e c h n o l o g i e s t h a t had h i g h payoff and h i g h r i s k , b u t which c o u l d be r e a d y f o r p r o d u c t i o n development by 1988 ( g i v e n s u f f i c i e n t f u n d i n g ) and f o r mass p r o d u c t i o n by t h e e a r l y 1990's.

These e n s i n e s c o u l d be as much as 1000 hp, b u t no more, b u t we emphasized sizes below 600 hp s i n c e w e p e r c e i v e d t h i s s i z e c l a s s t o be p o t e n t i a l l y t h e most rewarding - and c h a l l e n g i n g . W e a l s o emphasized a i r c r a f t c o s t o f ownership as a c r i t e r i o n of m e r i t d u r i n g t h e c o n c e p t u a l d e s i g n process.

The f i r s t t a s k o f t h e s e s t u d i e s was a 1988 market f o r e c a s t t h a t c o n s i d e r e d T h i s f o r e c a s t d e t e n n i n e d e n g i n e a l l t y p e s of s m a l l a i r p l a n e s and h e l i c o p t e r s .

power s i z e s and o t h e r r e q u i r e m e n t s of i n t e r e s t . Most o f t h e s t u d y e f f o r t , how- e v e r , was devoted t o broad-scope p a r a m e t r i c a n a l y s e s wherein v a r i o u s c y c l e s , s t a g i n g arrangements, and t e c h n o l o g i e s were s u b j e c t e d t o t r a d e - o f f and s c r e e n i n g e v a l u a t i o n s t o determine optimum e n g i n e c o n f i g u r a t i o n s f o r e a c h i m p o r t a n t m i s - s i o n i d e n t i f i e d i n t h e market f o r e c a s t . Then, a n t i c i p a t i n g t h a t t h e marketplace c o u l d n o t a f f o r d d i f f e r e n t optimum e n g i n e s f o r e a c h a p p l i c a t i o n , a n e v a l u a t i o n was made o f a s i n g l e common c o r e t o be used i n a f a m i l y o f engines. And f i n a l - l y , t h e r e q u i r e d RbT program was d e f i n e d .

W e a t N A S A d i d some of t h e s e a s s e s s m e n t s o u r s e l v e s , b u t most were done un- d e r c o n t r a c t t o t h e s e f o u r companies working independently: G a r r e t t 1 AiResearch, D e t r o i t D i e s e l A l l i s o n , Teledyne CAE, and Williams Research. Within a g e n e r a l framework, we p e r m i t t e d t h e company teams t h e freednm t o p u r s u e d i r e c - t i o n s and o p p o r t u n i t i e s t h a t t h e y ( r a t h e r t h a n we) p e r c e i v e d a s most a t t r a c t i v e .

T h i s freedom sometimes l e d t o u n i f o r m i t y , f o r example, a l l companies e x p r e s s e d a s t r o n g p r e f e r e n c e f o r t u r b o p r o p s i n s t e a d of t u r b o f a n s o r t u r b o s h a f t s , and some- t i m e s t o i n t e r e s t i n g d i v e r s i t y , f o r example, t h e e n g i n e c o n f i g u r a t i o n s and t e c h - n o l o g i e s v a r i e d c o n s i d e r a b l y a s d i d e n g i n e c o s t e s t i m a t e s .

The e n g i n e c o s t e s t i m a t e s were e s p e c i a l l y i n t r i g u i n g i n t h e GATE s t u d i e s because o f t h e obvious o p p o r t u n i t y t o t u r b i n i z e a p o r t i o n of t h e piston-powered market. The t u r b i n e e n g i n e i s p r e t t y rmch a c c e p t e d a s t h e most d e s i r a b l e t y p e o f powerplant because of i t s many v i r t u r e s - it h a s very low v i b r a t i o n l e v e l s , high r e l i a b i l i t y , m u l t i f u e l c a p a b i l i t y , a b e t t e r s a f e t y r e c o r d , low weight, few- e r emissions, l e s s maintenance, and s m a l l e r i n s t a l l a t i o n l o s s e s . D e s p i t e t h e s e advantages, t h e u s e of t u r b i n e e n g i n e s h a s been blocked a t about t h e 500 hp l e v e l because of i t s h i g h e r f u e l consumption and, e s p e c i a l l y , i t s 3:l p r i c e pre- mium ( f i g . 4). The c h a l l e n g e , of c o u r s e , is t o overcome t h e c o s t and f u e l bar- r i e r s without s a c r i f i c i n g a l l t h e s u p e r i o r q u a l i t i e s . The t r o u b l e i s t h a t c u r r e n t technology does n o t a l l o w t h i s .

i f we a t t e m p t t o lower c o s t s i g n i f i c a n t l y , t h e e f f i c i e n c y s u f f e r s t o o much ( f i g . 5 ) . But, i f advanced technology could move t h e c o s t - e f f i c i e n c y band down f a r enough, we c o u l d c e r t a i n l y d e s i g n a c o s t - e f f e c t i v e s m a l l t u r b i n e o r , i f we choose, a b e t t e r performer w i t h o u t c o s t reduction. On t h e o t h e r hand, i f ad- vanced technology could n o t lower t h e band s u f f i c i e n t l y , t h e n o n l y t h e high per- formance o p t i o n i s open. I n t h e end, t h r e e GATE s t u d y team pursued t h e low-cost t u r b i n e v e r s u s p i s t o n theme, and t h e f o u r t h pursued a high-performance, advanced t u r b i n e v e r s u s c u r r e n t t u r b i n e theme.

r e s u l t s of t h e f o u r c o n t r a c t e d s t u d i e s a r e p r e s e n t e d h e r e i n by s e l e c - The t i v e examples t h a t i l l u s t r a t e t h e main p o i n t s i n a r e p r e s e n t a t i v e fashion. The d e t a i l e d r e s u l t s a r e documented i n r e f e r e n c e s 1 t o 4.

CYCLES AND CONFIGURATIONS I n f i g u r e 6, d e s i g n t u r b i n e - i n l e t t e m p e r a t u r e s of 18000, 19000, and 2200° F a r e compared i n terms of a i r p l a n e t o t a l c o s t o f ownership, f u e l con- sumption, o p e r a t i n g c o s t , a c q u i s i t i o n c o s t s , and e n g i n e c o s t . For a l l of t h e s e c r i t e r i a , t h e optimum t e n p e r a t u r e l e v e l i s 2 2 0 F, o r about 400° F above c u r r e n t small-engine l e v e l s . Although i t a p p e a r s t h a t t e m p e r a t u r e s i n e x c e s s of 2200° F would be even b e t t e r , 2200° F was judged t o be t h e h i g h e s t t e m p e r a t u r e compatible w i t h t h e m a t e r i a l s a v a i l a b l e i n t h e 1900's. The e n g i n e c o s t of t h e 2 2 ~ 9 ~ F e n g i n e i s 40 p e r c e n t i e s s t h a n t h a t of t h e 1800" F e n g i n e because o f a combination of f a c t o r s . F i r s t , t h e p h y s i c a l s i z e i s about 40 p e r c e n t smaller because t h e s p e c i f i c power improves s u b s t a n t i a l l y and because a s m a l l e r a i r c r a f t is r e q u i r e d t o do a g i v e n mission. I n a d d i t i o n , t h e 2200° F e n g i n e i n c o r p o r a t e s more c o s t - r e d u c i n g technology, which r e t a r d s t h e normal growth of c o s t w i t h temperature and which keeps t h e c o s t p e r u n i t a i r f l o w n e a r l y c o n s t a n t . Like- wise, t h e 2200° F e n g i n e weighs a b o u t 40 p e r c e n t l e s s t h a n t h e 1800° F engine.

I n f a c t , t h e a i r p l a n e f u e l consumption i s improved 1 5 p e r c e n t , n o t because t h e c y c l e e f f i c i e n c y improves ( i n f a c t , i t i s only 1 p e r c e n t b e t t e r ) , b u t because t h e e n g i n e weight is reduced. ?he e n g i n e weight and c o s t s a v i n g s a l s o produce 15 t o 20 p e r c e n t improvements i n a i i p l a n e a c q u i s i t i o n c o s t , o p e r a t i n g c o s t , and t o t a l c o s t of ownership.

I n a s i m i l a r v e i n f i g u r e 7 d i s p l a y s c y c l e p r e s s u r e r a t i o e f f e c t s . I n t h e lower p l o t , e n g i n e c o s t i s d i s p l a y e d a s a band t h a t was drawn from f o u r com- p r e s s o r p o i n t d e s i g n s : a s i n g l e - s t a g e c e n t r i f u g a l a t 9 : l p r e s s u r e r a t i o and a r e l a t i v e c o s t of 1.0, a two-stage c e n t r i f u g a l a t 20:1, a n a x i c e n t r i f u g a l a t 11.3:1, and a t h r e e - s t a g e a x i c e n t r i f u g a l a t 1 5 : l . The band w i d t h i n d i c a t e s t h e i n c r e a s i n g c o s t a s s o c i a t e d v i t h more compressor s t a g e s a t a f i x e d p r e s s u r e r a t i o . Cost i n c r e a s e s r a p i d l y w i t h p r e s s u r e r a t i o a s more compressor and t u r - b i n e s t a g e s a r e required. Likewise, a t any g i v e n horsepower l e v e l , weight in- c r e a s e s t o o , s o t h e power-to-weight r a t i o , shown i n t h e upper p a r t , becomes worse. U r ~ f o r t u n a t e l y , a t t h e s m a l l a i r f l o w s r e q u i r e d i n t h e s e a p p l i c a t i o n s ( 2 o r 3 l b / s e c ) , t h e c y c l e e f f i c i e n c y i s n o t i n c r e a s i n g r a p i d l y enough t o o f f s e t t h e s e a d v e r s e t r e n d s . I n f a c t , a s t h e SFC bard shows, t h i n g s a r e even worse t h a n t h a t . While t h e 11.3:l e n g i n e is 6 p e r c e n t more e f f i c i e n t t h a n t h e 9 :1 e n g i n e , t h e 1 5 : l and 20:l e n g i n e s a r e a c t u a l l y € , l i g h t l y worse t h a n t h e 11:l en- g i n e because t h e component e f f i c i e n c i e s a r e s u f f e r i n g t o o rmch a t t h e very s m a l l c o r r e c t e d a i r f l o w s i n t h e f i n a l s t a g e s . Hence, t h e minimum f u e l s o l u t i o n i s about a 12 o r 1 4 : l compressor p r e s s u r e r a t i o , b u t t h e lowest a i r c r a f t c o s t of ownership s o l u t i o n i s a b o u t 9:l.

A s a t h i r d example o f t h e s e t r a d e - o f f s , f i g u r e 8 p r e s e n t s a suuunary of o n e team's e f f c r t s t o d e t e r m i n e t h e b e s t o v e r a l l e n g i n e c o n f i g u r a t i o n f o r a medium p r e s s u r i z e d twin. They c o n s i d e r e d a t u r b o f a n w i t h a g a s - g e n e r a t o r c o n s i s t i n g o f They a l s o a s i n g l e c e n t r i f u g a l compressor hooked to a one-stage r a d i a l t u r b i n e .

c o n s i d e r e d a f r e e - t u r b i n e t u r b o p r o p i n t h r e e d i f f e r e n t v e r s i o n s : The f i r s t h a s t h e same simple arrangement a s t h e t u r b o f a n ; t h e second h a s a two-stage c e n t r i - f u g a l compressor; and t h e t h i r d h a s a n a x i a l t u r b i n e r e p l a c i n g t h e r a d i a l t u r - A c t u a l l y , t h e l i t t l e diagrams o n l y show t h e g a s - g e n e r a t o r p o r t i o n s , b u t bine.

And a l l f o u r o f t h e s e c o n f i g u r a t i o n s a l s o have a two-stage axial-power t u r b i n e .

f i n a l l y , t h e y c o n s i d e r e d two a r r a n g e m n t s o f a s i n g l e - s h a f t turboprop: b o t h u s e a s i n g l e c e n t r i f u g a l compressor, b u t t h e f i r s t h a s J n e a x i a l t u r b i n e f o l l o w i n g a r a d i a l t u r b i n e , and t h e second h a s t h r e e a x i a l s t a g e s . The e v a l u a t i o n c r i t e r i a a r e a i r p l a n e t o t a l c o s t o f ownership, f u e l consumption, o p e r a t i n g c o s t , a c q u i s i - and t h e v a l u e s quoted a r e a l l r e l a t i v e t o t h e second t i o n c o s t , and e n g i n e c o s t , o p t i o n - t h e s i m p l e s t f r e e - t u r b i n e turboprop. The b a r s a r e o r d e r e d from l e f t t o r i g h t i n t h e same sequence a s t n c t o p l i t t l e diagrams. The most obvious r e s u l t i s t h a t t h e t u r b o f a n i s simply n o t i n t h e running a t a l l . I t s p e n a l t i e s , which a r e caused by i t s low e f f i c i e n c y a t low f l i g h t speeds, run from 25 t o 65 per- c e n t . A c ~ u a l l y , t h e optimum c h o i c e is n o t t h e s i m p l e f r e e - t u r b i n e b a s e l i n e b u t r a t h e r t h e even s i m p l e r s i n g l e - s h a f t c o n f i g u r a t i o n w i t h on2 l e s s t u r b i n e s t a g e .

However, i n c o n s i d e r a t i o n of o t h e r f a c t o r s , e s p e c i a l l y commonality w i t h h e l i - c o p t e r t u r b o s h a f t r e q u i r e m e n t s , t h i s team m a r g i n a l l y p r e f e r r e d t h e f r e e - t u r b i n e b a s e l i n e .

A f t e r many t r a d e - o f f s such a s t h e s e and i t e r a t i o n s w i t h t h e marketing a n a l y s e s , t h e f o u r teams s e t t l e d on t h e c y c l e s and c o n f i g u r a t i o n s shown on t h e right-nand s i d e of f i g u r e 9. These e n g i n e s a r e a l l t u r b o p r o p s ranging from 335 t o 565 hp and a r e aimed p r i m a r i l y a t t h e h i g h - p e r f o n n ~ n c e s i n g l e - e n g i n e and twin-er.gine a i r p l a n e a p p l i c a t i o n s . For comparison, both a r e p r e s e n t a t i v e cur- r e n t p r o d u c t i o n turboprop (uncooled, o l d technology) and a h y p o t h e t i c a l t u r b o - prop irlcorporating c u r r e n t l y a v a i l a b l e modern technology a r e i l l u s t r a t e d on ttle left-hand s i d e o f t h i s f i g u r e . ~ l l i s o n ' s c h o i c e is a c o o l e d , 22000 P maximum t u r b i n e - i n l e t temperatLLe, 1 4 : l p r e s s a r e r a t i o , f r e e t u r b i n e design. Two cen- t r i f u s a l compressors a r e d r i v e n by two a x i a l t u r b i n e s t a g e s , and a n o t h e r two- s t a g e t u r b i n e d r i v e s t h e p r o p e l l e r load on a second spool. T h i s d e s i g n d i f f e r s from b o t h t h e c u r r e n t p r o d u c t i o n e n g i n e s and t h e Fy pothet i c a l modern e n g i n e s mainly i n having a b e t t e r c y c l e and h i g h e r component e f f i c i e n c i e s . I t s p e r f o r - mance i s much b e t t e r , a l t h o u g h i t s e s t i m a t e d c o s t d i f f e r s l i t t l e . G a r r e t t a l s o chose a two-stage f r e e power t u r b i n e , b u t s e l e c t e d a s i n g l e 9 : l centrifugal com- p r e s s o r d r i v e n by a one-stage r a d i a l t u r b i n e i n t h e i r p u r s u i t o f lowering c o s t .

T e l e d y n e ' s low-cost e n g i n e q u e s t l e d t o an e n g i n e f a m i l y t h a t is d e s c r i b e d But i t s b a s i c element i s s very simple c o r e e n g i n e which c o n s i s t s of a l a t e r .

s i n g l e c e n t r i f u g a l compressor connected t o a s i n g l e r a d i a l t u r b i n e . T h i s t u r - b i n e a l s o d r i v e s t h e p r o p e l l e r l o a d )n a common s h a f t . T h e i r t u r b i n e r ? t o r i s very advanced d e s i g n , which is d i s c u s s e d uncooled and r e q u i r e s a very s p o r t y , low c o s t through a very u n c o n v e n t i o n a l l a t e r . F i n a l l y , Williams Research sought approach. i t a t h e r t h a n t h e c a n v e n t i o n a l i d e a o f e l i m i n a t i n g c o s t by e l i m i n a t i n g p a r t s , t h e y propose u t i l i z i n g known ways o f producing very inexpensive p a r t s at lower c o s t . The r e s u l t was a n uncooled, s i n g l e - s h a f t , a x i c e n t r i i u g a l arrange- ment w i t h s i x a x i a l compressor s t a g e s and f o u r a x i a l t u r b i n e s t a g e s a t modest t e q e r a t u r e b u t r e l a t i v e l y h i g h p r e s s u r e r a t i o . T h i s concept is d e s c r i b e d i n m o r e det:ll l a t e r also.

E N Z ~ N E FERPORMNCE AND COST Ane performance e s t i m t e s f o r t h e s e e n g i n e s a r e s u ~ a r a r i z e d i n f i g u r e 10.

Usually ue t h i n k o f SFC r i s i n g smoothly as w e d e c r e a s e e n g i n e s i z e b e c a u s e o f a d v e r s e s c a l i n g e f f e c t s . However, whenever a new e n g i n e is i n t r o d u r e d , it may d i s t o r t o u r e x p e c t e d c u r v e simply because o f its advanced technology r e l a t i v e t o o l d e r , -21-established engines. T h i s happened a few y e a r s ago when t h e T703 was i3troduced. It yanke.1 t h e c u r v e down t o form a "knee" i n t h e t r e n d c u r v e s a t 1500 hp. E x a c t l y t h e same t h i n g vuuld happen a g a i n i f GATE technology en- g i n e s were i n t r o d u c e d at k00 t o 600 hp, s i n c e t h e y would b e 20 p e r c e n t more e f - f i c i e n t t h a n c u r r e n t p r o d u c t i o n e n g i n e s o f t h e same size. The t e c h a o l a g i e s t h a t l e a d t o t n i s a r e d e s c r i b e d l a t e r .

The t h r e e l o r c o s t - t h e m e s t u d y t e a s p r o j e c t e d e n g i n e c o s t s i n t w o ways.

The f i r s t p r e s u a e s no i n c r e a s e i n p r o d u c t i o n r a t e s and s i q l y r e f l e c t s t h e in- t r i n s i c c o s t - r e d u c t i o n p o t e n t i a l o f u s i n g advanced technology ( f ig. 11). The magnitude of t h i s s a v i n g is about 40 percent. I n o t h e r words, GATE e n g i n e s would b e 40 p e r c e n t c h e a p e r t o produce t h a n t o d a y ' s engines. B u t , once a s a v i n g o f t h i s s i z e m a t e r i a l i z e s , i t would t r i g g e r i n c r e a s e d s a l e s , and t h i s opens up the p o s s i b i l i t y o f a new manufacturing f a c i l i t y d e d i c a t e d s p e c i f i c a l i y t o GATE e n g i n e s , which, i n t u r n , would c a u s e e v e n f u r t h e r s a v i n g s - f o r z t o t a l reduc- t i o n o f as much a s 60 p e r c e n t . A t t h e same t i m e , market demand would i n c r e a s e t o t h e neighborhood of 10 000 engines/year/coupany (assuming t h a t tro companies s p l i t t h e market e q u a l l y ) . Hence, w i t h o u t s a c r i f i c i n g t o o much performance, t h e p u r s u e r s of t h e i o r c o s t theme a r e p r e d i c t i n g t h a t GATE technology c o u l d p r o v i d e t h e key t h a t unlocks t h i s p o t e n t i a l . To p u t t h e c o s t e s t i m a t e s i n b e t t e r p e r s p e c t i v e , f i g u r e 12 shows e n g i n e s p e c i f i c c o s t e s t i m a t e s f o r a l l f o u r companies a g a i n s t a backdrop c.f t h e c u r r e n t c o s t s i t u a t i o n . C u r r e n t t u r b o p r o p s c o s t about t h r e e t i m e s a s much as p i s t o n engines.

But remember, t u r b o p r o p p r o d u c t i o n r a t e s a r e two o r d e r s o f magnitude l e s s t h a n p i s t o n e n g i n e r a t e s . A l l i s o n ' s r a t h e r s o p h i s t i c a t e d machine is e s t i m a t e d t o c o s t about t h e same a s c u r r e n t t u r b o p r o p s , and it t r i g g e r s modest i n c r e a s e s i n s a l e s . The l o r c o s t theme e s t i n a t e s o f Williams, G a r r e t t , and Teledyne a t 10 000 u n i t s p e r y e a r c l o s e l y approach t h e p i s t o n e n g i n e c o s t band, a2d t h i s o b v i o u s l y r e p r e s e n t % a major d e p a r t u r e from t o d a y ' s scenario. Of c o u r s e d o l l a r s p e r horsepower a l o n e is n o t s u f f i c i e n t s i n c e d i f f e r i n g l a p s e r a t e s , i n s t a l l a t i o n f a c t o r s , f u e l consumption, e t c . , a r e e q u a l l y important c o n s i d e r a t i o n s . The n e t ef f c c t of a l l t h e s e f a c t o r s is shown l a t e r i n t h e mi,sion a n a l y s i s r e s u l t s .

TECHNOLOGIES But what a r e t h e t e c h n o l o g i e u behind t h e s e improvements: C e r t a i n l y nobody c o u l d go o u t today and start b u i l d i n g e n g i n e s l i k e these. A c t u a l l y , it is some- u h a t d i f f i c u l t t o s u c c i n c t l y s u ~ l u r i z e t h e advanced t e c h n o l o g i e s i d e n t i f i e d i n t h e s e s t u d i e s because e a c h s t u d y t e a m i n c o r p o r a t e d d i f f e r e n t o n e s - a t l e a s t i n d e t a i l t h e y a r e d i f f e r e n t . N e v e r t h e l e s s , f i g u r e 1 3 lists s o ~ e o f them i n a corn- p o s i t e f a s h i o n , a l t h o u g h n o t a l l of t h e s e w u l d b e p r e s e n t i n a s i n g l e d e s i g n .

I n t h e gearbox a r e a t h e u s e o f powdered-metal g e a r s and laser hardening was C o u p e s i t e r e c o ~ l w n d e d by s e v e r a l ceams t o reduce c o s t and improve p r o p e r t i e s .

m a t e r i a l g e a r c a s t s were r e c o ~ r e n d e d by A l l i s o n f o r s t i f f n e s s and weight s a r i n g s , w h i l e T e l i d y n e s ~ l q g e s t e d d i e - c a s t aluminum f o r a c o s t saving. Teledyne a l s o recatmended a composite d r i v e s h a f t , and n e a r l y a l l teams r e c m e n d e d f u l l a u t h o r i t y d i g i t a l c o n t r o l s . However, t h e key e l e ~ e n t s i n a l l of t h e c o n c e p t s Except f o r U i l l i a m s , e a c h team sought high- involved t h e r o t a t i n g machinery.

p e r f n m a n c e c e n t r i f u g a l compressors u s i n g advanced a n a l y s i s t e c h n i q u e s , so= f o w o f p a s s i v e c l e a r a n c e c o n t r o l , and backward crrrvat:lre. High s t a g e l o a d i n g s without severe performance p e n a l t i e s bere p r e w l a n t . And new manufacturing p r o c e s s e s , such as u s i n g p w d e r e d - m e t a l t i t a n i u m f o r t h e r o t o r s , appeared.

Technologies f o r t h e c o r e t u r b i n e were e s p e c i a l l y d i v e r s e . Two companies s e l e c t e d high-temperature r a d i a l d e r s i g n : Te ledyne, w i t h a n uncooled. powered- laminated c o n s t r u c t i o n process. A l ~ i s o n m e t a l concept and G a r r e t t w i t h a c o o l e d , s e l e c t e d a high-temperature a x i a l arrangement wi:h a c o o l e d , d u a l - p r o p e r t y r o t o r and p o s s i b l y ceramic s t a t o r s . Again, p a s s i v e c l e a r a n c e c o n t r o l was c o s t e f - f e c t i v e , and improvement of e f f i c i e n c y through be t t e r three-dimens i o n a l f low a n a l y s i s is required. S i m i l a r improvements uere i d e n t i f i e d f o r cou&ustors and power t u r b i n e s .

The f i r s t o i s e v e r a l example key technology e l e m e n t s is i l l u s t r a t e d i n f i g - u r e 14. T h i s one r e p r e s e n t s t h e a t t a i n m e n t of a 9 :1 p r e s s u r e r a t i o compressor, a t high e f f i c i e n c y , i n a s i n g l e s t a g e . It r e q u i r e s advanced three-dimensional b l a d i n g w i t h high t i p s p e e d s and high i n d u c e r nach numbers, a n improved t h r e e - dimensional Gif f u s e r , b e t t e r f l o w a n a l y s i s and b e t t e r e x p e r i m e n t a l measurements, imprc ed s u r g e margin, and a low-cost f a b r i c a t i o n t e c h n i q u e t h a t y i e l d s e s s e n - t i a l l y a n e t shape p a r t from powered-metal t i t a n i u m . The b e n e f i t is t o improve compressor e f f i c i e n c y by 3-112 p o i n t s r e l a t i v e t o a c u r r e n t technology 9 :l.

s i n g l e - s t a g e machined compressor, w h i l e reducing c o s t t o be c o a p e t i t i v e w i t h c a s t d e s i g n s . While t h e 6 p e r c e n t , e n g i n e c o s t s a v i n g is n o t a s l a r g e a s f o r some o t h e r components, i t a l s o s a v e s 6 p e r c e n t i n e n g i n e w e i g h t , compared w i t h a two-stage compressor t h a t v o u l d o t h e r w i s e be r e q u i r e 4 b ' t h c u r r e n t technology.

The second example is Teledyne's proposed uncooled, but high-temperature, r a d i a l - i n i l o v t u r b i n e . Its c o n c e p i is based on t h e i r r e c e n t development expe- r i e n c e w i t h a 120-hp t u r b o g e n e r a t o r s e t f o r t h e Amy p l u s some encouraging ana- F i g u r e 15 shows t h e r e s u l t s of a p r e l i s i n a r y a n a l y s i s t o v e r l f y l y t i c a l work.

t h e c o n c e p t ' s l i f e p o t e n t i a l . The s t r e s s - r u p t u r e l i f e was e v a l u a t e d f o r two d i f f e r e n t b l a d e geometries. One is r e l a t i v e l y t h i n i n t h e r o o t r e g i o n and h a s a c r o s s - s e c t i o n a l a r e a t h a t t a p e r s dovn a t t h e t i p t o 1/16 t h a t a t t h e roor ( i . e . , it h a s an a r e a t a p e r r a t i o A T R of 15). The o t h e r is t h i c k e r a t t h e root and h a s a t a p e r r a t i o of 31. The e v a l u a t i o n a s s m e d t h e u s e o f equiaxed I N - l O O , 8 c u r r e n t m a t e r i a l , t o g i v e a high c o n f i d e n c e level. A c t u a l l y t h m u g h , advanced m a t e r i a l s and d i r e c t i o n a l s o l i d i f i c . :ion would probably be used to i n c r e a s e t h e d e s i g n ' s i n t e g r i t y . T h i s d e s i g n is wry h i g h l y loaded, w i t h t i p speeds ap- proaching 2500 f t l s e c and t r a n s o n i c e x i t v e l o c i t i e s a t a maxinun t u r b i n e g a s t e m p e r a t c r e of 22500 F. Under t h e s e c o n d i t i o n s , t h e b l a d e metal temperature i s l8t)oo F a t t h e t i p s , and t h e l i f e t i m e is o n l y 200 t o 100 hr. Hovever, t h e i r e n g i n e is f l a t r a t e d and w i l l n o t r e q u i r e s u c h high t e q e r a t u n s a t take- o f f o r a t any o t h e r normal c o a d i t i o n . A t c r u i s e , t h e g a s temperature is down t o 195W F , which y i e l d s 1552O P uxiru metal temperatures. T h i s y i e l d s a 3000-hr l i f e f o r t h e 1 b : l ATR d e s i g n o r a 10 000-hr l i f e f o r t h e 3 l : l ATR de- sign. However, it is n o t c e r t a i n t h a t t h e 31:l ATR is p r a c t i c a l because of in- c r e a s i n g flow-path r e s t r i c t i o n s i n t h e r o o t r e g i o n s (lore d e t a i l e d a n a l y s e s are r e q u i r e d t o d e t e r m i n e a n opt-- ATIL).

The t h i r d e x g p l e is G a r r e t t ' s cooled, r a d i a l t u r b i n e concept ( f i g . 16). I t c o n s i s t s of a set of photoetched l a m i n a t e s diffusion-bonded t o form i n t e g r a l c o o l i n g passages. A f t e r bonding, t h e p a r t is electroche!mically m i l l e d to t h e f u l 1 three-dimensional d e s i red a e r o d y r u r i c shape. Advanced p o w d e r e d - r t a 1 sheet s t o c k f a b r i c a t i o n methods must be used t o lwer c o s t and thereby permit t h e u s e of h i g h - s t r e n g t h m a t e r i a l s +uch as Astroloy.

The n e t b e n e f i t w u l d be a 9.8 p e r c e n t e f f i c i e n c y improvement r e l a t i v e t o c u r r e n t , c o o l e d a x i a l t u r b i n e s w h i l e reducing c o o l i n g bleed 20 p e r c e n t and e n g i n e c o s t 2 1 percent.

The f i n a i example is t h a t o f a n approach t h a t d o e s n o t a p p l y s p e c i f i c a l l y t o a s i n g l e c o q m e n t b u t r a t h e r i n f l u e n c e s t h e e n t i r e engine. I t is Williams' unconventional approach t o lowering c o s t through t h e u s e of restricted-geometry b l a d e and vane aerodynamic shapes. The concept is t o d e s i g n f o r very low, r a t h e r t h a n high, stress levels as d e p i c t e d i n f i g u r e 17. T h i s a l l o w s perhaps a 150° t o 200° F i n c r e a s e i n t u r b i n e metal temperature w i t h o u t c o o l i n g , o r , w i t h a n advanced p a t e r i a l , u c h h i g h e r t e q e r a t u r e s t o e x p l o i t t h e i r s p e c i a l l y shaped t e m p e r a t u r e - s t r e s s curves. With M 6000E, f o r example, a n e x t r a 3000 is p o s s i b l e . E i t h e r way, t h e lower stresses ( p e r h a p s 112 of c u n v e n t i o n a l ) imply lower d e s i g n speeds, and t h i s , i n t u r n , means lower b l a d e loadings, which permit t h e u s e of l o r c o s r , s i q l i f i e d b l a d e manufacturing techniques. S p e c i f i c a l l y , ~ 1 1 compressor blader c o u l d h a w t h e s o e a i r f o i l s e c t i o n , be o f c o n s t a n t chord and camber, and be uniformly t w i s t e d ; i n f a c t , because o n l y t h e l e m t h s w u l d d i f f e r , t h e p a r t s c o s t would b e d r a m a t i c a l l y lowered. The c o r o l l a r y is t h a t h i g h e r p r e s s u r e r a t i o s a r e o b t a i n a b l e without m c h c o s t penalty. Then a l l t h e b l a d e s a r e h e l d i n p l a c e a s t h e hub i s formed around them i n a s i n g l e opera- t ion. The compressor v a n e s and a l l o t t h e t u r b i n e a i r f o i l s a r e f o r r c d i n t h e sarac way. The t o t a l r e s u l t is a very d i f f e r e n t looking small e n g i n e concept which a t t e m p t s t o a c h i e v e low c o s t without performance s a c r i f i c e by incorpor- a t i n g a l a r g e nrrmber of very l o r c o s t p a r t s i n s t e a d of 8 very small number of r e l a t i v e l y expensive p a r t s .

ON CORE Another concept f o r reducing e n g i n e c o s t i n v o l v t s u s i n g a c o m n c o r e f o r a family of engines. R e t a i n i n g p a r t s comolonality w i t h o u t s a c r i f i c i n g t o o much performance i s t h e key h e r e because e a c h of t h e d i v e r s e mission a p p l i c a t i o n s p r e f e r s a d i f f e r e n t o p t i m m engine.

One approach t o t h i s dilemma is i l l u s t r a t e d i n f i g u r e 18 which shows Teledyne's C9 c o r e e n g i n e s l i g h t l y d i f i e d t o a c c o m - d a t e s a c a d d i t i o n a l p a r t s t h a t are r e q u i r e d t o r e c o n f i g u r e t h e e n g i n e f o r more pouer. T h i s is done by a d d i n g a s u p e r c h a r g i n g a x i a l compressor s t a g e , a n a x i a l t u r b i n e s t a g e t o p r o v i d e t h e e x t r a power, and a set o f e x t r a g e a r s , which are d u p l i c a t e s a f t h e f i r s t set t o h a n d l e t h e i a c r e a s e d pouer. T h i s a l l o w s a 335-hp e n g i n e t o grow 70 p e r c e n t t o a 565-hp d e r i v a t i v e w i t h on a 4-inch e x t e n s i o n ( E m - 3* to 3 8 in. 1 , a 31-lb weight i n c r e a s e (from 172 t o 203 l b ) , and a 5 6 per- c e n t i n c r c a s e i n c o s t ( f ig. i91. A t t h e sime time, t h e SFC is 10 p e r c e n t l o v e r d u e t o t h e i n c r e a s e d c y c l e temperature and p r e s s u r e and copponent rematching.

The p r i c e o f c o m o a a l i t y i n t h i s c a s e i s a 2 p e r c e n t SFC p e n a l t y f o r t h e b a s i c c o r e e w i n e . T h i s r e s u l t s f m - t h e l o v e r t u r b i n e temperature r e q u i r e d t o a c c o r d a t e a c o m o n fixed-area nozzle. The b e n e f i t s o f t h i s approach t o c o r o n a l i t y are a 7 p e r c e n t lower c o s t and a 1 6 p e r c e n t v c i g h t r e d u c t i o n f o r t h e C9 335-hp v e r s i o n r e l a t i v e t o t h e n e x t b e s t approach, v h i c h is u s i n g a s i n g l e , l a r g e con- f i g u r a t i o n and t h e n s h a v i n g t h e flowpath a r e a t o reduce pouer.

AIRCRAE T BENEFITS The e f f e c t t h a t t h e s e t e c h n o l o g i e s , b o t h i n d i v i d u a l l y and c o l l e c t i v e l y , would have i f GATE e n g i n e s were i n s t a l l e d i n conventionn! hct +lie.:?j- i+rc-.-=.'

a i r f r a w s f l y i n g m i s s i o n s moderately Pore d i f f i c u l t t h a n t o d a y ' s w i l l be i l l u s - t r a t e d v i t h s e v e r a l examples. I n each c a s e t h e h y p o t h e t i c a l a i r c r a f t is r e s i z e d t o a c c m t e t h e new engines. t h e A l l i s o n CATE e n g i n e was compared F i r s t , v i t h a s c a l e d turboprop v e r s i o n o f t h e i r most r e c e n t l y improved 250 series t u r boshaf t engine. Tireir CATE e n g i n e i n c o r p o r a t e s c o n s i d e r a b l e m a t e r i a l s and aero- therrcrdynamic improvement v h i c h a c c o u n t s f o r h i g h e r c y c l e e f f i c i e n c y and s m a l l e r s i z e . S p e c i f i c a l l y , a dual-property, u i a l , high-pressure t u r b i n e , w h i l e s l i g h t l y more expensive i n i t i a l l y , y i e l d s long l i f e and much less e n g i n e main- tenance c o s t . S i m i l a r l y , a t r a n s p i r a t i o n c o o l e d , Lamillog combustor, w h i l e n o t inexpensivr i t s e l f , a l l o w s t h e u s e o f a s h o r t , compact, and long-TBO ( t i m e be- tween o e r h a u l s ) combust i o n system. Ceramic r o t o r s e r e n o t judged a p p r o p r i a t e f c r manned a i r c r a f t a p p l i c a t i o n i n t h i s t h e frame, b u t A l l i s o n s u g g e s t e d t h a t ceramic s t a t o r s may be, a l t h o u g h even t h e y a r e o n l y marginal. L a s t l y , a f i b e r g l a s s / p o l y i m i d e c o ~ p o s i t e gearbox shoved a s l i g h t c o s t advantage. A 1 li- s o n ' s advanced tecnnology engine y i e l d e d 20 p e r c e n t b e t t e r SFC and 1 3 p e r c e n t l e s s weight, It c o s t s 3 p e r c e n t more t o buy, but 35 p e r c e n t less t o m a i n t a i n t h a n a comparable c u r r e n t turboprop. A range of a i r c r a f t b e n e f i t s a r e shovn i n t h e following list corresponding t o t h e t h r e e a i r c r a f t t y p e s t n a t t h e y i n v e s t i - g a t e d ( a n u n p r e s s u r i z e d twin, a heavy t u i n , and a twin-engine h e l i c o p e r ) : Technologies: 1. Advanced m a t e r i a l s and aerothermodynamics - h i g h e r c y c l e e f f i c i e n c y and s m a l l e r s i z e

2. Dual p r o p e r t y a x i a l high-pressure t u r t i n e - much lower maintenance

c o s t (5000 h r TBO)

3. Ceramic t u r b i n e s t a t o r - s l i g h t c o s t r e d u c t i o n

4. Lamilloy combustor - p e r m i t s 5000 h r TBO a t high temperature

5. Composite gearbox c a s e - s l i g h t c o s t r e d u c t i o n

E m i n e improvements :

1. SFC - 20 p e r c e n t

2. Weight - 2 3 p e r c e n t

3. Cost -3 p e r c e n t

k. Maintenance c o s t - 3 5 p e r c e n t

A i r c r a f t b e n e f i t s :

L . Fuel burned - 2 3 t o 32 p e r c e n t less

2. Gross e i g h t - 11 t o 21 p e r c e n t less

3. Purchase p r i c e - ~7 p e r c e n t

4. Ownership c o s t - 8 t o 20 p e r c e n t less

Although t h e purchase p r i c e s d o n o t change much, 23 t o 32 p e r c e n t less f u e l is burned, and ownership c o s t s d r o p 8 t o 20 p e r c e n t .

I n a similar way, t h e o t h e r companies l i s t e d t h e technology e l e n t s t h a t s u r v i v e d t h e i r s c r e e n i n g p r o c e s s e s and o r d e r e d them a s shown i n t a b l e 1, a Garrett e x i a p l e . The b e n e f i t s o f e a c h of t h e advanced t e c h n o l o g i e s are g i v e n r e l a t i v e t o a h y p o t h e t i c a 1, a1 1-new e n g i n e u s i n g c u r r e n t l y a v a i l a b l e t e c h n o l - ogy. For example, t h e high-pressure laminated t u r b i n e technology raises t h e c o r e t u r b i n e e f f i c i e n c y by 9.8 p e r c e n t , reduces e n g i n e c o s t 21 p e r c e n t , u e i g h t 7 p e r c e n t , and SFC 7.4 p e r c e n t and y i e l d s a b e n e f i t c o s t r a t i o o f 561. The bene- f i t is d e f i n e d a s t h e o v n e r s h i p s a v i n g o v e r 20 y e a r s f o r a f l e e t of 1 5 000 w d i w s i z e d , twin-engined a i r p l a n e s . The c o s t i s t h e r e s e a r c h investment re- q u i red t o d e r a n s t r a t e technology readiness.

The si ngle-stage powdered e r a 1 t i t a n i r n advanced compressor is 1 p e r c e n t less e f f i c i e n t t h a n a machined two- s t a g e c u r r e n t technology compressor; y e t it c o s t s and weighs enough l e s s t o o f f - set t h i s penalty. Another technology w i t h l a r g e b e n e f i t s i s a low-pressure t u t - b i n e t h a t o p e r a t e s a t a h i g h work f a c t o r b u t low speed. C o l l e c t i v e l y , t h e s e t e c h n o l o g i e s p r o v i d e a 36 p e r c e n t lower c o s t r e d u c t i o n , 20 p e r c e n t l i g h t e r u e i g h t , and 1 3 p e r c e n t b e t t e r SFC r e l a t i v e t o t h e b e s t t h a t we c o u l d d o v i t h t o d 8 y e s a v a i l a b l e technology. One o f t h e key e l m e n t s is c l e a r l y t h e laminated t u r b i n e technology, which p r o v i d e s roughly one-ha l f of t h e b e n e f i t s .

A i r c r a f t Fue 1 Now w e c a n r e t u r n t o t h e most c h a l l e n g i n g i s s u e , i d e n t i f i e d a t t h e o u t s e t : comparing advanced GATE t y p e e n g i n e s w i t h p i s t o n engines. One of tne disadvan- t a g e s of c u r r e n t t u r b o p r o p s is t h a t t h e y consume t o o rmch f u e l : about 10 per- c e n t more t h a n c u r r e n t p i s t o n e n g i n e s f o r a t y p i c a l twin-engine a i r c r a f t m i s - sion. T h i s is because t h e i r i n s t a l l e d c r u i s e thrust-SFC is i n f e r i o r . GATE en- g i n e s would e l i m i n a t e most of c h i s SFC d i f f e r e n c e as showr. i n f i g u r e 20.

S i n c e t h e i r i n s t a l l e d e n g i n e weight is o n l y 1 / 3 o r 1/4 a s larch as a r e c i p r o c a t i n g en- g i n e , t h e r e s u l t i n g GATE-powered a i r p l a n e would a c t u a l l y s a v e 5 t o 15 p e r c e n t S i n c e avgas c o s t s a s auch a s 20 p e r c e n t more p e r BTL!, t h e r e a l f u e l c o s t f u e l .

s a v i n g s are s u b s t s n t i a l l y g r e a t e r t h a n t h a t .

A i r c r a f t Economics A r e p r e s e n t a t i v e i l l u s t r a t i o n of how a GATE-powered a i r p l a n e compares w i t h a reciprocating-powered a i r p l a n e i n economic t e r m s is shorn i n t a b l e 11 f o r a l i g h t - t w i n a i r p l a n e t h a t c r u i s e s a t 10 000 f e e t a t 225 k n o t s f o r 1100 nmi, is flown 500 h d y r , and is s o l d a f t e r 3 years. The b a s e l i n e is a c u r r e n t - technology, reciprocating-powered a i r p l a n e t h a t r e q u i r e s two 380-hp p i s t o n en- g i n e s weighing 550 l b , each, t h a t t o g e t h e r b u r n 172 g a l l o n s o f f u e l . The air- p l a n e t a k e o f f t r r i g h t is 6200 l b ; t h e e n g i n e s c o 8 t $11 000 e a c h ; t h e a i r p l a n e c o s t s $207 000 t o t a l ; it c o s t s $ 5 l / h r t o o p e r a t e and, f o r t h e t h r e e - y e a r owner- s h i p p e r i o d , c o s t s a t o t a l of $170 OW. The p e r c e n t changes f o r t h r e e d i f f e r e n t advanced e n g i n e o p t i o n s are shown i n t h e right-hand columns. The f i r s t is a n improved r e c i p r o c a t i n g e n g i n e p r e s u a i n g simply 10 p e r c e n t lower SFC. I t pro- duces r a t h e r modest a i r c r a f t e c o n a a i c improvements: 5 p e r c e n t i n t o t a l c o s t of ownership.

Option 2 is a c u r r e n t technolcgy turboprop, b u t produced a t a rate of 1 0 000 u n i t s p e r year.

It too is n o t very a t t r a c t i v e - o n l y a 3 p e r c e n t n e t

savings. Option 3, one o f t h e l w - c o s t GATE turboprops, is a r c h more a t t r a c - t i v e .

T h i s a i r p l a n e w u l d be 20 p e r c e n t s m a l l e r and b u m 8 p e r c e 3 t Xess f u e l , and, a l t h o u g h t h e engine c o s t is up 2 3 p e r c e n t , t h e c o q l e t e a i r c r a f t c o s t is down 14 p e r c e n t , and t h e o p e r a t i n g c o s t i s down 28 p e r c e n t , f o r a total o v n e r s h i p s a v i n g of 20 percent.

I f w e expand o u r s c o p e t o i n c l u d e t h e o t h e r 1-cost GATE v e r s i m s and o t h e r a p p l i c a t i o n s ( f i g . 211, w e see t h a t , as a class, the twin-turboprop a i r p l a n e s would c o s t 1 5 t o 25 p e r c e n t l e s s t o buy and 30 t o 40 p e r c e n t less t o op- erate t h a n t h e i r piston-powered c o u n t e r p a r t s . However, the b e n e f i t s f o r high- performance, s i n g l e - e n g i n e a i r p l a n e s are o n l y one-third t o one-half as much, Nevertheless, any economic b e n e f i t s a t a l l =st be c o n s i d e r e d a bonus, i n a s m c h as t h e argument f o r t u r b i n i z a t i o n c o u l d be p r e d i c a t e d on noneconoric v i r t u e s alone, The obvious q u e s t i o n is: When d o t h e s e economic b e n e f i t s d i s a p p e a r ? A rough estimate of t h i s is shova i n f i g u r e 22 where a f e u d a t a p o i n t s from e a c h s t u d y are p l o t t e d i n terms of t h e r e d u c t i o n i n o m e r s h i p c o s t o f GATE-powred a i r p l a n e s r e l a t i v e t o c u f s e n t reciprocating-powered a i r p l a n e s as a f u n c t i o n o f t h e r c q u i r e d s h a f t horsepower f o r t h e r e c i p r o c a t i n g a i r c r a f t version. The twin- e n g i n e a i r p l a n e d a t a l o o k s impressive, s h w i n g 20 t o 30 p e r c e n t b e n e f i t s .

The single-engine d a t a a r e t o o s p a r s e t o be c e r t a i n , b u t it a p p e a r s as though t h e e c o m i c i n c e n t i v e goes t o z e r o somewhere i n t h e 20i)-hp region. Of c o u r s e , even a t z e r o o r s l i g h t l y n e g a t i v e economic change t u r b i n i z a t i o n is s t i l l a t t r a c t i v e .

Obviously, major b e n e f i t s of t h i s magnitude c a u s e a l a r g e impact i n t h e marketplace. The marketing f o r e c a s t s t h a t g o along w i t h t h e preceding a r e sum- marized i n f i g u r e 23 f o r e a c h s t u d y team i n terms of t h e t o t a l number of t u r b i n e e n g i n e s produced, both w i t h and w i t h o u t a n i n s t a n t a n e o u s l y w t u r e GATE e n g i n e i n 1988. S i n c e GATE technology e n g i n e s c o u l d n o t a c t u a l l y even e n t e r s e r v i c e u n t i l t h e e a r l y 19901s, t h i s is merely a n i n d i c a t i o n o f impact r a t h e r t h a n a n a c t u a l f o r e c a s t . The p i c t u r e is c e r t a i n l y s t r i k i n g because of t h e q u i t e d i f f e r e n t es- timates. ~ l l i s o n ' s modest f o r e c a s t is i n agreement w i t h t h e i r more c o n s e r v a t i v e c o s t e s t i m a t e s , w h i l e Teledyne p r e d i c t s a huge g a i n due t o t h e i r lower c o s t es- t i m a t e s and broad engine-size family. A l l of t h e e s t i m a t e s a r e much g r e a t e r t h a n t h e 1500 e n g i n e s pmduced i n 1976. Half of t h e s e were t u r b o s h a f t e n g i n e s f o r h e l i c o p t e r s . Hcwever, t h e f u t u r e GATE s c e n a r i o f o r e c a s t s t h a t t h e turboprop would s t r o n g l y dominate. A composite average of t h e s e f o u r f o r e c a r ~ s i s shown i n f i g u r e 24. A t o t a l of 2 0 000 GATE technology t u r b i n e e n g i n e s would be man- u f a c t u r e d a n n u a l l y , mostly t u r b o p r o p s , compared w i t h o n e - f o u r t h a s many w i t h o u t GATE technology. The a i r c r a f t market r e s u l t s a r e shown w i t h t h e p i e c h a r t s , U i t h o u t GATE, t h e t u r b o p r o p both w i t h and w i t h o u t GATE technology engines.

s h a r e is f o r e c a s t t o grow from i t s c u r r e n t l e v e l of 2 p e r c e n t to a l e v e l of 5 p e r c e n t . With GATE, i t would g r o v t o a b o u t 35 p e r c e n t , o r a s e v e n f o l d in- c r e a s e . The twin-piston market would p r a c t i c a l l y d i s a p p e a r , from 12 t o 2 per- c e n t , w h i l e t h e s i n g l e - e n g i n e p i s t o n p o r t i o n would s h r i n k from 68 t o 4 7 p e r c e n t , b u t it still would remain very l a r g e .

A sumaary o f what we p e r c e i v e t h e major s t u d y r e s u l t t o be i s d i s p l a y e d i n f i g u r e 25. The most c h a l l e n g i n g , b u t rewarding, o p p o r t u n i t y f o r s m a l l g e n e r a l - a v i a t i o n t u r b i n e e n g i n e s lies i n t h e 300 t o 600 hp region. Here, t h e p r o p e r improved m a t e r i a l s , h i g h e r component e f f i c i e n - c o a b i n a t i o a o f s i m p l e r d e s i g n , c i e s , c h e a p e r manufacturing t e c h n o l o g i e s , and c o r e coumwnality c o u l d r e s u l t i n s u f f i c i e n t l y lower e n g i n e c o s t , SFC, and weight t o overcome t h e t r a d i t i o n a l t u r - b i n e e n g i n e c o s t b a r r i e r a t t h e 500 hp s i z e . P l o t t e d h e r e a r e t h e t r e n d s of a i r c r a f t c o s t v e r s u s e n g i n e s i z e , and t h e l a r g e gap between r e c i p r o c a t i n g - powered and turboprop-powered a i r c r a f t is apparent. GATE technology permi ts l a r g e improvements i n a i r c r a f t economies a t t h e upper end o f t h e r e c i p r o c a t i n g - powered c l a s s and f i l l s i n t h e g a p between t h e r e l a t i v e l y i n e x p e n s i v e r e c i p r o - c a t i n g a i r c r a f t and t h e expensive t u r b o p r o p a i r c r a f t . I n t u r n , t h i s b r i n g s t h e many o t h e r v i r t u e s of t u r b i n e e n g i n e s t o a auch b r o a d e r spectrum of u s e r s and a p p l i c a t ions.

REFERENCES 1. Baerst , C. F. ; and F u r s t , D. G. : General A v i a t i o n T u r b i n e Engine (GATE) Study F i n a l Report. N A S A CR- 159482. ( AiResearch-21-2997, AiHesearch Manu- f a c t u r i n g Co. of ~ r i z o n a ; USA C o n t r a c t NAS3-20755. ) NASA CR-159482, 1979.

2. G i l l , J. C., e t d l . : Study of a n Advanced General A v i a t i o n T u r b i n e Engine (GATE). (EDR 9528, D e t r o i t D i e s e l ~ l l i s o n ; N A S A C o n t r a c t W3-20756. ) N A S A CR-159558, 1379.

3. Smith, R. ; and Benstein, E. H. : Advanced General A v i a t i o n T u r b i n e Engine (GATE) Study. (CAE-lb00, Teledyne CAE; N A S A C o n t r a c t NAS~-20757.) N A S A CR-159624, 1979.

4. Lays, E. J.; and Murray, C. L. : Advanced General A v i a t i o n T u r b i n e Engine (GATE) Concepts. (wCR-78-113-15, Williams Research Corp., N A S A C o n t r a c t NAS3-20758. ) N A S A CR-159b03, 19 79.

Table I A D V A N C E D TECHNOLOGY BENEFITS i r ' A WT. A COST. m F R l A SFC.

TLCHmKOtY A 1, 8 COST PTS RATIO -7 -21 %1 -7.4 HP W w T E E +QS

4 - 4 232 + 1 4

W T I S W L E STAtE CDMPRLSSOR -LO

---

0 - 1 U1 0 L o l l con n r L NOZZES

---

0 - 2 1 3 2 0 ELECTRONIC C W O L - 7 . 0 - 5 4 8 - 7 . 0 H I G H W ~ S R E D L P T U R B M 6 0

-

0 - 3 226 0 . USER H A - CLARS

--

-8 -3 rPUVGI 1 TOTAL m s L CHANGES ARE R U T W E TO HYPOlWlICAL C U R E N l lEUmYDGY N R B H D Y G W 2 C~~ CONlROL lENFlTS ARE N C W N A M yr FOR l5 om m l r m 1 B E N f R # F W D AS W M R S H I P COST SAVHCS OYER T W H N C 4 C O S T l S M S A R & T C O S l Table 11 GATE T U R B O P R O P A I R P L A N E S WOULD BE CHEAPER lEmoGY

' ADV IECH CURRENl GATE

~ECH ~ R B O P R O P Z SHP. SiS TO 3 0 E N C N WEIGHT HO I b 1 7 2 gal MISSION M L GROSS WEIGHT 6 2 0 0 l b E K N COST $11 @O ACOUISRW COST t Z O l K OPERATMG COST S 5 U h r TOTAL COST OF MlHERSHlP S 1 7 0 K SOURCE: GARRffl 0 - 7 9 - 4 1 9 0 MAINTENANCE COST n-19-1712 Figure 3

CURRENT ENGINE SELECTION FOR LIGHT AIRPLANES

r U I 0 1 N t 3 1 S A O V A N T l l a s Figure 4 EXPLOITING ENGINE TECHNOLOGY DIFFERENTLY I 0 L I I 1 . 1 . 2 . 3 . 4 .5 ,->--.,-.:;< ENGINE EFFICIENCY Figure 5

TURBOPROP MEDIUM PRESSURIZED TWIN

FhEE TUaBlNE SINGLE CENTRIFUGAL COMPRESSOR TURBl NE RADIAL (;AS GENERATOR TURBINE TEMP.

COST CONSUMPTION COST COST COST SOLIRCE: GARRETT Figure 6 EF FECTS OF iNCREASlNC CYCLE PRESSURE RATIO Q 6 RLLATIVE lblhp LOO R f L A T M COST CYCLE PRESSURE RATIO SOURCE: TELEDYM CAE . - .- Figure 7 GAS GENERATOR CONFIGURATIONS MEDIUM PRESSUR'ZED N ! I N TURBOFAN FREk-TURBINE TURBOPROP SINGLE-SHAFT TURBOPROP

a E X 3 0 0

L RELAT NI VALUE L ACQUISITION COST iONSUhlPTlON COST COST COST SOURCE: GARRETT .:,..: ..

Figure 8 AN OVERVIEW OF NASA RESEARCH ON POSITIVE DISPLACEMENT GENERAL-AVIATION ENGINES Erwin E. Kempke, J r .

National Aeronautics and Space Administration Lewis Research Center NASA i s involved i n a research and technology program r e l a t e d t o improved and advanced general a v i a t i o n engines. The o v e r a l l goals o f t h e program a r e t o develop t h e technology t o improve f u e l economy, reduce engine weights and i n s t a l l a t i o n drag, and provide f o r brsad-specif i c a t i o n f u e l o r mu1 t i f u e l usage.

I t s two major technic21 t h r u s t are d i r e c t e d a t t h e near-term improvement o f conventional a i r-cooled spark-i gni t i o n p i s t o n engines and a t f u t u r e a1 t e r n a t i v e engine systems based on all-new s p a r k - i g n i t i o n p i s t o n engines, l i g h t w e i g h t diesels, and r o t a r y combustion engines t h a t show p o t e n t i a l f o r meeting program goals i n t h e midterm and long-term future, The conventional p i s t o n engine a c t i v i t i e s i nvol ve e f f o r t s on applying e x i s t i n g technology t o improve f u e l economy, i n v e s t i g a t i o n o f key processes t o permit leaner operation and reduce drag, and t h e development of c o s t e f f e c t i v e technology t o permit f l i g h t at high-a1 t i t u d e s where f u e l economy and s a f e t y a r e improved.

The advanced engine concepts a c t i v i t i e s i n c l ude engine conceptual design studies and enabling technology e f f o r t s on t h e c r i t i c a l o r key technology items.

NEAR-TERM IMPROVEMENT OF CONVENTIONAL ENGINES The o b j e c t i v e of t h e ongoing near-term improvement o f conventional engine e f f o r t i s t o i d e n t i f y and f o s t e r t h e near-term technology base t o reduce fuel consumption by 20 percent, extend t h e c r u i s e a l t i t u d e c a p a b i l i t y and decrease emissions.

The program addresses several s p e c i f i c technology elements through a combination of c o n t r a c t and in-house projects. The most s i g n i f i c a n t are (1) An i n v e s t i g a t i o n o f using e x i s t i n g technology t o modify a p i s t o n engine f o r t h e purpose o f improving f u e l economy and reducing exhaust emissions.

An e f f o r t t o improve c o o l i n g and reduce i n s t a l l a t i o n drag.

( 2 ) An e f f o r t t o improve understanding o f t h e combustion process t o a l l o w ( 3 ) leaner burn operation.

( 4 ) Research t o determine improved f u e l -i n j e c t i o n system c h a r a c t e r i s t i c s .

A contractual e f f o r t t o d e f i n e the b e n e f i t s and optimum design ( 5 ) requirements o f an advanced, c o s t - e f f e c t i v e general-aviation h i g h - a l t i t u d e turbocharger.

Each o f these e:ements w i l l nowQe discussed i n more d e t a i l .

Near-Term M o d i f i c a t i o n s Teledyne Continental Motors A i r c r a f t Product D i v i s i o n , under a NASA c o s t sharing contract, i s researching and developing methods t o improve t h e fuel economy and reduce t h e exhaust emissions o f i t s a i r c r a f t p i s t o n engines. T h e i r research has r e s u l t e d i n t h e development o f f o u r concepts which, when appl i e d t o such an engine, permit leaner operation rlnd thus improved fuel economy and simultaneously reduced exhaust emissions o f hydrocarbons and carbcn mnoxide.

The f3ur chosen concepts as shown i n f i g u r e 1 are (1) A timed, air-density-compensated f u e l - i n j e c t i o n system, which replaces t h e f a m i l i a r low-pressure continuous-flow system.

(2) A thermal b a r r i e r exhaust p o r t l i n e r f o r improved c y l i n d e r head cool ing.

(3) A i r i n j e c t i o n which i n combination w i t h t h e exhaust p o r t 1 i n e r s reduces t h e exhaust valve stem temperatures t o l e v e l s below t h e b a s e l i n e engine w h i l e increasing CG H C o x i d a t i o n i n t h e exhaust.

(4) Variab'e spark t i m i n g t o maintain best power spark t i m i n g over a broader operating range, A comparison of t h e emissions and f u e l economy f o r t h e standard 10-520 engine c o n f i g u r a t i o n and t h e engine w i t h these f o u r concepts i n t e g r a t e d i s presented i n t a b l e I. Emissions are shown as percentages o f t h e proposed 1980 standards. EPA has announced i t s i n t e n t i o n t o withdraw these b u t have not y e t done so. Note t h a t t h e modified engine meets a l l o f t h e standards and i n a d d i t i o n demonstrates a 10 percent improvement i n t h e h i g h performance c r u i s e fuel economy. It a l s o e x h i b i t e d a 22 t o 30 percent improvement i n t h e LTO f u e l economy.

The f i n a l round o f t e s t s are presently being conducted w i t h a f l ightworthy engine, and t h e c o n t r a c t i s being extended t o i n c l u d e an f l i g h t t e s t phase i n e a r l y 1980.

Cool i ng Drag Reduction The o b j e c t i v e o f t h e c o o l i n g drag reduction program, a j o i n t e f f o r t of t h e Lewis and Langley, i s t o develop and demonstrate t h e technology t o improve t h e performance and economy o f p i s t o n engine a i r c r a f t v i a reduced c o o l i n g and i n s t a l l a t i o n drag. Contemporary engine c o o l i n g and i n s t a l l a t i o n designs are based i n p a r t on technology and data developed f o r r a d i a l engines. These data and technology are not adequate f o r precise design o f an engine i n s t a l l a t i o n using a h o r i z o n t a l l y opposed engine f o r which few data on t h e h e a t - r e j e c t i o n patterns are known. I t has been estimated t h a t t h e c o o l i n g drag f o r c u r r e n t designs ranges from 5 t o 27 percent o f t h e t o t a l a i r p l a n e c r u i s e drag. Recently completed t e s t s i n the Ames Research Center 40 by 80 ft wind tunnel have shown a c o o l i n g drag o f 14 percent o f t o t a l a i r p l a n e c r u i s e drag f o r one c o n f i g u r a t i o n .

An i n t e g r a t e d approach t o engine cool i ng t h a t i n c l udes reduced c y l i n d e r cool i n g requirements and improved i n t e r n a l and externpl aerodyna~aics can reduce t h i s drag by a t l e a s t 50 percent.

This c o o l i n g study i s s t r o n g l y encouraged by t h e i n d u s t r y because many This turbocharged i n s t a l 1 a t ions now operate near detonation o r cool i n g 1 i m i ts.

problem, which a f f e c t s both s a f e t y and economy, w i l l be g r e a t l y aggravated by e f f o r t s t o f l y h i g h e r through use o f improved turbochargers.

NASA Lewis w i l l award a c o n t r a c t by t h e end o f 1979 f o r t h e f i r s t p o r t i o n o f t h e c o o l i n g program.

This p a r t o f t h e program w i l l determine t h e c u r r e n t p r a c t i c e and actual minimum cool i ng requirements f o r representative, present-day c y l inder-head and b a r r e l assembl ies. Using these base1 i n e r e s u l t s , various c o o l i n g concepts w i l l be evaluated on t h e i r a b i l i t y t o reduce t h e cool i n q requirements. A t l e a s t two o f t h e most promising concepts w i l l be selected, designed, integrated, and t e s t e d on the c y l inder-head and b a r r e l assembl ies. To extend these experimental r e s u l t s t o other cases, an a n a l y t i c a l computer simulation model on t h e c y l i n d e r head/barrel assemblies w i l l be developed i n a j o i n t e f f o r t o f Lewis and t h e contractor.

A planned b u t as-yet unfunded follow-on contracted e f f o r t w i l l use t h e above improved techno1 ogy base and design i nformation t o design an optimized c y l inder-head and b a r r e l assembly and t o conduct experimental v e r i f i c a t i o n t e s t s on a s i n g l e - c y l i n d e r engine. P a r a l l e l e f f o r t s i n aerodynamics methodology w i l l improve n a c e l l e i n t e r n a l f l o w paths, i n 1 ets, and e x i t s . F o l l owing t h e v e r i f i c a t i o n t e s t i n g , a f u l l - s c a l e engine and n a c e l l e w i l l be b u i l t f o r wind tunnel and f l i g h t t e s t s .

Combustion Process Studies Spark-i gni t i o n combustion process s t u d i e s are being conducted t o provide t h e data base from which p r e d i c t i o n s o f f l o w process and chemical r e a c t i o n s i n homogeneous and s t r a t i f i e d charge engines can be made. Included i n t h i s a c t i v i t y are e f f o r t s i n developing diagnostic instrumentation, conducting experimental studies, and developing sophisticated computer models.

Combustion-Diagnostic Instrumentation Instrumentation has been designed a t Lewis t h a t w i l l determine on a per cycle, per c y l inder basis, r e a l -time measurements of the i n d i c a t e d mean e f f e c t i v e pressure and ~ e r c e n t mass o f charge burned as a f u n c t i o n of crank angle.

Today, our systems are being used by both t h e a i r c r a f t and automotive industry.

W e a l s o have a good design f o r i o n i z a t i o n probes, which are placed i n t h e c y l i n d e r head t o measure flame p o s i t i o n and thickness as a function of crank angle.

Laser Doppler velocimetry (LDV) measurements o f t h e v e l o c i t i e s and turbulence l e v e l s f o r c o l d flow w i t h i n the combustion chamber have been undertaken through a grant t o Carnegie-Me1 1 on U n i v e r s i t y .

A contractual study w i t h Barnes Engineering Co. has indicated t h a t a l a s e r can extend t h e usefulness o f an i n f r a r e d spectral radiometer. The study concluded t h a t spacial r e s o l u t i o n could be improved by using a high-energy l a s e r t o change t h e energy s t a t e o f the specie o f i n t e r e s t a t the measurement point w i t h i n t h e chamber. It can then be detected by t h e i n f r a r e d device a t a different wavelength than the surrounding carbon monoxide i n t h e chamber.

A unique charge sampl i n y system has been developed a t Lewis which measures the local f u e l - a i r r a t i o . The information provided by the system i s used t o establish the c y c l i c and spacial v a r i a t i o n o f f u e l - a i r r a t i o w i t h i n t h e combustion chamber a t selected times i n the cycle o f an operating engine.

B r i e f l y , the sampling system works as follows: A very small volume o f gas i s sampled by a fast a c t i n g valve a t any selected crank angle up t o the s t a r t of combustion. The sample valve ( f i g . Z ) , which was developed from a General Motor's design, i s shown i n s t a l l e d i n the c y l i n d e r head o f a V-8 engine used f o r some o f our combustion studies. The sample enters a high vacuum chcmber and i s analyzed by a mass spectrometer f o r f u e l and a i r concentration. A d i g i t a l e l e c t r o r i c instrument was designed by NASA t o control t h e sample valve, measure the output from the mass spectrometer, and perform t h e c a l c u l a t i o n t o determine the f u e l - a i r r a t i o .

A l l of t h e above instrumentation has proven t o be extremly valuable i n studying the r o l e o f turbulence and gas motions i n combustion chambers. A p r i n c i p a l goal i s t o formulate a general mass o f charge burned equation which includes engine a i r - f u e l r a t i o , speed, a ~ d torque. This r e s u l t i s important i n Otto cycle modeling where u n t i l now the mass f r a c t i o n burned curve was assumed t o have a simple cosine r e l a t i o n .

Otto Cycl e Computer Model For the past 5 years Lewis has had a zero-dimension Otto cycle code e f f o r t .

And since 1977 w e have supported a program on i n t e r n a l combustion engine flame propagation and emissions a t Princton University (and now Carnegie-Me1 l o n ) w i t h Dr. William A. Sirignano as the Principal Investigator. The major goal o f t h e grant program reniains the development of a theory and mu1 tidimension computer code f o r flows and combustion i n a reciprocating engine.

The variables t o be considered are detai 1s o f engine geometry and operating conditions and fuel chemistry. Predictions w i l l include engine performance, f u e l consumption, heat transfer, exhaust compositioc, and l o c a l flow velocities. The code w i 11 also predict the e f f e c t o f turbulence i n t e n s i t y and scale effects, t h e e f f e c t s o f manifold and valve flow, and t h e chemical composition of the flow close t o the wall s. The basic approach i ncl udes not only developing t h e computer program, but also the experimental v a l i l d a t i o n o f key hydrodynamical features of t h e flow model. The status o f t h i s e f f o r t i s as follows: ( I ) The code p r e d i c t s t h e flow f i e l d f ~ r a x i s y n e t r i c , unsteady, moving boundary, compressible, t u r b u l e n t ( s c a l e and i n t e n s i t y ) p i s t o n - c y l i n d e r flows.

( 2 ) The code extends beyond t h e valve i n t o t h e manifold and consequently t h e i n i t i a l c o n d i t i o n s do n o t use assumed v e l o c i t y and turbulence p r o f i l e s but r e a l engine v a r i a b l e s such as valve diameter t o c y l i n d e r bore r a t i o s , valve lift curves, and entrance and exhaust f l o w angles and s w i r l .

(3) The predicted f l o w f i e l d includes t h e boundary l a y e r s and hence surface effects.

s c a l e ( 4 ) The predicted f l o w f i e l d i s s e n s i t i v e t o both l a r g e and small turbulence conditions.

( 5 ) Measurements o f the a x i a l and t a n g e n t i a l components o f mean flow and t h e rms o f v e l o c i t y f l u c t u a t i o n s f o r two d i f f e r e n t operating conditons have been made.

( 6 ) Comparisons w i t h t h e t h e o r e t i c a l p r e d i c t i o n s are being made. The f i r s t examples dre t h e low speed, open o r i f i c e , and t u r b u l e n t nonreacting flows.

The v e l o c i t y p r o f i l e s agree i n shape, and good matching occurs i n t h e v i c i n i t y o f t h e o r i f i c e j e t o r i n t h e r e q i o n near t h e j e t . Comparisons near t h e w a l l agree i n shape, but the predicted v e l o c i t y p r o f i l e gradients a r e l a r g e r than t h e measl,red gradients. Good agreement occurs w i t h t h e l o c a t i o n and magnitude of t h e flow revers21 region.

Fuel I n j e c t i o n Research i s being conducted t o improve the i n l e t - p o r t f u e l - i n j e c t i o n systea To accompl i s h t h i s r 2 q u i r e s a more complete by extending t h e lean 1 i m i t .

understanding o f the re1 a t i o n s h i p o f t h e f u e l - a i r m i x t u r e preparation before The i n d u c t i o n i n t o t h e combust i o n chamber and o v e r a l l engine performance.

extent t o which the microscopic and macroscopic degree o f homogenity o f t h e m i x t u r e e n t e r i n g t h e combustion chamber a f f e c t s performance i s not w e l l known.

The c l a s s i c a l theory concerning m i x t u r e preparatian has been t h a t a w e l l mixed, homogeneous charge was necessary f o r lean operation. Various i n v e s t i g a t i o n s have been conducted which tend t o support t h i s theory. However, these r e s u l t s are i n c o n f l i c t w i t h t h e recent work o f General Motor's researchers who concluded t h a t some form o f heterogeneous i n t a k e charge "wetted" w i t h fuel d r o p l e t s and p o s s i b l y w i t h bulk s t r a t i f i c a t i o n may be optimum f o r lean combustion. Accordingly, t h e NASA i n v e s t i g a t i o n i s designed t o provide a d d i t i o n a l i n f o r m a t i o n on t h i s important aspect without assuming t h a t complete v a p o r i z a t i n y e i l d s optimum heat engine performance.

A l o g i c a l f i r s t step i n t h i s i n v e s t i g a t i o n i s t h e c a r e f u l c h a r a c t e r i z a t i o n of fuel i n j e c t i o n spray nozzles. The physical s t a t e o f t h e f u e l - a i r m i x t u r e inducted i n t o t h e c y l inder i s influenced by t h e p r o p e r t i e s o f t h e spray emitted from the i n j e c t o r . Hence, key v a r i a b l e s such as d r o p l e t size, v e l o c i t y , and s p a t i a l d i s t r i b u t i o n s must be known as a f u n c t i o n o f nozzle desigil and operating parameters.

Spectron Deveiop~nent Laboratories, Inc., under a NASA c o n t r a c t i s usina a l a s e r v i s i b i l i t y method t o o b t a i n p a r t i c l e f i e l d measurements o f d i f f e r e n t i n j e c t o r s under simulated engine manifold conditions.

Concurrently, manifold f l o w - v i s u ~ l i z a t i o n t e s t s t o e s t a b l i sh t h e d i s p o s i t i o n of f u e l - a i r m i x t u r e are being conducted a t Lewis. One c y l i n d e r head of t h e TS10-360 engine has been modified t o i n c l u d e a transparent a c r y l i c i n t a k e s e c t i o n ( f i g . 3). Under a wide range o f motored evgine conditions, high-speed photographs are taken through f i b e r o p t i c s l o c a t e d a t t h e r i g h t s i d e of t h e i n t a k e section. The p r e v i o u s l y discussed sampling valve w i l l be i n s t a l l e d t o measure t h e f u e l - a i r r a t i o w i t h i n t h e c y l i n d e r . From t h i s informatian a c o r r e l a t i o r , o f t h e i n j e c t o r spray and p o s i t i o n w i t h t h e f u e l - a i r m ~ x t u r e i n t h e i n t a k e p o r t and c y l i n d e r can be establ ished.

A f t e r these f l o w v i s u a l i z a t i o n t e s t s , s i n g l e c y l i n d e r hot performance and emission t e s t s w i 11 be conducted.

High A1 t i tude Turbocharger Techno1ogy ?rogram Turbochargers have served General A v i a t i o n we1 1, improving comfort, economy, safety, and performance by p r o v i d i n g h i g h e r t a k e o f f power and high-a1 t i t u d e capabi 1 i t y a t reasonable cost. The t r e n d i s toward even higher a l t i t u d e c a p a b i l i t y f o r a l l s i z e a i r c r a f t engines. Thus, an a i r c r a f t could safely f l y over t h e weather and a t t h e same time improve i t s f l i ~ h t e f f i c i e n c y .

I n response t o a sgggestion by i t s advisory committee, NASA has i n i t i a t e d an e f f o r t t,, i d e n t i f y , develop, and demonstrate t h e technology f o r a Fdmily o f advanced but c o s t - e f f e c t i vc turbochargers appl i c a b l e t o a spectrum of conventional and a1 t e r n a t i v e engines. The projected program, which w i l l emphasize near-term improved s p a r k - i g n i t i o n engines as t h e baseline, has t h r e e phases: system performance analysis and conceptual design r t u d y (phase I ; ; reference c o n f i g u r a t i o n design and component v e r 4 f i c a t i on t e s t i n g (phase i I ) ; and f i n a l design and v e r i f i c a t i o n t e s t i n g (phase 1 I I ) . Phase I has been funded and, depending on t h e r e s u l t s , funds f o r t h e remainder o f ,.., program ill be considered .

ALTERNATIVE PROPULSION SYSTEMS Although c u r r e n t a i r c r a f t engines operate a t h i g h l e v e l s o f e f f i c i e n c y and re1 i a b i l i t y , changins requirements i n terms o f f u e l economy, f u e l avai i abi 1 i t y , and environmental concerns have brought about t h e consideration o f s i g n i f i c a n t l y NAS!. has improved o r completely new types of engines f o r f u t u r e a i r c r a f t .

addressed t h i s issue through a series o f conceptual design study c o n t r a c t s w i t h engine manufdcturers.

The conceptaal e.:gi ne candidates under study are ( I ) improved/ advanced spark-igni t i o n p i stor engines, ( 2 ) 1 i yhtweight d i e s e l engines, (3) s t r a t i f i e d - c h d r g e r a t a r y engines, and ( 4 ) aJvanced turboprop engines which were i n t h e preceeding presentation.

The above-mentioned a c t i v i t i e s t y p i c a l l y i n c l u d e several technical tasks, such as ( 1 \ ::ethnology evaluation and c o n f i g u r a t i o n selecti..n, ( 2 ) conceptual design, ( 3 ) a p r e l iminary airframe i n t e g r a t i o n study, and ( 4 ) proyram recornmendat i ~ n s t.o address each candidate's key techno1 ogy requirements. The status and r e s u l t s of the three i n t e r n a l conibustion engine candidates w i l l be presented f o l l o w i n g t i r i s overview.

The preliminary airframe i n t e g r a t i o n study portions o f t h i s work w i l l assess t h e apparent advantages o f t h a t p a r t i c u l ar concept compared w i t h current engines.

Contractual studies w i l l also be conducted t o obtain from two general- a v i a t i o n airplane manufacturers (Cessna A i r c r a f t Co. and Beech A i r c r a f t Corp. ) a comparative evaluation o f the f o u r candidates i n airplanes and on missions t h a t are representat i ve o f the manufacturer' s expected market share. The eval uat i ons w i l l be conducted on a consistent basis; t h a t is, t o the greatest extent practicable, t h e same o r equivalent airplane and engine technology w i l l be used for a1 1 d i r e c t comparisons, I n a l l cases, the a i r f r a a e w i l l be t a i l o r e d t o take maximum advantage of t h e candidate engine's unique features, o r t o minimize the adverse effect of i t s less-desirable features. The c t i t e r i a o f comparison w i 11 include airplane size, economics, and f i i ght performance; f u e l consunpt i o n and f u e l to1 erance; business factors; and an assessment o f re1 a t i ve technol ogical risks.

I n the above comparative studies two types o f ~ i r c r a f t w i l l be used - a

high-performance, pressurized single and a ~ r e s s t i r i i e d twin. The nominal engine s i z e i s 250 bhp (net, i n s t a l l e d cruise power a t 2 5 000 ft a l t i t u d e ) . It i s expected t h a t r e s u l t s o f these two cofitracts w i l l allow a r a t i o n z l selection o f one o r more candidates f o r a contemplated NASA engine technology development and demonstration program.

Broader in-house studies are being conducted t o evaluate additional types o f a i r c r a f t and missions. Longer t e n supporting research and technol ogy e f f o r t s are being conducted a t Lewis. These e f f o r t s are aimed a t evaluating selected key technology areas i n order t o v e r i f y concept p o t e n t i a l and t o enlarge the o v e r a l l technology base. The diesel and r o t a r y engine t e s t c e l l s a r e now operational, w i t h baseline mapping ccmplete. The buildup of t h e stratified-charge, single-cyl inder f a c i 1 i t y i s under way.

Table I

EMISSIONS AND FUEL ECONOMY COMPARISON

COIUBUUED 8ASWNL MODS EMISSIONS, % €PA 1 m W U S 0 1 s HC 122 11 I BSFC. 5 CHANGE -3Js LTO 0 -1- CRUISE 0

-

INTEGRATED NEAR TERM MODS

IMPROVED CY LlNMR M A D COOLING FIN DE SIGN 7 ' COATtNGS %TAKE V A L E CoMBUSn@N 'VARIABU TIWNG lGNllON S Y S m FOR LIAN FIh MIXTURE OPERARW Figure 1 THE SPARK-IGNITION AIRCRAFT PISTON ENGINE OF THE FUTUREf Kenneth J. Stuckas Aircraft Products Division Tdedym Continental M o t w s A study is unden-ay t o define and apply those areas o f advanced technoi~w appropriate t o t h e design of a spark-ignition a i r c r a f t p5ston engine for t h e lste 1980 t i n e period. !?esults of the study, zo f a r , show that s i g n i f i c a n t iiuprovements i n f u e l econoqy, w e i g h t and s i z e , s a f e t y , r e l i a b i l i t y , d u r a b i l i t y and performance m y be ach5eved v i t h high degree o f success, predicated on t h e continued dei-elqxaent of advances i n combustion systems, electronics, materials and control s y s t e m .

The modern aircraft piston engine has represented t h e b e s t compromise among f u e l econocly, veight, s i z e , cost,ease of maintenance, d u r a b i l i t y and v e r s a t i l i t y . The evolution of t h e a i r c r a f t piston engine over t h e p a s t 50 years has included t h e incorporation of a p ~ r o p r i a t e new technology on a systematic b a s i s , minimizing exposure t o r i s k as t h i s technology became w e l l established an4 pmven i n lower r i s k m i l i t a r y - a d automotive~applications.

A s a result, t h e product which has evolved f ~ o m t h i s process has a demon- s t r a t e d reputation of s a f e t y and r e l i a b i l i t y . Today, t h e spark-ignition f i r c r a f t piston engine serves as a prime mover f g r 93% of t h e nearLv 200,000 a c t i v e a i r c r a f t i n t h e general aviation f l e e t .

I n recent years, t h e r e a l i t y of r i s i n g f u e l prices coupled with t h e p o s s i b i l i t y of reduced f u e l a v a i l a b i l i t y has added impetus t o t h e search for items of advanced technology which, when incorporated i n a newly designed &.reraft piston engine, w i l l continue t o Freserve t h e increasing u t i l i t y of t h i s segment of our U. S. t r a r s p o r t a t i o n system.

AJIVMiCED EIG INE CHRONOLOGY Table 1, shows a chronology of events which we know from errperience must be accomplished before an a i r c r a f t p i s t o n engine cf a t o t a l l y new design *NASA Contract NAS3-21272, Advanced Spark-Ignition Aircraft Piston Engine Design Stukv can e a t e r t h e marketplace.

I n generating t h i s schedule, ve s t a r t e d at Dece!mber 31, 1989 and worked backwards allowing time f o r Marketing and Custoaer Acceptance Testing, Engine a d Airfkame C e r t i f i c a t i o n Testing, Prototype Engine Build-up and Testing and Parts and Materials Rocwement f o r t h e Prototype. A t t h i s point, v e a r r i v e a t t h e time when t h e level of technology t o be included i n t h e engine must be frozen.

This leads t o a sonrevhat discouraging reveiaticm. O w advanced technology engine of t h e late 1980's w i l l reflect a l e v e l of technology t h a t is f i v e and a h a l f years old!

If we begin ts develop t h e technology t h a t ve s e e emerging as cf January 1, 1980, then we have about four and a h a l f years t o develop it t o t h e point where it can be r e a l i s t i c a l l y included i n our a6varrced technology engine.

As p a r t of our study ue took ell those areas of advenced technology w e deemed a ~ p r o l p r i a t e t o an advanced spark-ignition a i r c r a f i piston engine, and put them i n t o categories which v e ranked, komtop t o bottom, i n or&er 09 importance o r dependence, as shovn i n Figure 1.

hr study covered t h e topics of Fuels; combustion systems; various means f o r extracting a d d i t i ~ n a l power from v a s t e exhaust gases - supercharg- ing, turbocompounding and bottoming cycles; engine operational systems such as f u e l iqfection ignition and engine governing systems; cozfiguration and

cooling - shown here on t h e same l e v e l because of t h e i r interwining relation-

s h i p (some engines because of t h e i r configuration require l i q u i d cooling); materials, from t h e standpoint of weight reduction and increased d u r a b i l i t y ;

manufacturing; engine auxiliary systems - such as air conditioning and

e l e c t r i c a l p e r generation m d , f i n a l l y , lubricants.

The -st important decision ve had t o nake, and i n many ways t h e most d i f f i c u l t , was t h e determina+,ion of f u e l a v a i l a b i l i t y . W e did a very thorough l i t e r a t u r e survey covering t h e past, ?resent and f u t u r e of t h e energy industry. W e looked a t not only the teckaicdl aspects of develop- ment of primary energy resources, but a l s o t h e economic, s o c i a l and p o l i t i c a l trends which might a f f e c t our choice of a f u t u r e fuel.

Considering t h e f a c t t h a t t h e U.S. has t h e energy equivalent of 33 times as nuch o i l shale, coal and u r a n i m as t h e r e i s crude o i l i n t h e e n t i r e world, we came t c ? t h e conchsion t h a t petroleum-bas,cod f u e l s would be around f o r e long time t o come t o meet t h e needs o f trahspcrrtation. The assumption being, of course, t h a t t h e s e needs w i l l be m e t by t h e s a t i s f a c t o r y development of the technology necessary t o e f f i c i e n t l y produce s y n t h e t i c crude o i l fron our o i l s h a l e and c c a l resources within economic and environmental constraints, and that- non-transportation needs w i l l be m e t by continued con- serwation measures a n d development cf a l t e r n a t i v e non-petroleum fuels.

- .

This study shaued t h a t t h e r e w e r e two i d e n t i f i a b l e prospects f o r Puel f o r our advanced engrne. F i r s t , for t h e near term, t h e continued use of l O O U avgas is indicated, vhich d i c t a t e s t h e use of a homogeneous charge combustion system similar t o t h a t which is used today, and second, f o r t h e f a r tern, ve s e e the d e s i r a b i l i t y of moving away from t h i s highly specialized aviation gasoline, vhich comprises less than one percent of all t h e gasoline produced i n t h e country today. For t h e f a r tern, o w choice of f;lel is kel-nsene-based commercial jet f u e l , which suggests t h e lrse of a s t r a t i f i e d charge combustion system (Table 2).

A parallel can be drawn between t h e use of J e t f u e l i n an advanced, spark-igniticn z i r c r a f t piston engice and t h e increased ~ r o d u c t i o n of d i e s e l engine powered cars. One of t h e biggest problems associated with t h e ictro- duction of a powerplant desigosd t o o-perate on an a l t e r n a t e f u e l is t h e availa'cility o f t h a t f u e l t o t h e consumer. A s i n t h e case of ',he diesel- powered c a r where diesel fuel is widely available because of t h e e x i s t i n g d i s t r i b u t i o n system f o r long haul trucking, jet f u e l is becoming more videly available due t o t h e increased use of jet-povered busir.ess and commuter airplaues.

'Ihe a v a i l a b i l i t y of tvo f u e l s suggested t h a t our s t u a should address t h e p o s s i b i l i t y C I ' two advanced engines r a t h e r than one. The two engines w e have chosen v e v i l l c a l l moderate r i s k technology and high r i s k technology engines. Both engices are similar, except t h e oder rate r i s k technology engine is Cesigne? t o use l O O U avgas i n a hoaogeneous charge caab*lstion sgsteni and t h e high r i s k technology engize with a s t r a t i f i e d charge combustion system w i l l use jet f w l .

Once t h e matter of f u e l a v a i l a b i l i t y was decided, then the choice of c o b u t i a n systems could be deternineti. S h o k ~ i n Figure 2, on t h e l e f t , is a standard cmbustion chamber usea o r nearly a l l a i r c r a f t piston engines.

The comb-istion chamber volume is hemispherical i n shape, w i t h one intake valve, m e ex9aust valve and two spark plugs per cylinder. O n t h e r i g h t is t h e combustion chamber we are proposing for both t h e moderate r i s k and high r i s k technology engines. I n t h e case of t h e moderate r i s k technology engine, t h e combustion system v i l l use a lov ;;ressure f'uel i n j e c t i o n system where gasoline is indected i n t h e intake manifold J u s t zpstream of t h e intake valve.

f u e l a t The hi&- r i s k technology s t r a t i f i e d charge system w i l l i n j e c t j e t high pressure d i r e c t l y i c t o t h e combustion chwber Just before t h e piston reecnes t o p dead center.

This combustion chmber ve have c a l l e d t h e HTCC, o r high turbulence combustion chamber. Ttlrough t h e use of s w i r l and high turbulence, it permits t h e c a b u s t i o n of lcbn mixtures of fuel and air a t high compression r a t i o s without the detonation vhich limits the campression r a t i o of t h e standard engine. With the HTCC conrbustion chamber, we have recently demonstrated t h e detonation- free operation of a homogeneous charge, 6-cylinder engine a t a compression r a t i o of 12:1 compsred t o a compression r a t i o of 8.5:1 for a stand- ard engine. This increase i n compression r a t i o had t h e effect of impmVin& fuel economy at cruise paxers by 7 percent.

Among the various means of extracting paver from t h e waste exhaust gases of an internal combustion engine a r e turbocharging and turbocompound- F i r s t , ing. I n a n a i r c r a f t piston engine, turbocharging serves two purposes.

it is a means of extracting greater pwer f r o m a given engine displacement, and second it is possible t o maintain tk:at paver from sea level t o high altitudes. Turbocharging is a common practice i n t h e a i r c r a f t piston engine industry. In 1979, about 65% of all a i r c r a f t engines manufactured by Teledyne Continent a1 Motors w i l l be turbocharged.

For our advanced engines ve a r e proposing the use of turbocompounding i n addition t o turbocharging. The schematic i n Figure 3, shows one method of employing turbocoerpounding. The exhaust gases leave the engine, " E ? , and pass first through a power turbine Ti, vhich transmits p e r back into t h e engine crankshaft through a speed reduction unit. The exhaust gases then carry t h e i r remaining energy t o a turbocharger. The advantages of turbo- c olc pounding are that it is possible t o extract one horsepower for every pound of weight added, and the combination of turbocharging plus turbocom- pounding i s more efficient thac turbocharging alone.

Although turbocompounding is not a novel idea i n its application t o a i r c r a f t piston engines, turbocompounding does consititute advanced technology of the basis that we w i l l be atterapting t o apply it t o a n engine of only 350 horsepower, compared t o the 3000 horsepower of the Wright engine and the Napier Nomad of t h e post- World War I1 era.

ENCZNE OPERATIONAL SYS-S W e see for the f u t i r ~ a significant impact on our i n h t r y by t h e work that i s now going on i n the f i e l d of automotive e l e c t m l i c s . There is no doubt that the auto industry represents the greatest potential for f a r term growth for the electronics industry. Partly responsible for t h i s growth i s the development of inexpensive and reliable signal transducers and t h e development of sophisticated electronic control system strategies.

For both the moderate r i s k and hi&% r i s k technology engines w e see the adaptation of all engine operational systems t o electronic control.

This rne;~ns that the present three levers now i n use t o control engine speed, m a n i f ~ l d pressure and fuel flow w i l l be combined i n t o a single lever by which the pilot controls power, The t r i c k is t o be able t o accolpplish t h i s task so thak the systems exhibit fail-soft behavior. This lsesns t h a t a mechanical backup aptem v i l l be required.

I n vicv of t h e increasing coarplexity of our air t r a f f i c control system and the increasing cuaount of single-pilot IFR W a g , t h e extent t:, uhich ve can reduce p i l o t workload impinges directly on safety of flight. T a b l e 3 lists those operational systems uhich w i l l be converted t o electronic control and the benefits derived fram each.

We examined many different engine configurations and reduced our choices to t h e three shown i n Figure 4. A s f a r as cooling is concerned, our conclusions f o r these three configurations and an engine of 350 horsepower, vas t h a t liquid cooling provides no distinct advantage i n either weight o r cost over air cooling. In fact, uhen considering the added systems required for liquid cooling, a certain additional risk is involved, Since all engines are ultiaately air-cooled, and because of t h e lav temperature differentials present vith liquid cooling, t h e placement of a radiator large enough t o remnre t h e rejected heat would pose a problem i n t h e already compact design of an airplane f o r vhich t h i s s i z e engine is intended.

Only one of these configurations looked promising compared t o t h e horizontally-opposed, six-cylinder design w e ultimately chose, and that w a s the inverted V-8. The V-8 engine would be =re vibration-free than the horizontally-opposed six, but from a cost a d maintainability standpoint, six-cylinders are preferable. The radial design was rejected because of its large frontal area.

ADVMCEf) MATERIALS The use of advanced materials was considered from the standpoint of weight reduction and increased durability. Table 4 compares three engines it contains. The where the velght of each is divided up among the materials f i r s t engine is a TSIO-550 engine representing the present level of technology.

It contains 8 lbs. of miscellaneous materials such as plastic, rubber and copper, 332 l b s of s t e e l , and 245 lbs. of aluminum, for a t o t a l weight of 585 lbs. Our moderate r i s k technology engine contains only 253 Ibs. of s t e e l and

215 lbs . of aluminum while we have added 10 lbs . of advanced materials tor

a t o t a l weight of 485 lbs., which is a weight reduction of 17%over the present engine, The reduction i n use of s t e e l and aluminum i n t h i s engine is brought about primarily by the more judicious use of these materials.

In the high risk technology engine we =e w i n g only 80 lbs. of s t e e l , reduction gears, cylinders and exhaust valves.

primarily i n the crankshaft, The use of aluminum has been reduced somewhat and a t o t a l of 119 lbs. of

advanced materials are used for an engine weight of 40 5 lbs . , a 31% improve-

ment over t h e present dey engine.

I n t h i s engine the greatest part o f t h c advanced material weight is titanium, with a small m u n t of reinforced plastic and ceramics.

Titanium is one of the most abundant metals t o be found on earth.

While titanium is not a rare metal, it is very costly t o produce. The problem being that it i s not usually found i n great quantity i n one spot, but it's pretty much evenly distributed over t h e earth. Another problem is t h a t it takes 1 3 times as much energy t o produce a pound of titanium iroln ore as it does t o make a pound of s t e e l . What we are counting on here i s an advancement i n titanium production and metallurgy which would permit an overall savings i n energy consumption t o be realized. The question i s whether t h e fuel saved by reducing t h e weight o f t h e engine by 31% will be overcme by the energy used t o produce t h e titanium i n t h e first place.

ENGINE SPECIPICATIOlO Many of t h e d e t a i l s of t h e design study have been omitted f o r t h e sake of brevity, but Table 5 shows comparison of some of t h e specifications of t h e three engines we have discussed. All three are six-cylinder, horizontally- opposed. The current technology engine has a displacement of 550 cubic inches, with 420 cubic inches f o r the moderate and high r i s k technology designs. All three a r e rated a t 350 HIP and can cruise at 25,000 feet at 250 BHP. A t t h i s cruise parer the brake specific fuel consumptions a r e ,446, .358 and .331, The service ceilings of our advanced engines a r e increased t o respectively.

32,000 feet compared t o 35,000 feet for the present engine.

We've already discussed t h e installed weight and type of fuel.

The TBO, or time between overhaul, for our current technology engine i s 1400 hours, which w e have increased t o 2000 hours for t h e advanced engine.

To get an idea of the f u e l econow improvements which were made, compare The current technology t h e power wasted i n t h e exhaust of the three engines.

engine dumps the equivalent of 3l9 HP out the exhaust a t maximum cruise power.

For the moderate r i s k technology engine t h i s loss has been reduced by 33% t o 214 HP, and by 51% t o only 156 HP for the high r i s k technology engine.

IMPROVED AIRPLAlUE PERFORMANCE Well, what does all t h i s bw- us? It's not enough t o consider only W e must look a t the bottom line. That t h e improvements i n the engine.

i s , what benefits do we see when the engine i s installed i n an airplane?

But here are W e are not quite finished with t h i s part of our study.

some preliminary results based on the installation of the three engines i n a current technology single-engine airframe.

What we've done here i n Figure 5 is t o simulate t h e installation of all three engines i n an identical air- designed f o r a chosen arbitrary mission profile f o r t h e high r i s k technology engine. W e chose a range of 1000 nautical miles v i t h 45 dnutes fbel i n reserve, a cruise a l t i t u d e of 25,000 feet at 250 horsepwer vhich corresponds t o maximum cruise power for all three engines, The resulting calculgtions shaw t h a t t h e present technology engine vould have a raage of only 518 nautical miles and t h e mo&erate risk technology engine, 814 nautical miles, A relative efficiency was calculated f o r all three cases based on the payload each airplane could carry, multiplied by its speed and divided by its fuel consuenptim. The efficiency factor w a s then normalized t o t h e value of 1.00 for t h e present technology engine. Based on this factor, t h e relative efficiency of t h e moderate r i s k technology engine v a s increased by 32% and t h a t of the h i e r i s k technology engine by 49%.

O f course, these factors vill change depending upon t h e mission profile selected.

A similar analysis w a s done f o r the case of atwin-errgine airplane (Figure 6), with similar results.

In t h i s case t h e mission profile was set a t 1300 nautical miles for t h e high r i s k technology twin.

The results show normalized relative efficiencies of 1.00, 1-37 acd 1.57, respectively.

Figure 7 shows the top, side and r e a r external viers of an advanced engine, pointing out some of the important features. Both t h e moderate risk and high r i s k technology engines w i l l appear sq&stantially the same, externally.

Compared t o present engines, t h e gear-driven propeller shaft has been extended somewhat t o accommodate a more streamlined cwling. The exhaust system is designed for good pulse recovery t o enhance t h e paver recovery of t h e turbocompounding power turbine. The speed reduction system From t h e power tcrbine t o t h e crankshaft is a lasvytis traction drive which w a s chosen over a gear reduction drive because of its potential for damping torsional vibrations and its lighter weight. The engine also includes an integral o i l sump/oil cooler t o save weight and volume, In order t o achieve a compact design, the exhaust system is on t o p of the engine rather than on the bottom. Because of t h i s , the engine is designed for updraft cooling, instead of the usual downdraft method. This, i n conjunction w i t h an advanced airframe design and well-designed baffling w i l l permit the design of a more efficient r a m a i r pressure rise recovery plenum.

While t h i s study has not been completed, ve hsvt identified several items which require development i n t h e next four and a half ye- i n order for the proposed engines t o become a r e a l i t y by t h e end of t h e next decade, as outlined i n Table 6.

The c r i t i c a l developmen+, items we have identified include; s t r a t i f i e d charge ETCC conibustfon system and a compatible advanced ignition system; an improved efficiency, high pressure ratio, lightweight turbocharger; a reduction drive system f o r t h e turbocompounding paver turbine; electronic control strategies appropriate t o a turbocompounded a i r c r a f t piston engine, and a method t o improve engine cooling and reduce cooling drag.

Other items of a non-critical nature which have been identified include the reduction of engine friction, t h e lov cost production of titanium, the development of lightweight accessories and impr ~ v e d heat exchangers for o i l cooling and induction air intercooling.

ADVANCED SPARK-IGNITION AIRCRAFT PISTON E W E TECHMOLOGY BASE CHRONOLMY [OPTIMISTIC) PROTOTYPE ENGINE IUILO.UP YARKETING AN0 CUSTWEP.

ACCEPTANCE TESTING

FUELS F OR ADVANCED SPARK-IGNITION

AlRCUAF T PISTON ENGINE

N E A R T E R M 100 LL A V G A S O l i WIDE.CUT V E R S I O N 'HOMOGENEOUS CHARGE COMBUSTION) FAR TERM

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a KEROSENE BASE COMMERCIAL JET F U E L (JET A ) ( S T R A T I F I E D CHARGE COMBUSTION) TABLE 2 MODERATE RISK HIGH R I , K JMOMOGENEOUS CHARGEl jSTRATlFlE0 CHARGE) ELECTRONIC FUEL CONTROL ELECTRONIC L I R CONTROL ELIMINATES MANUAL MIXTURE CONTROL WOVIOES AIR THROTTLIYB FOR OPTlMUY FUEL ECONOMY A REOUCES U L O T WORKLOAD ELECTRONIC 'GNITION A PROVIDES O ? l l M U M FUEL ECONOMY COUPLES IGNITION TIMING WITH FUEL 4 PREVENlS ENGINE DAMAGE DUE TO IWECTION FOR OPTIMUM COMlUSTlON IWROPER MIXTURE CONTROL TECHNIOUES ELECTRONIC SINGLE .LEVER M W L R CONTROL ELECTRONIC SINGLE-LEVER WWER CONTROL CAME AS MODERATE RD;O A ELIMINATES SEPARATE THROTTLE (RACK) A N 0 PROP CONTROLS A REDUCES R L O T WORKLOAO PROVIDES onmuw ENGINE SPEEDS AND THROTTLE (RACK) SETTINOS TABLE 3 ADVANCED MATERIALS FOR ENGINE WEIGHT REDUCTION TOTAL MISCELLANEOUS AOVANCEO ENGINE E R C E N T (36) MATERIALS STEEL ALUMINUM MATERI&LS* WEIGHT WEIGHT 0% R E O U t n O N Bb! (Ib) Ubl I I U

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PRESEWT 350 hp I 132 245 - H 6 fl

t S I O - S U ENGINE MODERATE RISK I 2~ ns 10 u s 17 TECH#OLOGY ENGINE HIGH RISI( 8 k 210 118 105 31 TECHNOLOGY ENGINE *TITANIUM, CARBONIGRAPMITE/BORON REINFOT.:EO PLASTICS. CERAMICS TABLE 4 W A X € PEClFlC FUEL -1 AT U X # Y tkuJsc rarROlW =*I)? a h mar 051, i n n r m m e ADVANCED SPARK.IGNITION AIRCRAFT PISTON DESIGN STUDY TECHNOLOGY PRGGRAM RECOMMENDATIDNS fiEDilCED ENGINE F R l C T t O N STRATIF !ED CHARGE HTCC COM8LISTIG"i SYSTE.2 AT)\ A'ICED *G\'TlO\ SYSSEV LW COST PRODL'C'1O"c OF T I T A V I U K YPPCbE3 EFF'CiEYCY HIGH P R E S S Y E RAT:O ilGrT.%E:SHT ACCESS0P:ES -1GHT #.EIGHT T~RBOCIAQGFP IMPROVED HEAT EXCHAkGERS 8i3:'CTION DRIVE S Y C T I Y AND CLUTCH FOR T U R B O C0MPOUND:;di TURBINE ELECTROVIC CONTROL STRATEGIES A?PROPR!ATE TO TURBOCOUMPOUNaEL AIRCRAFT PISTON ENGINES IMPROVED EFJGlhlE C C Z L I N G A h C COOLING DRAG REDUCTlOk TABLE 6 ADVAllCED TECHWOLOGY CATEGORIES H I E W I I I C A L STRUCTURE I E W H E AUXiLlARl S m H16H CWPRESSION RAllO/UAW BURR COYBUSTION CHAMER HTCC STANDARD 8.5: 1 COYIRESSlOIl RATIO RADIAL

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RAMBE WITH 45 MIMUTE RESERVE 7 1 0 a mi L: LIGHTWEIGHT DIESEL AIRCRAFT ENGINES FOR GENERAL AVIATION

Steven C . Berenyi and Alex P. Brouwers

General Products Division Teledyne Contimtal Motors This design study reintroduces the d i e s e l engine as an a i r c r a f t powerplant.

A methodical design study was conducted t o a r r i v e a t new d i e s e l engine con- figurations and applicable advanced technologies.

ILo engines a r e discussed and t h e description of each engine includes concept dravings. A performance analysis, stress and weight prediction, and a c o s t study were a l s o conducted.

This information w a s then applied t o two airplane concepts, a six-place M a and a four-place s i n g l e engine a i r c r a f t . The a i r c r a f t study consisted of in- s t a l l a t i o n drawings, computer generated performance data, a i r c r a f t operating costs and drawings of the r e s u l t i n g airplanes.

The performance d a t a s h w s a vast improvement over current gasoline-pwered a i r c r a f t . A t the ctnnpletion of t h i s basic study, the program was expanded t o evaluate a t h i r d engine configu- ration. This third engine i ~ c o r p o r a t e s the b e s t features of the o r i g i n a l two, Preliminary information on t h i s engine and its design is currently i n progress.

is presented.

INTRODUCTION Encrgy conservation, uncertainties of f u e l supply and limited a v a i l a b i l i t y of high octane gasoline, have renewed the i n t e r e s t i n the d i e s e l a i r c r a f t engine, since its f u e l economy is b e t t e r than any type of a i r c r a f t engine currently i n product ion.

Aircraft d i e s e l engines have been developed before, notably the Junkers "JUMO", O f these, only t h e the Napier "NOMAD" and the McCulloch TRAD 4180.

Junkers opposed piston, 2-stroke cycle engine ever reached the production stage.

Its complexity The Napier Nomad was a 2-stroke cycle, turbocompounded design.

and the f a c t t h a t it invaded the t e r r i t o r y of turbine engines probably accounted for its demise. The McCulloch engine came close t o f l y i n g when the program was terminated for non-technical reasons.

New technologies, now under a c t i v e development, w i l l r e s u l t i n even b e t t e r f u e l economies than can be obtained with current state-of-the-art d i e s e l engines.

These technologies a l s o make it possible t o develop a powerplant which is more compact and l i g h t e r than current gasoline a i r c r a f t engines.

Tuo engines were investigated i n the study, a 298 kW (400 HP) d i e s e l f o r a twin engined airplane and a 149 kW (200 HP) d i e s e l f o r a s i n g l e engined a i r c r a f t .

\ C The study consisted of three major phases: 1 . Technology Analysis.

A survey of available aviation and automotive sources was conducted to identify new developments which offer potential benefits to an aircraft engine. These technologies were ranked and then evaluated on the basis of performance and adaptability.

2 . Engine Concept Design.

The technologies which were chosen as a result of the evaluation and ranking process were applied to the design of the 149 and 298 kW engines. Performance, stress, weight, and cost calculations were made concurrently.

3 . EnginefAircraf t Integration Study .

The results of Step 2 were then used in an engine-aircraft integration study to determine the performance improvement of an airplane equipped with these diesel engines.

Some of the technologies which were applied in these engine designs are anticipated to be available in the late 1980's.

These technologies result in high level of performance and, although advanced, are not untried. The adiabatic engine, the catalytic combustor, and the high speed alternator envisioned are currently under development under various contracts. It should be noted here that, although the concept engine proposes the use of ceramic combustion system components, the use of such materials for "man-rated" aircraft may be 20 years These concepts were included primarily to shov what may be ultimately away.

possiblz. However, alternate less advanced solutions are also given which will result in a small reduction of performance when compared to the ultimate; but nevertheless will result in a powerplant which far outperforms the current gasoline aircraft .

RESULTS AND DISCUSSION Advantages of the Diesel Engine The diesel engine has always been burdened with the stigma of being heavy, thus offsetting its advantage of low fuel consumption for aircraft applications.

If it is possible to build an engine that combines low fuel consumption and low weight, then that engine becomes a very attractive aircraft powerplant. Old and once discarded concepts can become attractive by applying new technologies.

A conventional diesel er~gine requires high compression ratios for starting This results in high firing pressures at full load when and low load operation.

in fact the engine could run adequately at a much lower compression ratios.

New technologies make i t possible t o combine good s t a - r t a b i l i t y with lov f i r i n g pressures at f u l l load. This study shows t h a t the weight of t h e d i e s e l can be reduced below t h a t of current gasoline a i r c r a f t engines.

The d i e s e l engine o f f e r s other advantages i n addition t o low f u e l consumption, i.e. : 1. Lower operating cost.

- Lover c o s t of f u e l

- Reduced maintenance

- Extended TBO

2. Greatly reduced f i r e and explosion hazard.

3. Better in-flight r e l i a b i l i t y . N o i g n i t i o n and mixture control problems.

4. Multi-fuel capability.

5. N o carburetor icing problems.

6. Improved a l t i t u d e performance.

7. Safe cabin heating from exhaust s t a c k s ( l e s s danger of carbon monoxide) .

8. Exact f u e l metering indicator. The rack position determines t h e f u e l flow.

9. N o e l e c t r i c a l interference from i g n i t i o n system Previous A i r c r a f t Diesel Engines Table I shows a l i s t i n g and design data of past a i r c r a f t d i e s e l engines.

N o c l e a r trends follow from t h i s tabulation. Seven of the t h i r t e e n engines have a r a d i a l configuration, seven w e r e aircooled, eight were 2-stroke cycle.

The tabulation becomes more meaningful i f specific. r a t i o s a r e used. See Table 11.

Some observations can be made from these tables. Average s p e c i f i c weight values are: 4-Stroke cycle engines .710 kW/Kg 2-Stroke cycle engines .926 kW/Kg Aircooled engines .783 kW/Kg Liquid-cooled engines .923 kW/Kg 'fie numbers i n d i c a t e t h a t a 2-stroke cycle engine can be expected t o be l i g h t e r than a 4-stroke cycle engine. A comparison of aircooled and l i q u i d cooled engines would seem t o favor t h e liquid-cooled engine. However, t h e engine weights of liquid-cooled engines tabulated do not include t h e weight of the cooling package. With t h i s modification, the corrected values then becose: Aircooled engines .78 3 kW/Kg Liquid-cooled engines ,805 kW/Kg New Engine Design Study

- 298 kW Engine -

In addition to considering the historical background of aircraft diesel engines, a literature search was made and a technology base was established to evalute any concepts that may be considered for an all new engine design.

The follwing criteria were observed in considering any new ideas: The engine must be a pistonfcrankshaft type powerplant.

Be compatible with conventionally designed aircraft (size and arag).

nilow manufacture of an experimental model in five years.

Be ready for production in the late 1980's.

M ?P+ 1979 EPA Emission Standards (guide-ref erence only) .

Have multi-fuel capability.

Have engine performance comparable to current aircraft engine.

Have lower BSFC than present engines.

Maximum specific weights of .852 KgIkW for the 298 kW and 1.095 Kg/kW for the 149 kW engine.

Life cycle costs equal or less than present aircraft engines.

Avoid problem areas encountered in current aircraft engine designs.

Design attributes that had to be considered included: Performance, weight, size, C. G . , fuel economy, reliability, multi-fuel capability, noise, life cycle cost, component costs, and technology required.

Figure 1 is a flow chart that shows the different possible combinations of design features that were considered and evaluated within the frame work outlined above.

A detailed description of each of these features is beyond the scope of this paper, but was included in the original basic study. The evaluation then result- ed in the technologies that follow from Figure 1 by taking the high score items along the "common to all versions" and "radial 2-stroke cycle" lines.

Common to All Radial 2-S troke Versions Line Cycle Line Open Chamber Individual Cylinders Ceramic Pistons Low Compression Ratio Insulated Exhaust Manifolds Geared Prop Drive No Cylinder Cooling !.oop Scavenge Tool Steel Piston Rings Independent Turbo Loop Composite Connecting Rods Catalytic Combustor Synthetic Lube Oil High Pressure Fuel Injection Electronic Controls Conventional Oil Filter Conventional Fuel F i l t e r Pendulum Damper The engine concept was then l a i d out around these features.

Figure 2 s h w s an a r t i s t rendering of the proposed 298 k W 6-cylinder engine.

Figure 3 shows t h e schematic of t h i s engine. The design incorporates t h e technologies which were defined before.

The chosen system uses the C u r t i s loop scavenging. The i n t a k e p o r t s and intake manifold a r e located at t h e propeller s i d e of the engine, exhaust p o r t s and exhaust manifold a t the back end.

The cylinder l i n e r and piston top are ceramics and, therefore, cylinder cooling w i l l not be required. Tool steel p i s t o n r i n g s w i l l be required. Cool- ing air w i l l be used only f o r t h e a f t e r c o o l e r , o i l cooler, and t h e f u e l i n j e c t o r s .

The exhaust p o r t s w i l l be o i l cooled.

Each cylinder receives f u e l from a separate i n j e c t i o n pump located i n f r o n t of t h e cylinder (cool s i d e of t h e engine). F a i l u r e of one pump still leaves 516 of engine paver available. A high i n j e c t i o n l i n e pressure w i l l be required t o limit i n j e c t i o n duration a t high engine speeds.

The turbocharger can run independent of t h e engine. For t h a t pilrpose a high speed s t a r t e r / a l t e r n a t o r and an o i l pump a r e mounted on t h e turbocharger. A two-way valve is placed i n t h e intake manifold. To s t a r t t h e engine, t h i s valve is i n t h e v e r t i c a l position of t h e schematic, which r e s u l t s i n a turbocharger loop independent of t h e engine. Combustor f u e l is ignited by t h e heater. This heater can be turned off a s soon a s the c a t a l y s t becomes s u f f i c i e n t l y hot. The cycle w i l l become self-sustaining a t approximately 1 / 3 of t h e maximum turbo speed, and t h e s t a r t e r now runs a s an a l t e r n a t o r . Hot, high pressure a i r w i l l flow t o t h e engine when t h e two-way valve is partia1l.y opened. The cylinder intake p o r t s a r e opened during approximately 120 craxrk-degrees, so hot a i r can flaw through two cylinders f o r preheating on cold days. The high pressure a i r w i l l next be admitted t o the engine mounted bleed a i r s t a r t e r t o crank t h e engine. The whole sequence would be automatic.

This system o f f e r s many advantages: 1. The a v a i l a b i l i t y of hot induction a i r a t s t a r t reduces the need f o r a high compression r a t i o . The engine w i l l s t a r t and i d l e a t a 10:l compression r a t i o provided t h i s hot, high pressure a i r is a v a i l a b l e t o i t during cranking. Thus, with t h i s low compression r a t i o , the f i r i n g pressures a r e held down t o 9650 kPa (1400 psig) a t f u l l load r e s u l t i n g i n low engine weight.

2. The engine w i l l s t a r t e a s i l y under cold conditions, a problem with current engines.

3. Hot s t a r t problems a r e eliminated.

4.

The engine can be -hut-off and t h e turbocharger kept running when t h e a i r c r a f t i s on the ground f o r some period. Meanwhile, e l e c t r i c p w e r , cabin heat o r a i r conditioning remain available.

This i n e f f e c t converts t h e turbocharger i n t o an APU.

5 . The b a t t e r y requirement is g r e a t l y reduced s i n c e engine cranking is accomplished by a i r pressure.

The use of synthetic o i l is required i n t h i s engine design due t o the hot cylinders. The synthetic o i l can take higher temperatures and requires fever changes than conventional petroleum bases o i l s .

Over t h e long term perhaps a method can be found t o generate an a i r f i l m between pistons and cylinder walls, i n e f f e c t , using a i r bearing technology. This would a l s o reduce the expected r e l a t i v e l y high o i l consumption which is inherent t o 2-stroke cycle engines.

The engine design concept is shown i n Figures 4 through 8. The cylinders a r e arranged i n two o f f s e t banks of three cylinders each, a c t i n g on a s i n g l e crankpin.

The r o t a t i n g and reciprocating i n e r t i a s are 100% balanced by coumer- weights on the crank cheeks.

The pendulum dampers a r e mounted t o the counter- weights and w i l l be tuned f o r t h e 4-1/2 and 6th orders. The cylinders a r e un- cooled and provided with ceramic l i n e r s .

The intake ports and the intake mi- fold a r e located a t t h e f r o n t s i d e - the cool s i d e of t h e engine. The exhaust p o r t s and exhaust manifolds a r e located a t t h e backside - t h e hot s i d e of the engine. Two exhaust manifolds a r e required t o prevent the exhaust pulse of one -vlinder t o i n t e r f e r e with t h e scavenging of the previous cylinder i n t h e f i r i n g s tlence. The piston tops a r e ceramic.

The s n a l l end of t h e connecting rods is designed t o allow f r e e r o t a t i o n of the piston. This should reduce t h e wear r a t e of the piston rings. The b i g end of t h e connecting rods is designed a s a s l i p p e r , i.e., each rod contacts only 113 of the circumference of t h e crankpin. This is possible f o r 2-stroke cycle engines because the combined load of gas pressure and i n e r t i a s is always d i r e c t - ed toward the crankpin. The bearing material w i l l i n i t i a l l y be conventional, but a study could be conducted l a t e r of s e l f - l u b r i c a t i n g and gas bearings t o eliminate the need f o ~ o i l i n t h e crankcase.

Immediately i n f r o n t of t h e f i r s t main bearing a r e 6 individual i n j e c t i o n Individual pumps were chosen t o pumps, operated by a s i n g l e lobed cam ring.

improve engine r e a l i a b i l i t y - f a i l u r e of one pump still leave 5 cylinders

operable. Also, a l l f u e l l i n e s can have the same length r e s u l t i n g i n t h e same i n j e c t i o n timing f o r a l l cylinders.

A bevel gear i n f r o n t of t h e cam ring drives the prop governor and t h e f u e l priming pump.

-

- .

- . .

A gesr reduction reduces t h e cranksh=ft speed of 3500 rpm a t take-off down

t o 2300 rpm propeller speed.

A t the back of the crankcase i s an accessory housing which contains t h e gearing f o r t h e engine o i l pump, t h e vacuum pump, and the bleed a i r s t a r t e r .

The a i r s t c r t e r drive is provided with a s l i p clutch t c ?revent engine damage i n the case of a h y d r o s t ~ t i c lock i n one of the cylinders (accumulation of f u e l Four engine mounting points a r e provid- due t o t h e leakage of a f u e l i n j e c t o r ) .

Above t h e accessory case is t h e c a t a l y t i c combustor ed on the accessory housing.

Leading t o it a r e t h e two exhaust manifolds and the a i r bypass f o r assembly.

operation i n t h e APU mode.

Figure 8 shows The turbocharger is located behind t h e accessory housing.

To the r i g h t is a t h e turbine t o the l e f t and the compressor i n t h e center.

gear housing with the high speed a l t e r n a t o r and turbo o i l pump drives.

The a f t e r c o o l e r and o i l cooler a r e located below t h e engine accessories.

The engine w i l l operate with a dry sump.

The operating parameters f o r t h e 298 kW engine a r e shown i n Table 111 and t h e sea l e v e l performance curve is included a s Figure 9.

In addition t o t h e performance projections, d e t a i l e d calculations were made of weights, t o r s i o n a l vibrations, power component stresses, turbomachinery sizing, cooling requirements, and projected c o s t s t o manufacture.

Results of each of these s t u d i e s are very favorable f o r the d i e s e l engine, however, these d e t a i l s a r e beyond t h e scope of t h i s presentation.

Emissions were not q u a n t i t a t i v e l y addressed, however, the following q u a l i t a t i v e statements a r e valid: 1. Hydrocarbons and carbon monoxide w i l l be oxidized by the use of a c a t a l y t i c converter.

2. NOx concentration w i l l be minimized due t o the r e l a t i v e l y low peak pressures (9650 kPa) and lower peak temperatures.

3. Smoke l e v e l s should be r e l a t i v e l y low s i n c e the minimum trapped A/F r a t i o w i l l always be on the order of a t l e a s t 24:l.

A s with emissions, only q u a l i t a t i v e evaluations were made of the anticipated engine noise a s l i s t e d below: 1. The c a t a l y t i c combustor and insulated exhaust stacks i n s e r i e s with the turbocharger should minimize d i r e c t combustion noise.

2. The absence of cylinder cooling f i n s should reduce externally generated vibratory noise.

3. The absence of valves, rocker arms, push rods, and camshaft should minimize i n t e r n a l l y generated mechanical noise.

4. The geared drive w i l l allow a r e l a t i v e l y low propeller speed, thereby, reducing prop generated noise.

5. Two-stroke cycle operation, however, tends t o o f f s e t some of the gains noted above.

A s was described e a r l i e r , t h i s engine's f e a s i b i l i t y relies heavily on nev technology. Following are t h e areas where e x i s t i n g technologies need t o be advanced t o make such an engine feasible: 1, Piston rings - operating i n uncooled cylinders.

2. Cylinders - ceramic components and t h e i r I n t e r f a c e with m e t a l l i c

hardware.

3. Turbo s t a r t e r / a l t e r n a t o r operating a t high speeds.

4. C a t a l y t i c combustor and its associated controls.

5. Cooling of the cylinder exhaust ports.

6. Piston lubrication.

7. Spherical connecting rod end.

8. E f f i c i e n t f u e l i n j e c t i o n systems.

A comparison of t h e 298 kW engine was made with t h e 4-stroke cycle GTSIO-520-H gasoline engine.

Table I V shows t h i s comparison i n a tabular form.

Figure 10 is a s i z e comparison.

The f r o n t a l area of t h e d i e s e l ensbl 78% of t h a t of a comparable gasoline engine.

- 149 k W Engine Design -

The technologies epplied t o t h e 149 kV engine a r e not a s f a r advanced a s i n the case of the 298 kW engine.

The 149 kW engine w i l l primarily serve t h e p r i v a t e owner market where i n i t i a l cost and e a s t of maintenance carry more weight than i n the case of the corporate a i r c r a f t .

The engine w i l l be e a s i e r t o develop and manufacture.

Figure 11 shows an a r t i s t rendering of t h e p.:oposed engine.

Figure 12 shows t h e schematic of the engine.

The following f e a t u r e s a r e incor,orated i n t h e 149 ! 7 design concept: 1. Radial configuration.

2. Two-stroke cycle C u r t i s luop scavenging.

3. Minimum cylinder cooling - reduced f i n area.

4. Variable compression r a t i o pistons (VCR) .

5. Mechanically driven c e n t r i f u g a l blower, declutched when not needed.

6. G ' o w plug s t a r t i n g a i d i n cylinders.

7. Conventional s t a r t e r and a l t e r n a t o r .

8. Conventional exhaust system (no combustor) .

9. Direct propeller drive.

Calculations of t h e heat t r a n s f e r through cylinder walls, as well a s s i n g l e cylinder engine t e s t s have conf!nned t h a t the heat f l u x is highest through the cylinder walls surrounding the comkustion chamber (when the piston is i n top The maximin gas temperature t o vhich t h e cylinder v a l l is l o c a l l y dead center).

exposed drops off f a s t a s the p i s t o n t r a v e l s dounrard, r e s u l t i n g i n a lower l o c a l average cycle gas temperature and, therefore, r reduced heat flwt. It can be safeiy predicted t h a t most cooling f i n s below t h e p i s t o n r i n g b o l t (piston i n TDC) can be eliminated vithout an appreciable e f f e c t on cylinder w a l l , piston and piston r i n g temperatures. Using t h i s approach r e s u l t s i n an increase of cooling drag vhen compared t o uncooled cylinders; but elidnates t h e need for ceramic :orponents, thus making the engine a n u & . more v i a b l e a l t e r n a t i v e f o r nearer cem applications.

S i w e t k i s smaller engine does not havz t h e independent turbocharger loop, Other means must be found t o low coapression r a t i o pistons cannot be u t i l i z e d .

It becomes necessary t o reduce the keep f i r i n g pressures d m t o 9,650 kPa.

compression r a t i o under load t o 10:l.

However, t h e engine cannot be s t a r t e d o r run i d l e a t such a low compression r a t i o . I n t h e case of t h e larger 298 kW engine, t h i s was solved by m a n s of the inaependent turbocharger loop which provides intake a i r of s u f f i c i e n t pressure and temperature t o start the engine and the c a t a l y t i c combustor vhich keeps t h e turbocharger a t a high speed during sugine i d l e operation. This is not t h e case here, therefore f o r t h i s case a variable compression r a t i o piston is rc:ommended.

The VCR piston, Figure 1 3 v a r i e s t h e compression r a t t o from 17:l at start and lw load t o 10:l a t f u l l load. This high C . R . is s u f f i c i e n t under normal ambient conditions t o s t a r t t h e engine. Even t h e 17:l compression r a t i o , h w - ev-i, does not provide a s u f f i c i e n t l y high compression temperature t o i g n i re t h e f u e l a t very low ambient temperatures.

Operation of t h e glow plug may be re- quired t o assure good s t a r t a b i l i t y . It is a l s o intended t h a t g l w plug oper- a t i o n would automatically be i n e f f e c t a t lw t h r o t t l e s e t t i n g s . This would be an added safety feature t o assure absolutely no misfiring during descent mode operation.

Scavenging of a 2-stroke cycle cylinder requires t h a t t h e intake manifold pressure exceeds the exhaust manifold pressure at any load and engine speed.

The turbocharger, however, produces a negative a P a t l c r w load. This is no

problem for 4-stroke cycle engines where t h e piston does t h e scavenging. The 2-stroke cycle engine without a combustor requires an engine drxven blower t o

produce a p o s i t i v e A P across t h e cylinders a t low loads. The blower w i l l be

disconnected a t the load point where the turbocharger provides a p o s i t i v e n P .

It a l s o 3ecame obvious e a r l y i n the design phase of t h e 149 kW engine t h a t a d i r e c t drive would r e s u l t i n a smaller engine paclcage and a weight reduction.

The e ~ g i n e r e l i a b i l i t y is somewhat improved by t h i s appru.-.ch due t o fewer p a r t s required.

The chosen BWEP of approximately 1200 Wa 2s 100 kPa higher than t h e BMEP of t h e larger 298 kU engine.

The much lower crankshaft speed d i c t a t e d by t h e d i r e c t propeller drive w i l l r e s u l t in b e t t e r scavenging and, hence, a l a r g e r

amount of a i r trapped i n the cylinde .. It should, therefore, be possible t o

obtain :.iis higher BMEP without a- i n c e a s e of cylinder temperatures. The detailed cycle calculation^ bear t h i s out.

'fie 143 kU engic;e concept design then is shown i n t h e Figures 14 thrcugh 18.

The cylinders a r e arrangdd i n one bank of four cylinders. The r o t a t i n g and reciprocating i n e r t i a s a r e 100% balanced. 'Ihe cylinders have a liptited number of cooling f i n s t o cool t h e c o d u s t i o n chamber. The necessity f o r a gear driven blower a t the back s i d e of t h e engine made it more p r a c t i c a l t o have t h e cylinder intake port a t t h e back s i d e and the exhaust manif@lds a t t h e front. The exhaust manifolds w i l i be insulated t o avoid radiation t o the i n j e c t i o n pupps.

TWO exhaust manifolds a r e required t o avoid pulse interference between cylinders.

The connecting rods a r e t h e s l i p p e r type. The b i g weds are wider than i n t h e case of the 298 %!r' engine t o coapensate f o r t h e reduced circmaferential coatact length.

The use of synthetic o i l is not e s s e n t i a l i n t h i s engine because of lower cylinder temperatures (compared t o the ceramic), but may be advantageous t o extend the periods betveen o i l changes.

Four individual i n j e c t i o n pumps a r e provided driven off a single fabe cam ring, The c e n t r i f u g a l blower is driven off t h e propeller shaft through a lay s h a f t which is located above t h e crankcase betveen t h e cylinders d l and 14.

This arrangwent was chosen r a t h e r than a d r i v e f r a m tile rear end t o avoid torsional problems. The nodal point lies c l o s e t o t h e l a r g e s t i n e r t i a member of the crankshaft system, t h a t is the propeller. Putting t h e blower d r i v e gear nepr t h i s point reduces t h e input of torsional amplitudes i n t o the blower drive.

The lay s h a f t , which is a q u i l l s h a f t , further i s o l a t e s che blower from t h e crank- s h a f t vibrations. Houever, t h i s feature forced t h e use of a d i r e c t propaller drive. Ta put a propeller reduction gearing i n f r o n t of the blower d r i v e vould have led t o an unacceptable length of the engine. A weight a n a l y s i s f o r t h i s particular engine shoved t h a t the d i r e c t d r i v e with the inherent larger p i s t o n displacement still r e s u l t s i n a l i g h t e r engine than the geared drive.

The blower d r i v e is provided with two clutches. One, t h e magnetic clutch, disengages the blover d r i v e once t h e t u r t o c b r g e r has c o w up t o speed, The location of t h e magnetic clutch is such t h a t =s much of blower d r i v e a s possible is disengaged t o prevent unnecessary drag on tile engine. A d i s c type s l i p clutch is provided t o prevent large t o r s i o n a l amplitudes a s they occur a t lw engine speeds due t o c y c l i c i r r e g u l a r i t y from reaching t h e blower.

The turbocharger is mounted behind the engine, a s a r e t h e o i l cooler and the aftercooler. ZLo versions of t h e engine w e r e drawtr. One, a s shown, f c r a n a i r c r a f t v i t h fixed landing gear. A second versiim of t h e engine was d r a m which accommodates a r e t r a c t a b l e nose gear. The coolers a r e moved outboard and the turbocharger raised t o provide space between cylinders #2 and 83 f o r t h e nose gear s t r u t .

Table V presents the operatinp parameters of the 149 kV engine concept and Figure 19 shows th:. pro2ected f u e l consrrmption curve for t h i s engine.

A s with the 298 kW engine, calculations were made t o define power component s t r e s s e s , turbocharger and cooler s i z i n g , t o r s i o n a l vibration d e f i n i t i o n , as well a s c o s t and t light projections.

Table V I and Figure 20 aze presented t o show comparisons v i t h today's state- of-the-art gasoline engines.

a11 comparisons favor t h e d i e s e l Again, a s v t t h t h e l a r g e r engine, ewine.

EngineIAirframe I n t e g r a t i o n This study was conducted as a subcontract by Beech A i r c r a f t Corp. t o e v a l u s t e t h e i n t e g r a t i o n of the proposed d i e s e l a i r c r a f t engines I n t o f u t u r e airframes and t o detenaine t h e e f z e c t of t h e engiae on a i r c r a f t perforaaace m d operating c o s t s . The r e s u l t s were then compared with corresponding d a t a f o r c u r r e n t p r d u c t i o n type gasoline engine powered a i r c r a f t .

Engine I n s t a l l a t i o n I n s t a l l a t i o n design layouts were made which show t h e 298 kk' d i e s e l mounted on a w i n engine a i r p l a n e and the 169 kg engine i n s t a l l e d i n a s i n g l e engin* a i r c r a f t with r e t r a c t a b l e landfng gear. The Figures 21 through 23 show t h e twin engine i n s t a l l a t i o n ; t h e Figures 2 1 through 26 show t h e s i n g l e engine install- a t i o n . Figure 27 and 28, t h e t h r e e view drawings, are based on t h e wing a r e a s indicated by t h e performarace s y n t h e s i s program, t h e engine drawings and standard a i r p l a n e proportions. Some p e r t i n e n t f e a t u r e s about t h e i n s t a l l a t i o n a r e as follous:

1. Engine m u n t s a r e of two b a s i c types - c a n t i l e v e r and bed mount. A

e a n f i l e v e r mount from the f i r e v a l l was used i n t h e twin and a bed mount incorporating the nose gear support s t r u c t u r e was used i n the single. "Dynafocal" type Eocnts woull be used with the c a n t i l e v e r method t o minimize v i b r a t i o n transmission t o the airframe.

1. The induction system i n both cases would be a SACA f l u s h i n l e t , duct- ing and an a i r f i l t e r . Alternate a i r would be a v a i l a b l e t o t h e engine through a door operated by d i f f e r e n t i a l pressure.

3.

Both engi.nes have a dry o i l sump and require e x t e r n a l o i l tanks mounted i n the engine cmpartments.

I. Both engines would have cooling a i r i n l e t s p-oviding a i r t o a plenum chamber. Ducts from the plenure would d i r e c t a i r t o individual c y l i n d e r s , o i l coolers. a f t e r c o o l e r s and f u e l i n j e c t o r s a s needed.

On the s i n g l e .

c o d i n g s i r e x i t s a r e outboard of the nose gear on the lower s i d e of the cowling. Exits from the twin n a c e l l e would be a t t h e laver a f t end.

5. The i n s t a l l a t i o n drawings were done i n enough d e t a i l t o i n d i c a t e t h e f e a t - , e s noted above and t o provide reasonable assurance t h a t no major i n s t a l l a t i o n problems would be encountered with t h e proposed d i e s e l engine concepts.

A i r c r a f t Configurations Three v i e w sketches of the airplane are shown i n t h e Figures 27 and 28.

Following a r c some c h a r a c t e r i s t i c s of both planes: 1. Propeller Data.

Prop. diameter 2.057 n Prop. speed a t take-of f 2,345 rpm Tip speed a t take-off 253 mr'sec = .74a a = Velocity of sound = 20.06 m m / s e c (T i n OK) A t standard aPlbient temp. 15.50C a = 20.06 v m 5 . 5 = 341 d s e c Prop. speed a t economy c r u i s e 1,790 r p m Tip speed a t ecommy c r u i s e 193 d s e c

Prc? . ground clearance 330 nr

2. Sight Angles.

The p i l o t ' s s i g h t angles f o r t h e tvia are indicated by A and B (Figure 27).

The c e n t e r l i n e angle over t h e nose, A, as i3dicated is about 120. If t h e a i r p l a n e were l o f t e d , t h e angle from t h e p i l o t ' s a c t u a l eye p o s i t i o n vould be about 1 8 O which is considered more than zdequate. The srallest l a t e r a l angle E is 100.

This is a l s o more than adequate e s p e c i a l l y c o q a r e d t o sorae current pis-on engine twins with l a r g e r nacelles.

3. A i r c r a f t Data.

Twin Engine Twin Engine Diesel Gasoline Airfraae minus engine (a) kg 1,860 1,860 Engines (2)

Empty weight (a) + (b)

Payload Fuel load

Useful load (d) + (e)

Max. tzke-of f weight ( c ) Wing span Length T a i l height T a i l span Wing area Figure 28 shows the s i n g l e engine a i r c r a f t . C h a r a c t e r i s t i c s are: 1. Propeller Data.

Prop. diameter 2.134 m Prop. speed a t take-off 2,400 r p m Tip speed a t take-off 268 m/sec = .79a prop speed a t economy c r u i s e 1,800 rpm Tip speed at ecaooly cruise Prop ground clearawe 2 . The c e n t e r l i n e angle over t h e nose f o r t h e s i n g l e engine a i r p l a n e C, is This should correspond t o z c t u a l p i l o t ' s viewing angle of about 12O.

go.

This is probably adequate, e s p s c i a l l y vhen compared t o some of today's long nose s i n g l e engine a i r c r a f t .

3. A i r c r a f t Data.

Single mine Single E a g h D i e s e l Gasoline Airframe minus engine 667 (a) ke. 667

Engine -

4 t y weight (a) + (b)

Payload Fuel load Useful load (d) + (e) Max. take-of f veight ( c ) Wing span Leagth T a i l height T a i l span Wing area A i r c r a f t Performance Evaluation The m j o r t o o l used i n t h e a i r p l a n e design synthesis was a sa-t modified version of t h e synthesis method o r i g i n a l l y developed f o r t h e NASA =TE (General Aviation Turbine Engine) Study. The process - w a s simplified f o r t h i s purpose s i n c e taks-off and c r u i s e pover could be specified a s program inputs. The program is not accurate enough nor does i t account f o r enough v a r i a b l e s t o a c t u a l l y design airplanes, but it is considered adequate t o i n d i c a t e trends i n r e l a t i v e s i z e and performance f o r airplanes t h e o r e t i c a l l y equipped with s u f f i c i e n t engines. The w i n point t o bear i n mind when looking at t h e r e s u l t s of the program is t h a t t h e objective is t o provide an indication of t h e differences i n perforaance and c o s t b e t w e n d i e s e l and gasoline powered airplanes. The methods used i n estimating throughout are no b e t t e r than 5 t o 1CZ accurate, but the uniform assumptions and metilods used i n a l l cases vould make the r e s u l t i n g differences good indications of t h e trends t o be expected.

This is t h e proper objective f o r a conceptual investigation.

Hypothetical gasoline and d i e s e l powered a i r p l a n e s were synthesized and compared i n two ways. I n one case, t h e airframe was held constant and t h e mission p r o f i l e was allowed t o change when t h e powerplant type changed. I n the o t h e r case, the mission requirements w e r e held constant and the a i r p l a n e needed t o perform t h a t mission changed s i z e a s necessary t o meet the mission requirements.

These comparisons were made f o r both the s i n g l e 149 kW 3nd t h e twin 298 kU engine air?lanes.

The r e s u l t s of t h e a i r c r a f t performance simlatioa program are shaun in the Tables F f I and VIII.

Table VII shovs t h e differences in a i r c r a f t p e r f o m e f o r a fixed airplrrw site.

The fixed parameters are:

- b i . take-off weight

- Max. land- veight

- Take-off distance

- Landing distance

- S t a l l speed

- Wing area

The advantages of the d i e s e l s with t h e i r high c r u i s e power output sad lw fuel ccmsu~prioas can be r e a d i l y seen in the basic parareters of range, speed, and payload.

The advantages of t h e d i e s e l s with t h e i r high c r u i s e power output and l w f u e l c o n s ~ t i o n s can be reatilly seen i n t h e basic p a r a r e t e r s of range, speed, and payload.

Table VIII shows t h e dizferences i n a i r p l a n e s i z e f o r a fixed performance.

The fixed parameters are:

- Paylead

- Xax. c r u i s e speed

- Range

The gasoline pwered a i r p l a n e s a r e bigger and considerably less e f f i c i e n t .

Operating Cost Csi'mates Production c o s t s were estimated by assuming t h a t nev a i r p l a n e s would be designed and equipped with the d i e s e l engines and, a l t e r n a t i v e l y , coapatible gasoline engines. Development, material, ad labor c o s t s w e r e chosen t o be of roughly the correct magnitude, but a r e intended primarily t o i l l u s t r a t e c o s t differences due t o using d i e s e l rnstead of gasoline engines. Operating c o s t estimates were made using figures obtained from current estimates of average operating costs.

The acquisitton cost estimates were based on information from the a i r p l a n e synthesis process. The a i r p l a n e empty weights vere t h e main parameters used v i t h FY79 r a t e s f o r labor, material costs, and OPI engine costs. The estimating methods used a r e based on h i s t o r i c a l data and "learning curve" theory. An a i r - frame weight was estimated from the operating empty weight. This w a s used with estimating d a t a t o get material weights t o which material c o s t could be applied.

Manhnur per pound data were used t o get labor content t o which labor rates were applied. A productten run of 600 units was used.to amortize a s s 4 development costs and to locate facrors on the learning curves. When a basic factory cost was swmned up, assumed ~anufacturer's and dealer's mark-ups vere applied.

Costs were included for currently typical optional equipment a d avionics selections.

The final total represented a dealer's price tag figure for a typically equipped airplane. Both the single and the twin were considered to be all neu designs.

The same sets of reasonably realistic assumptions vere used throughout so the results are quite adequate for looking at differences between gasoline and diesel airplane prices within the overall accuracy of this study. Acquisition price percentage changes from the diesel to the gasoline engine pavered airplanes is shown on the cost srpaaries. See Tables IX and X for the twin and single engine airplanes, respectively.

The columns headed "gasoline" refer to the airplanes for equivalent size to the diesels but with the iaission capability as indicated in the performance estimates. The "equal plane performance gasoline" column refers to the airplanes that will do the same missions as the diesels but are bigger and less efficient.

The cost suaary tables s h w the considerable overall cost advantages of the diesel powered airplanes.

Gasoline airplanes of equivalent size cost less initially but this advantage is not sufficient in view of the reduced mission capability and higher overall costs. The biggest factors in raising the gasoline airplanes operating costs are fuel and overhaul expense, as indicated.

Propeller Noise Estimates Propeller performance estimates were made to get some idea of the propeller sizes needed to realize a cruise propulsive efficiency of -85 for both the twin and single engine airplanes. These calculations indicated that a two-blade, 84 inch diameter, constant speed propeller will vurk for the single engine airplane.

The propellers indicated for the twin are 81 inch three-blade. Estimates of 1000 ft. flyover noise predict valstes of 72 dB(A) for the single and 74 OB(A) for the twin. These compare favorably to the limits of 77.5 dB(A), respectively.

Limits are based on airplane weight as set out in FAR 36, Appendix F. A favorable carrection factor can reasonably be expected, creating a greater margin relative to the limits.

The correcticn factor is based on detailed take-off performance estimates that are beyond the scope of this st-ady. Even without correction factors, the noise regulations appezr to present no problem for the conceptual diesel airplanes.

186 kW (250 HP) Engine Configuration -9 third engine configuration is currently under evaluation and design definition. This engins is rated at 186 kW (250 UP) net shaft power at 25.000 ft. cruise altitude. Table XI outlines the pertinent features of this engine configuration. Engine specifications are summarized in Table XII.

As shown in t h i s table, a cruise fuel consumptioa value of .36 lb/hp-ht i.

projected f o r t h i s engine. Some pertinent data comparing t h i s projected fuel consmption with other engines is shown in Table X I I I . Mote tbat the projected v a : -. is relatively conservative compared t o actual n u m h g engines, c m f S ~ the tact that an engine such as proposed is within reach.

A study indicates that the diesel engine praises to be a superior pouer- plant for general aviation a i r c r a f t for the follaving reasons: 1. ?he diesel engine offers high cruise power a t a l t i t u d e and l o w fuel scmsumption. This w i l l result in improved range, high cruialng speed and more payload for a diesel englned a i r c r a f t .

2. The diesel powered airplane has a considerable overall cost advantage.

Gasoline airplanes of equivalent s i z e cost less initiaLly, but t h i s advantage is offset by reduced mission capability and higber operatlag costs.

3. The diesel engine presents no installation p r o b l w . Altbough tbe radial configuration is different than current gasollne engines, the mounting t o the airframe is essentially tbe sare and requires no major a i r f rare modifications.

5 . The independent turbo loop provides:

- Easy cold and hot s t a r t s

- Can crank engine indefinitely

- Electric power available independent of engine operation (APU wde)

- Reduced battery capacity

- Cabin cooling or heating available while a i r c r a f t is on the ground

6. The r a d i a l cylinder configuration results Ln:

- Low engine weight

- Reduced engine f r i c t i o n

- Absence of piston i n e r t i a forcaa

- Compactness of the power package

7. The two-stroke cycle feature results in:

- Weight reduction

- Improved r e l i a b i l i t y due to fewer parts

- Reduced frontal area

8 . The following key technologies will be required to demonstrate the feasibility of the engines proposed in this study:

- Coaabustionlscavenging in 2-cycle loop scavenged system

- High pressure ratio, high efficiency turbocharger

- High pressure fuel injection system

- High speed starter/alternator .

1985-2000 (400 HP Engine)

- All of the above

- Ceramic components

- Advanced labricants (solids, air bearings, etc .)

- Catalytic combustor

TABLE I Pnrkus Aircmft D h d s N o . k n s l r o L . D l r O L ~ .

Ylb ro61 Conlig.CldrCodlngCIL- - -

Rdw *

1. PwLVd 8ir 2. Guiblrson R.dw liquid 30.A S(kschmpQ Radii1 4. efistd pho.nix air 5. LbmlovLI LOD ma .h 6. H i m CNWt l4F2 Rldial air 7. Ynr#w, SH18 Rdw 8il OF2 60- v IiQuid 8. MOWOOOS leuid 9. Junksn 204 oRw=d rquid 10. Judms 205 mw=d liquid 11. Junkers (1)' 207 Turbo opposed liquid 12. N.pier 0 . Nomd Flat 13. McCulbch (3)' TRAMlaO Radial air 'Nurnbsrr in parentheses refer to list of references at the end of this rapwt.

TABLE 11 Specific Data of Previous Aircraft Diesels Pman PWonHw( w e Pactcard Guibewn Deschamps Bnstol Zbropvka Hlspano Salmson Merccdcs Junkers 204 Junkers 205 hnkers 207 Nap~er McCulloch TABLE Ill

Operating Panmeters - 2911 kW E n g h

Altitude 0 6.096 meters 6.096 Power 298 298 194 kW RPM 3 . 5 0 0 3,500 2,675 Displacement 4 . 7 1 4 . 7 1 4 . 7 1 liters Bore x Stroke 100 x 100 100 x 100 100 x 100 mm BMEP 1,085 1,085 923 kPa Compressor Pressure Ratio 4 . 0 6 : 1 8.30: 1 6 . 2 5 : l Nominal Compression Ratio 1 3 . 1 8 5 : l 1 3 . 1 8 5 : 1 1 3 . 1 8 5 : 1 Effective Compression Ratio 1 0 . 0 : 1 1 0 . 0 : 1 1 O . O : l Barometric Pressure 1 0 1 . 4 46.4 6 6 . 4 kPa Ambient Temperature 1 5 . 5 -25 -25 'C lntake Manifcld Pressure 402.4 370.2 277.6 kPa lntake Mnifold temperature 116 116 116 'C Exhaust Manifold Pressure 309.5 284.8 245.5 kPa Scavenge System Curtis Loop Curtis Loop Curtis Loop Sca~enge Ratio 1 . 5 1 . 3 1 . 3 Ratio BoosUBack Pressure 1 . 3 1 . 3 1 . 1 3 1 He~ght lntake Ports 20.65 20.65 2 0 . 6 5 mm Height Exhaust Ports 26.14 26.14 mm 26.14 61' 4 7 ' lntake Ports OpenlClose 6 1 ' 4 7 ' 6 1 ' 4 7 ' BBDCIABDC Exhaust Ports OpenlClose 6 9 ' 3 9 ' 69-39' 6 9 ' 3 9 ' BBDCIABDC BSFCsngine 212.9 1 9 4 . 6 gkW-hr.

206.8 BSFCcombustor 1 8 . 2 6 . 1 0 glkW-hr.

BSFC-powerpack 225.0 219.0 194.6 glkW-hr.

Fuel Flow Powerpack 6 7 . 1 65.3 37.8 kglhr Air Density .00279 .00256 ,00205 kg/f AirIFuel Ratio 27.50 24.59 25.47 TABLE I'J Comparison of GTSIO-520-H Gasoline and GTDR-290 Aircraft Diesel Engine 2.slmk* C* &stroll* c m GTSK)-SZDH 0TDRJ)O

- Engicr

-- * - --E!

6 cyl. opposed 6 cyl. radial Configuration 4.71 Displacement .! 8.52 Take-off RPM 3400 3500 Rated max. take-off power kW 280 298 Rated max. for cruising kW 210 298 Prop speed at take-off RPM 2278 2345 BSFC glkW-hr: Take-off 100?/0 power cruise 65% power cruise Dimensions: Length mm V1:dth mm Keight mn, Er,gtne weight dry, kg TABLE V E n g h Opofoting Pwrmohn r o o v r k r r 1 - clr*.

Altitude 0 3,048 Power 149 119 RPM 2100 2400 Displacement 3.14 3 . 1 4 Bore x Stroke 100x 100 ' 0 0 x 100 BMEP 1,187 1 . 1 8 7 4 . 1 6 : l 6 . 1 0 : l Compressor Pressure Ratio Compression h t i o variable Variable Max. C.R. 173 (effective) 1&1 (eftective) Min. C.R.

Barometric Pressure 1 0 1 A 69.6 69.6 kPa

Ambient Temperature 1 5 . 5 - 5 -5 'C

411.8 411.8 lntake Manifold Pressure 200.9 kPa lntake Manifold temperature 116 r 16 1 1 6 ' C Exhaust Manifold Pressure 316.8 316.8 255.4 kPa Scavenge System Curtis Loop Curtis Loop Curtis Loop Scab~~rge Ratio 1 . 3 1 . 3 1 . 3 1 . 3 1 . 3 1 . 1 Ratio BoosUBackpressure 2 0 . 1 3 mm Height Intake Ports 2 0 . 1 3 2 0 . 1 3 Height Exhaust Ports 27.15 27.15 2 7 . 1 5 mm

* 6 1 ' *6l0 a 6 1 BBDClABOC

lntake Ports OpenlClose Exhaust Ports OpenlClose t 7 1 ' a 7 1 ' 7 BBDClABDC 222.0 2 2 8 . 1 209.8 glkW-hr.

BSFC Fuel Flow 20.3 kglhr.

3 3 . 1 34.0 AirIFuel Ratio 26.6 26.0 2 4 . 0 TABLE VI Comprrlron d TSIO-3o.E Gasoline and TOR-192 Aircnft D M Engine 2afeka C* -cFI.

tSKIJ(M TDIEln2 -EnOln -Engh.

Configuration 6 cyl. opposed 4 cyl. radial Displacement 1 5 . 9 1 3.14 Takeoff RPM 2800 Rated m u . takeoff power kW 149 149 Rated max. for cruising kW 112 149 Prop drive direct direct BSFC glkW-hr: Takeoff 3 7 7 . 1

100% power cruise -

65% power cruise 267.6 Dimensions: Length mm 1188 Width mm 795 Height mm 672 Engine weight dry, kg 174.6 TABLE VII Comparison Gasoline and Diesel Aircraft Engines Airplane Size Fixed, Variable Performance Twin-Englm Twin E n d n Gasolirw' - -:- ---- - -. - - - -.

Rated power kWlRPM 14912400 29812300 (ea) 29812267 (ea) Max. take-off weight (gross) kg 1349 3654 Max. landing weight kg 1349 3654 3654 Standard empty weight kg 829 2275 2385 Useful load kg 520 1378 1269 Usable fuel 2511180 9081653 8321598 Payload (with full fuel) kg 340 726 671 Altitude -ml".b power 30481 100 ',o 304e175% 7620181.5% 7620175% Ma,. crurse speed kmlhr 324 291 474 448 Range km 1481 1468 2592 1726 Altttude - rnl0,~ power 304817540 30481 75 % 7620/81.5% 7620175% Speed kmlhr 289 291 474 448 Range km 1 968 1468 2592 1 726 Take-off dtstance (normal. OV. 15 m) n 579 579 70 1 701 Landlng dis:ance (normal. OV 15 m) m 369 369 677 677 Stall speed (land~ng) kmlhr 85 85 135 135 m : Wlng area 17.7 17 7 22.6 22.4 'All englnes are turbocharged w- krtglma Gaalm-

- - --

+ l l ' r ine Configuration Features ADVANCED RCTARY ENCINE STUDIES Charles Janes Curtiss-Wright Corporation INTROWCTION The i n t e n t of t h i s paper is t o review recent Rotary Engine Developments r e l e v a n t t o a S t r a t i f i e d Charge Rotary A i r c r a f t Engine. I n addition, present s t a t u s of t h e NASA-funded Advanced A i r c r a f t Engine Study, which is c u r r e n t l y unden-ay, w i l l be b r i e f l y described.

Background Work Although Cur tiss-Wright designed t h e i r f i r s t Wankel-type Rotary Engine i n 1958 and ran t h i s engine i n e a r l y 1959, developments continued i n t o 1962 before a r e l i a b l e , durable and e f f i c i e n t b a s e l i n e engine was demonstrated, The f i r s t S t r a t i f i e d Charge trials were made t h a t same year, d i r e c t e d towards s multi-fuel m i l i t a r y engine. During t h e mid-601s period, two proto- type S t r a t i f i e d Charge Rotary Engines were designed, b u i l t and developed through t h e e a r l y operational test stand s t a g e (ref. 1). The RC2-60U10 ( f i g u r e

1) w a s 3 liquid-cooled two r o t o r vehicular engine i n t h e 160 - 200 HP class

and the RC2-90 ( f i g u r e 2) an a i r cooled 300 HP h e l i c o p t e r drone engiite. The trochoid dimensions of these engines was the same as t h e 1958-designed 60 cubic inch s i n g l e r o t o r engine (the RC1-60), but t h e r o t o r width was increased 50% f o r the RC2-90. Both engines proved t h e i r multi-fuel c a p a b i l i t i e s , but n e i t h e r could match t h e f u e l economy of our carbureted RC2-60U5 automotive prototype engine of t h e same e r a , which was comparable t o e x i s t i n g automotive engines (ref. 2). Furthermore, the RC2-60U10 performed w e l l a t the lower powers and speeds, with shortcomings apparent a t the other end of the operating regime, whereas the 90 cubic inch combustion configuration was subsequently developed t o meet high power goals, only t o show low end d e f i c i e n c i e s . I n both cases, however, t h e engines showed s u f f i c i e n t technical promise f o r t h e i r s p e c i f i c designed a p p l i c a t i o n s , but a s a r e s u l t of changes i n m i l i t a r y planning, the intended uses did not m a t e r i a l i z e and development was shelved.

Although thermal e f f i c i e n c y equal t o our homogeneous charge Rotaries was never demonstrated with these engines, t h e inherent compatibility of the Rotary geometry with unthrottled and d i r e c t chamber i n j e c t e d S t r a t i f i e d Charge combus- t i o n ied some t o believe t h a t t h e p o t e n t i a l f o r s u p e r i o r performance had t o be there. Figure 3 i l l u s t r a t e s how t h e Rotary provides the required r e p e t i t i v e scheduled turbulence, without l o s s e s , while d i r e c t chamber i n j e c t e d recipro- cating engines have t o generate t h e required v e l o c i t y g r a d i e n t s a t a c o s t of both volumetric and mechanical efficiency, f u r t h e r widening the s p e c i f i c power advantage of the Rotary.

Following the f u e l c r i s e s of 1973, R&D e f f o r t s were r e i n i t i a t e d in an attempt t o resolve whether o r not t h i s higher efficiency p o t e n t i a l r e a l l y ex- isted. This time, our f e a s i b i l i t y t r i a l s were directed towards automotive appli- cations which peaat not only vide power and speed range f l e x i b i l i t y with f u e l economy, but lw o r controllable m i s s i o n s as w e l l . Since hydrocarbon emissions a t the very lw speeds and povers typical of an automotive operating regime had proved t h e -st d i f f i c u l t area for the homogeneous charge Rotary, new configura- tions were screened on the b a s i s of road-load brake s p e c i f i c f u e l consumption (BSFC) and brake s p e c i f i c hydrocarbons (BSHC). The 1973 attempt t o combine the best f e a t u r e s of RC2-60U10 and f i n a l RC2-90 i n j e c t i o d i g n i t ion designs i n t o a s i n g l e configuration which could run f u l l range was successful and, f o r the f i r s t time, achieved b e t t e r f u e l consumption, on a variety of fuels, than the gasoline carbureted engine. This design improvement led to, i n 1974, a -re f l e x i b l e arrangement whereby a separate p i l o t nozzle, with r e l a t i v e l y small fuel f l w , is used t o t r i g g e r combustion. This design, shown i n f i g u r e 4, uses a multi-hole main nozzle, located close t o the trochoid surface t o modulate f u e l f l w i n response t o power demand.

A number of variations of t h i s basic approach were tested during the 1975 and 1976 periods of increased R&D a c t i v i t y and the r e s u l t s showed that the localized and controlled combustion could produce low "raw" hydrocarbons. The test findings did indicate, however, t h a t increased rotor combustion pocket temperatures were required. In t h i s case, these temperatures were achieved by use of an air-gap insulated surface p l a t e attached t o the rotor combustion face.

The r e s u l t s , f o r two successive 8.5:i compression r a t i o hot r o t o r designs (figure 5) show t h a t the best of these w a s able t o match the shaded area which represents modern automotive engine untreated HC levels. These data a l s o i l l u s - t r a t e t h a t the r e s u l t s were similar f o r the d i f f e r e n t f u e l s tested.

Although BSFC vs. BMEP a l s o showed r e l a t i v e l y small differences with these f u e l s there was no s i g n i f i c a n t reduction of BSFC with the increased rotor temperatures used t o reduce BSHC.

This e a r l y t e s t work indicated that further HC reductions a r e possible with moderate intake t h r o t t l i n g a t the very low power/speed end of the regime and by an increase of compression r a t i o . Accordingly, the f i r s t t e s t on the RC1-60 s i n c e i n t e r r u p t i o n of the automotive-directed a c t i v i t y between e a r l y 1977 and the present, is now being run with a 10:l compression r a t i o rotor. Since t h i s compression r a t i o increase w i l l a l s o improve SFC, it is germane to look a t the comparative trends. The test evaluation, which s t a r t e d i n October, has not yet completely covered the nozzle-matching and i n j e c t i o n dynamics sorting out pro- cess. Preliminary data, presented on an Indicated b a s i s i n figure 6, shows some promise; however, gain on a brake basis w i l l be s l i g h t l y less as a r e s u l t of some f r i c t i o n increase.

The 1976 s p e c i f i c fuel consumption baseline curves (8.5:l compression r a t i o ) a r e shown i n f i g u r e s 7 and 8. Figure 7 compares r e s u l t s , f o r the same designs t h a t demonstrated low hydrocarbons, t o data which a r e representative for f u l l - sized European automobiles powered by Diesel engines. Figure 8 adds other speeds and compares VW sub-compact "Rabbit" Diesel 4 cylinder engine d a t a de- veloped f o r D O T ( r e f s . 3-5). The r e l a t i v e s i z e s and weights of t h e canparable output Volkswagen Diesel 6 cylinder engine and a S t r a t i f i e d Charge Rotary sine a r e s h u n i n f i g u r e 9.

The "cast i r o n r o t o r housingr' curve i l l u s t r a t e s the type of SFC improve- ment t h a t w a s a t t a i n e d with a moderate i n c r e a s e of trochoid s u r f a c e temperature.

While a c a s t i r o n r o t o r housing w a s used f o r t h i s test exploration, t h e tern- peratures t e s t e d do not preclude use of aluminum. Further work is required t o d e f i n e gain at higher levels.

It follows t h a t i f one can match t h e s w i r l o r pre-chamber d i e s e l on an en- gine-for-engine f u e l consumption b a s i s , then t h e smaller dimensions and reduced bulk has t o mean b e t t e r t o t a l v e h i c l e system f u e l efficiency. Furthermore, it is s i g n i f i c a n t t o note t h a t Texaco has developed data (refs, 6, 7) t o show t h a t t h e United S t a t e s would be able t o obtain m r e usable Btu's per barrel of crude o i l i f the r e f i n e r i e s were optimized t o produce a broad base middle d i s t i l l a t e f u e l .

Testing of Other S i z e s The combustion e f f i c i e n c y shown f o r the automotive size6 module is of i n t e r e s t f o r o t h e r a p p l i c a t i o n s only t o the extent t h a t the same technology can be scaled t o the s i z e s required f o r the p a r t i c u l a r application. The s c a l i n g f l e x i b i l i t y of the homogeneous charge engine has been demonstrated adequately over a per r o t o r displacement range of about 500:l and 1 to 4 r o t o r s b u t u n t i l 1978, S t r a t i f i e d Charge Rotaries with the c u r r e n t full-range design f e a t u r e s had not been run i n any other s i z e . me e a r l i e r configurations ( f i g u r e 2) had been run i n t h e wider r o t o r 90'13 chamber and shown t h e same thermal e f f i c i e n c y (ISFC) a s t h e RC1-60 ( r e f . 8).

In e a r l y 1977 the RC1-60 t e s t i n g program was deferred f o r Engineering ac- t i v i t y on a l a r g e r 350 cubic inch module. The 350 cubic inches per r o t o r was achieved by enlarging the trochoid by approximately two-thirds and widening r o t o r proportions by 25 percent.

The same technology and basic configurations developed i n the RC1-60 were used f o r the 350 cubic inch engine, including a "reversed" configuration ( A X p i l o t ) where t h e p i l o t and main nozzle r e l a t i v e posi+,ion (BTC p i l o t ) shown in f i g u r e 4 a r e interchanged. A s of the end of 1976, t h i s reversed design had showed promise but had not been evaluated t o the point where it had surpassed t h e BTC p i l o t . The output t a r g e t s f o r the l a r g e r engine were a l l established from the RC1-60 t e s t r e s u l t s .

Although emissions w i l l be measured subsequently i n the program, none have been evaluated up t o t h i s point which has thus f a r concentrated on b a s i c con- f i g u r a t i o n and s y s terns evaluations. The f u e l economy and power milestones f o r t h i s program to develop a m i l i t a r y engine which, s i m i l a r t o an a i r c r a f t engine, emphasizes t h e higher output spectrum have a l l been m e t t o date. Nonetheless, a comparison of excerpted basic performance r e s u l t s is of i n t e r e s t f o r those phases of technology which are d i r e c t l y a p p l i c a b l e and because of t h e i l l u s t r a - Although the r e s u l t comparisons w i l l t i o n of s c a l i n g e f f e c t s t h a t it affords.

be from RCl-350 test r e s u l t s , t h e complete 4 r o t o r engine, t h e RC4-350, is shown i n f i g u r e 10 f o r r e l a t e d i n t e r e s t i n a multi-rotor engine.

The b a s e l i n e performance work on t h e 1-350 engine has a l s o been conducted w i t h t h e same i n s e r t e d r o t o r design and an 8.5:l compression r a t i o , although higher compression r a t i o r o t o r s w i l l be evaluated i n t h e near future.

The l a r g e r module s i z e has the g e n e r a l advantage of more a v a i l a b l e space to accomodate nozzle v a r i a t i o n s within a given r o t o r housing and, o p e r a t i o n a l l y , is l e s s constrained by spray impingement on t h e r o t o r and housing surfaces.

There a r e other advantages t o the l a r g e r combustion chamber s i z e , which include reduced s e a l i n g l i n e and leakage a r e a r a t i o t o charge volume, a s i m i l a r Eavor- a b l e r a t i o f o r h e a t losses, and the same type of reduction i n FMEP with s c a l e t h a t is generally observed with reciprocating engines.

To f a c i l i t a t e a d i r e c t comparison, the current a v a i l a b l e d a t a f o r the tn, engine s i z e s , both having t h e design configuration shown i n f i g u r e 4 (BTC p i l o t ) are compared on an Indicated b a s i s and equivalent (same apex seal v e l o c i t y ) RPM From f i g u r e 11 it can be seen t h a t the RC1-350 and RC1-60 are i n f i g u r e 11.

very c l o s e a t the lower IMEP1s, whereas t h e 1-60 d a t a shows lower ISFC (or b e t t e r thermal e f f i c i e n c y ) at t h e higher loads, i n d i c a t i n g f u r t h e r probable h- provements f o r t h e l a r g e r engine. The d i f f e r e n c e is believed t o r e f l e c t t h e concentration of e f f o r t a t t h i s speed f o r the smaller engine, i n v i e w of its automotive s i g n i f i c a n c e , whereas the low speed range of the l a r g e r engine is of less i n t e r e s t f o r c u r r e n t applications. For t h e reasons j u s t s t a t e d t h e r e is less a v a i l a b l e RC1-60 d a t a a t the higher speeds, but what is a v a i l a b l e suggests t h a t t h e thermal e f f i c i e n c i e s a r e reasonably c l o s e f o r both engines.

Figure 12 compares the RC1-350 d a t a of figure 11 plus available RC1-350 d a t a f o r the "reversed" configuration (ATC p i l o t ) mentioned e a r l i e r , versus E l : , r a t i o . The observed d a t a shows t h a t f o r a given mixture s t r e n g t h the RC1-60 develops higher IMEP's a t equivalent speed, which would imply more e f f e c t i v e air u t i l i z a t i o n .

This IMEP trend may be misleading because the engines w e r e n o t run with s i m i l a r induction systems. I f the IMEP d a t a is "normalized" by c o r r e c t i o n t o an equal volumetric e f f i c i e n c y b a s i s (which has l i t t l e e f f e c t on o t h e r plot- ted v a r i a b l e s ) , the RC1-60 and RC1-350 with B T C p i l o t a r e very c l o s e and t h e RC1-350 with ATC p i l o t is s l i g h t l y higher at t h e increased power end. The higher thermal e f f i c i e n c y of the RC1-350 A X p i l o t does not say t h a t the d i f f e r - ences noted w i l l n e c e s s a r i l y hold f o r the RC1-60 s i z e but it does imply t h a t t h e r e is a d d i t i o n a l p o t e n t i a l t o be realized.

Figure 13 shows both curves on a BSFC b a s i s , r e f l e c t i n g the d i f f e r e n c e s i n f r i c t i o n . Figure 1 3 shows t h a t , d e s p i t e the lower thermal e f f i c i e n c y a t higher power with the BTC p i l o t design, the RC1-350 shows a brake b a s i s advantage over t h e RC1-60 because of the lower s p e c i f i c f r i c t i o n . The A T C p i l o t configuration curve r e f l e c t s both f r i c t i o n and combust ion advantages. I n a d d i t i o n t o lower f r i c t i o n , the 350 cubic inch engine enjoys the advantage of b e t t e r i n j e c t i o n and i g n i t i o n equipment. The influence of t h i s l a s t point w i l l be c l e a r e r when the c u r r e n t RC1-60 t e s t i n g , which a l s o enjoys a s i m i l a r equipment advantage over the e a r l i e r work, has progressed f u r t h e r .

The conclusion of t h i s comparison is t h a t the engine scales d l , although demonstrated only i n the l a r g e r sized direction. The baseline d a t a of the RCl- 350 at higher powers and speeds, with the s c a l i n g trends noted, v i l l be used t o estimate performance f o r the a i r c r a f t engine regime. This i n p u t w i l l be impor- t a n t when weighing the system advantages of a l i g h t e r , smaller multi-rotor air- c r a f t engine versus a somewhat heavier, but less expensive and s l i g h t l y more e f f i c i e n t , l a r g e r module s i n g l e r o t o r engine. The f a c t o r s influencing t h i s balance process f o r the current NASA contract a r e thus i n c l e a r focus and Cessna A i r c r a f t Co., under sub-contrac t t o Curtiss-Wright , w i l l study the a i r c r a f t system trade-off s e n s i t i v i t y of various engine s i z e choices.

Cur r e n t N A S A Advanced Engine Study Approach and S t a t u s The objectives of the current =A Advanced Rotary Combust i o n A i r c r a f t Engine Design Study contract are t o define advanced and highly advanced engines which w i l l s a t i s f y the following goals and c r i t e r i a : Engine performance and e f f i c i e n c y improved a s compared t o current en- 1.

gines: BSFC <, 0.38 lblhp-hr (2 75% power cruise; s p e c i f i c weight 5 1.0 l b / hp @ takeoff power; cooling a i r f l o w x pressure drop product decreased by a f a c t o r of 2.

2 . E f f i c i e n t operation on 100/130 octane a v i a t i o n f u e l and one o r more a l t e r n a t i v e f u e l s such as j e t o r d i e s e l f u e l , o r low octane unleaded automotive f u e l .

3. Emissions t h a t meet the EPA 1979 piston a i r c r a f t standards. ( I f and when the revocation of these standards occurs, t h i s goal w i l l be reevaluated).

4. Engine d i r e c t manufacturing c o s t s comparable t o or less than present day spark-ignition p i s t o n a i r c r a f t engines.

5. Overall l i f e cycle c o s t s and maintenance lower than f o r current air- c r a f t engines.

6 . A l t i t u d e c a p a b i l i t y equal t o present day spark i g n i t i o n a i r c r a f t en- g i n e s .

The approach t h a t has been taken was t o f i r s t survey a l l p a r a l l e l and re- l a t e d technologies f o r a p p l i c a t i o n t o an extension of the S t r a t i f i e d Charge developments summarized e a r l i e r . A t o t a l of 35 s i g n i f i c a n t sources were iden- t i f i e d and s o l i c i t e d f o r information. In addition many hundreds of a b s t r a c t s located by source search were read and 220 papers obtained.

From a review of data from the above contacts, papers, and previous tech-

nology in£ ormation developed by Curtiss-Wright , the candidate technologies

shown i n Table I were selected f o r more detailed evaluation. The evaluation form ( f i g u r e 14) was developed as a means of carrying out the procedure f o r ranking of t h e c a n d i d a t e technologies. The technology e v a l u a t i o n c r i t e r i a were u t i l i z e d i n a system p a t t e r n e d a f t e r t h e one described i n r e f . 9.

A technology base was defined from which new approach s e l e c t i o n s were made f o r a n "advanced" engine. They were t h e approaches estimated t o be t h e most advanced t e c h n o l o g i e s s u f f i c i e n t l y proven and h i g h l y ranked t o be a v a i l a b l e to an engine design i n i t i a t e d i n 1985 o r 1986. It is estimated commercial i n t r o - d u c t i o n would t a k e p l a c e i n t h e e a r l y 1990's.

I n a d d i t i o n a s e l e c t i o n of d e s i g n approaches f o r a "highly advanced" en- g i n e were made. These were h i g h e r r i s k approaches l i k e l y t o r e q u i r e a more e x t e n s i v e development program a n d / o r a l a t e r i n t r o d u c t i o n t o t h e c o w e r c i a 1 market.

As a r e s u l t of t h i s ranking process, w i t h t h e a d d i t i o n a l balancing over- view r e f l e c t i n g concentrated r o t a r y engine experience of those who did not p a r t i c i p a t e i n ranking, s p e c i f i c c a n d i d a t e t e c h n o l o g i e s were s e l e c t e d (Table 11).

These i n p u t s w i l l be used t o d e f i n e a conceptual d e s i g n f o r t h e advanced en- gine. The "highly advanced" engine w i l l be d e s c r i b e d but not defined w i t h in- s t a l l a t i o n , c r o s s - s e c t i o n a l drawings, performance d a t a , e t c . which w i l l be de- veloped f o r t h e advanced engine. Comparative system a n a l y s i s w i l l be performed, however, by Cessna f o r both engine concepts i n compatible g e n e r a l a v i a t i o n a i r c r a f t .

U n t i l t h e d e s i g n study has been completed and we can assess t h e r e l a t i v e t r a d e - o f f s of t h e s e c a n d i d a t e t e c h n o l o g i e s a g a i n s t t h e c o n t r a c t o b j e c t i v e s and g o a l s , w e cannot s p e c i f i c a l l y weigh c o n t r i b u t i o n s i g n i f i c a n c e . However, t h e promising c h o i c e s have been s u f f i c i e n t l y d e f i n e d i n t h e aforementioned s c r e e n i n g p r o c e s s t o s i n g l e o u t s e l e c t e d items which can i l l u s t r a t e , i n t h e following paragraphs, t h e n a t u r e of o u r choices.

1. Turbocharging The requirement of a n e a r - f u t u r e a i r c r a f t engine (250 H P c r u i s e c l a s s ) f o r i n c r e a s e d a i t i t u d e (25,000 f e e t p l u s ) c a p a b i l i t y h a s focused more a t t e n t i o n on t h e e f f e c t s of turbocharging. Here, t h e Rotary S t r a t i f i e d Engine more c l o s e l y resembles a D i e s e l than a conventional g a s o l i n e f u e l e d engine, because of its a b i l i t y t o run w e l l on extremely l e a n mixture r a t i o s . I n c r e a s i n g the a i r charge rate t o t h e engine n o t only improves t h e f u e l economy by r a i s i n g the mechanical e f f i c i e n c y ( i . e., g e t t i n g more o u t p u t f o r e s s e n t i a l l y t h e same f r i c - t i o n l o s s e s ) , b u t i t p e r m i t s o p e r a t i o n a t A/F r a t i o s which g i v e t h e b e s t combus- t i o n and thermal e f f i c i e n c y . The c h a r a c t e r i s t i c curve shape f o r XSFC vs. mix- t u r e s t r e n g t h , shown i n f i g u r e 12 f o r low speed, holds f o r c r u i s e speeds a s w e l l a l t h o u g h t h e a b s o l u t e v a l u e s change w i t h speed. I n essence, the BSFC curve can e f f e c t i v e l y be d r i v e n down t o lower l e v e l s a s shown q u a l i t a t i v e l y i n f i g u r e 15 ( r e f . 10).

The q u a n t i t a t i v e degree t h a t can be p r a c t i c a l l y r e a l i z e d remains an unknown a t t h i s p o i n t , but from t r e n d s observed on t h e c u r r e n t n a t u r a l l y a s p i r a t e d s t r a t i f i e d engines, an SFC r e d u c t i o n of 17 percent can be p r e d i c t e d by high power c r u i s e turbocharging t o i n c r e a s e t h e a i r f l o w from an approximately 18 t o 28 a i r - f u e l r a t i o . The b a s e l i n e a b s o l u t e v a l u e of BSFC f o r t h e s t r a t i f i e d charge n a t u r a l l y a s p i r a t e d engine is probably n o t t h e same as it would be f o r t h e corresponding c r u i s e p o i n t of a g a s o l i n e engine a t its approximately l 5 : l a i r - f u e l r a t i o , b u t t h e turbocharged s t r a t i f i e d charge c r u i s e BSFC would be lower than e i t h e r t y p e ( s t r a t i f i e d o r homgeneous) n a t u r a l l y a s p i r a t e d engine.

2 . Increased LYEP and SpeedIImproved Apex S e a l Wear M a t e r i a l s I R e t r a c t i n g Apex S e a l s The i n c r e a s e of mean e f f e c t i v e p r e s s u r e i s accomplished by t h e tslrbo- charging d e s c r i b e d above, trading-off t h e c o m p i e x i t i e s of boost r a t i o s h i g t ? r t h a n can be a t t a i n e d from commercial low-cost turbocharger u n i t s a g a i n s t en- g i n e s i z e . However, wherever t h i s b e s t p o i n t r e s o l u t i o n o b t a i n s a s a r e s u l t of our c u r r e n t a n a l y s e s , t h e f a c t remains t h a t h i g h e r e f f e c t i v e p r e s s u r e s w i l l be required. These h i g h e r o p e r a t i n g l e v e l s of temperature and p r e s s u r e have b o t h s t r e s s and d u r a b i l i t y i m p l i c a t i o n s , which i n t u r n w i l l be r e f l e c t e d i n t'le s e l e z t i o n of s p e c i f i c o p e r a t i n g limits and d e s i g n c o n f i g u r a t i o n s , some of which w i l l be b r i e f l y reviewed i n succeeding paragraphs.

The same t y p e of trade-off h a s t o be made w i t h r e s p e c t t o maximum o p e r a t i n g speed. Higher speeds obviously i n c r e a s e t h e engine o u t p u t and t h u s improve s p e c i f i c power d e n s i t y . The Rotary engine h a s s i g n i f i c a n t growth p o t e n t i a l in t h e h i g h e r speed d i r e c t i o n because it is n o t l i m i t e d by v a l v e dynamics and v a l v e b r e a t h i n g r e s t r i c t i o n s , h a s complete dynamic balance, does not r e v e r s e d i r e c t i o n of its s e a l i n g elements a t t o p c e n t e r , and has a r e l a t i v e l y modest i n c r e a s e of f r i c t i o n w i t h speed. Xonetheless, f r i c t i o n i n c r e a s e s e x p o n e n t i a l l y w i t h speed and u n l e s s t h i s h i g h speed c a p a b i l i t y is reserved only f o r take-off power, t h e b e s t s p e c i f i c f u e l consumption w i l l d i c t a t e r a t i n g a t t h e lowest p o s s i b l e speed c o n s i s t e n t with a c c e p t a b l e s p e c i f i c weight. Again, t h e Cessna s e n s i t i v i t y s t u d y w i l l provide some i n s i g h t s i n t o how t h i s higher speed c a p a b i l i t y can be b e s t u t i l i z e d .

a.

Improved Apex Seal/Trochoid ? f a t e r i a l Combinat i o n s The i n c r e a s e i n engine e u t p u t may r e q u i r e f u r t h e r development of s u p e r i o r apex seal and t r o c h o i d wear s u r f a c i n g m a t e r i a l s which have e i t h e r been i d e n t i f i e d by o u r p r i o r r e s e a r c h e f f o r t s o r have emerged a s new t e c h n o l o g i e s .

The c u r r e n t t u n g s t e n c a r b i d e t r o c h o i d wear s u r f a c i n g m a t e r i a l has t h u s f a r shorn r e l a t i v e l y low apex s e a l v e l o c i t y s e n s i t i v i t y and is adequate, with a c c e p t a b l e TBO and r e l i a b i l i t y s t a n d a r d s , f o r any o p e r a t i n g speed under consi- d e r a t i o n ( r e f . 11). It w i l l probably a l s o prove a c c e p t a b l e , p o s s i b l y with lower wear apex s e a l s , f o r any of t h e INEP l e v e l s which can be obtained w i t h s i n g l e s t a g s turbocharging. However, t o i l l u s t r a t e p o t e n t i a l , f i g u r e 16 shows t h a t use of a Titanium c a r b i d e t r o c h o i d c o a t i n g , in t h i s c a s e i n a s t e e l m a t r i x , and w i t h apex s e a l s of t h e same m a t e r i a l , shows s u b s t a n t i a l l y l e s s wear than c u r r e n t m a t e r i a l s . The p a r t i c u l a r m a t e r i a l shown i n t h i s f i g u r e was plasma sprayed, which i s a l e s s expensive a p p l i c a t i o n technique than t h e c u r r e n t deto- n a t i o n gun p r o c e s s , A t t h i s s t a g e of development, plasma-spraying cannot a t t a i n t h e spme bond s t r e n g t h s , but plasma-spray technology is moving very f a s t and is expected t o be a s e r i o u s contender w i t h i n a s h o r t time.

b. R e t r a c t i n g Apex S e a l s For a more a m b i t i o u s technology s t e p , which we r e s e r v e f o r t h e "Highly is p o s s i b l e t o have t h e h i g h s p e c i f i c o u t p u t of h i g h speed Advanced ~ e s i g n " , i t w i t h o u t f a c i n g t h e more s e v e r e apex s e a l wear environment of h i g h e r s e a l pres- s u r e s p l u s h i g h e r speed. S i n c e apex s e a l leakage is a tine-weighted f a c t o r , a t high e n g i n e speeds a s m a l l leakage a r e a can be t o l e r a t e d w i t h o u t s e r i o u s consequence. S e a l d e s i g n s which r e t r a c t from t r o c h o i d c o n t a c t a t high r o t a - t i o n a l speeds a r e a v a i l a b l e , but n o t t e s t e d . One of s e v e r a l a l t e r n a t e ap- proaches, i n t h i s c a s e t a k i n g advantage of t h e c e n t r i f u g a l f o r c e s t o p u l l t h e s e a l back a t high speeds, is shown i n f i g u r e 17.

3. High S t r e n g t h High Temperature Alminum C a s t i n g Alloy The i n c r e a s e s i n IMEP and speed, a s s t a t e d e a r l i e r , w i l l i n t r o d u c e h i g h e r o p e r a t i n g temperatures. The a n t i c i p a t e d degree of temperature i n c r e a s e , t o be confirmed a s t h e c u r r e n t s t c d y p r o g r e s s e s , can be paced by t h e degree of s t r e n g t h improvement t h a t new m a t e r i a l s have i n t r o d u c e d . The c h o i c e of l i q u i d c o o l i n g f o r g e n e r a l a v i a t i o n e n g i n e s ( r e f . 10) on t h e b a s i s of improved system e f f i c i e n c y and b e t t e r m e t a l temperature c o n t r o l is p a r t i c u l a r l y s i g n i f i c a n t a t t h e h i g h e r o u t p u t s of t h e advanced engines.

E s s e n t i a l l y a l l of our Rotary engine c a s t aluminum r o t o r housings have been A M 4220, based on our r e c i p r o c a t i n g a i r c r a f t e n g i n e experience. I t has proven t o be a d u r a b l e h i g h temperature m a t e r i a l w i t h good f a t i g u e l i f e under c y c l i c loading. X new aluminum high temperature c a s t i n g a l l o y , AMS 4229, h a s been on t h e scene f o r s e v e r a l y e a r s . It h a s n o t been t r i e d h e r e because o u r a p p l i c a t i o n s have n o t r e q u i r e d t h e additions: s t r e n g t h and, u n t i l r e c e n t l y , v e r y few f o u n d r i e s were ~ i l l i n g t o c a s t t h e new a l l o y . Today, however, 1 5 f o u n d r i e s i n t h e U.S. use t h i s a l l o y , which h a s markedly h i g h e r s t r e n g t h and d u c t i l i t y t h a n XEIS 4220.

F i g u r e 1 8 shows c a l c u l a t e d p r e d i c t i o n s , based on u l t i m a t e t e n s i l e s t r e n g t h , d u c t i l i t y , and modulus of e l a s t i c i t y , of low c y c l e thermal f a t i g u e l i f e a t 400°F, r e p r e s e n t a t i v e of high power c r u i s e peak t e m p e r a t u r e s , f o r MlS 4220 and 4229.

The s a n e type of improvements can be demonstrated a t h i g h e r temperature l e v e l s , should they prove d e s i r a b l e a s t h e s t u d y p r o g r e s s e s .

4. Rotor Combustion Flank I n s u l a t i o n / ~ d i a b a t i c Engine The background d i s c u s s i o n of S t r a t i f i e d Charge hydrocarbon t e s t i n g made r e f e r e n c e t o r o t o r combustion s u r f a c e s which were r a i s e d i n temperature, by use of i n s u l a t e d p l a t e s , t o reduce HC formation. HC formation is not expected t o be a c o n s i d e r a t i o n for a n a i r c r a f t engine o p e r a t i n g regime ( r e f . l o ) , b u t t h e insu- l a t e d r o t o r s u r f a c e w i l l reduce o i l heat r e j e c t i o n and thus reduce system weight and bulk.

E a r l y t e s t i n g with t h e RC engine had shown t h a t Zirconium o x i d e , plasma sprayed on t h e r o t o r combustion f a c e , was e f f e c t i v e with g a s o l i n e homogeneous charge e n g i n e s , but d i d not have adequate thermal shock s t r e n g t h i n a s t r a t i - f i e d charge a p p l i c a t i o n where d i r e c t f u e l impingement was p o s s i b l e . However, considerable development of thermal b a r r i e r coatings of t h i s type has taken place, l a r g e l y a t NASA-Lewis, s i n c e t h a t time, p a r t i c u l a r l y f o r gas turbine components.

The "Highly Advanced Engine" (Table 11) r e f l e c t s inclusion of a zirconium oxide r o t o r h o t s u r f a c e coating of .060" thickiiess. Figure 19 shows t h a t , f o r an assumed 90"3 r o t o r t h i s coating thickness is calculated t o reduce the r o t o r heat r e j e c t i o n t o t h e engine o i l by approximately one t h i r d .

The same type of coating would a l s o be a p p l i c a b l e f o r an " ~ d i a b a t i c " Rotary engine. However, d e s p i t e t h e f a c t t h a t w e consider Rotary engines in- t r i n s i c a l l y more adaptable to the completely unlubricated "Adiabatic" engirie approach than a reciprocating engine ( l a r g e l y because e i t h e r r e t r a c t i n g o r the uni-directional ceramic apex s e a l s with t h e i r zdvantage of gas hydrodynamic f i l m l u b r i c a t i o n , against a ceramic trochoid s u r f a c e appear c l o s e r t o r e a l i z a - t i o n than t h e i r reciprocating counterparts) w e did not consider an engine of t h i s type t o be within c o n t r z c t o b j e c t i v e g u i d e l i n e s of even t h e "highly ad- vanced design technology" and did not consider it f u r t h e r .

Direct i o n s Our present carbureted prototype a i r c r a f t engine, t h e liquid-cooled RC2-75 ( r e f . 10) shown i n f i g u r e s 20 and 21 was t h e obvious s t a r t i n g point.

To i l l u s t r a t e what t h e presently planned t r e n d s of higher IMEP and RPM would mean i n terms of t h i s engine, mock-ups of both s i n g l e and twin r o t o r 75 cubic inch S t r a t i f i e d Charge and turbocharged a i r c r a f t engines have been prepared t o supplement t h i s presentation. The RC1-75, which measures 34 1/2" x 21 1/2" x 20", without coolers, is predicted t o develop 235 HP @ take-off under c u r r e n t l y envisioned l i m i t s f o r IMEP and speed f o r the advanced engine and 300 HP f o r t h e "Highly ~dvanced" technology. The same numbers f o r the RC2-75 a r e 40 114" x 21 1/2" x 20" and 470 H P and 600 H P f o r r e s p e c t i v e technology l e v e l s . Taking t h e same vdlues of IMEP and speed used to make t h e advanced engine e s t ~ m a t e s , and applying it t o the c o n t r a c t goal engine of 250 HP c r u i s e to 25,000 f e e t , r e s u l t s i n a two r o t o r engine o u t l i n e as shown i n f i g u r e 22. The coolact and o i l coolers, which presumably would be remotely located f o r system optimization, a r e not included.

This par t f cu, lr arrangement, with accessories held t o a minimum o v e r a l l diameter, would be biased towards a twin engine i n s t a l l a t i o n . However, it should be understood t h a t t h i s is a preliminary look s i n c e , a s s t a t e d e a r l i e r , t h e economic advantages, a s well a s f u e l in5kction technology l i m i t s , favor a l a r g e r s i n g l e r o t o r engine and, secondarily, more p r e c i s e d e f i n i t i o n of IMEP and speed, upon which these projections were based, is still being determined.

Wl~ile t h e s p e c i f i c s may vary from those defined a t t h e point when a l l work on t h i s c o n t r a c t has been completed, the e f f o r t t o d a t e has c i a r i f i e d what we s e e a s the advantages of t h i s type of Rotary S t r a t i f i e d Charge engine. These a r e l i s t e d i n Table 111.

Closure The Rotary Engine, in its carbureted form, is rmiquely suited t o a i r c r a f t engine propulsion because of its ad\-antages of size, weight, simplicity, saooth- ness, scaling f l e x i b i l i t y and grout:& po &ential. Recent p a r a l l e l hardware and test developments have shown t h a t t h i s engine type is p a r t i c u l a r l y adaptable t o unthrottled d i r e c t injected s t r a t i f i e d charge, resulting in additional f e a t u r e s of wide range f u e l capability and superior fuel economy. A s a r e s u l t of the S t r a t i f i e d Charge Rotary Engine's a j i l i t y t o perform e f f i c i e n t l y over a broader range of mixture strengths, w-thout regard f o r f u e l octane o r cetane rating, turboctarging ran extead the demonstretea p o t e n t i a l to the complete a i r c r a f t engine operating regire. This canbination of systen; efficiency plus f u e l choice optimization p o s s i b i l i t i e s is !xi= carefully exa~lined as we continue i n t o an e r a of c r i t i c a l energy resource allocations.

Jones, Charles; Lamping, Harold D,; Myers, David M.; at$ Loyd, R.W.: An Update of t h e D i r e c t I n j e c t e d S t r a t i f i e d Charge Rotary Conbustion

Enginc Develqnnents at Curtiss-Wright . SAE T r a n s a c t i o n s 1978, Paper

No. 770044.

J o n e s , Charles: A Survey of ~urtiss- right's 1958-1971 R o t a t i n g Combustion Engine Teclw- logical Developments. Paper No. 720468, Hay 1979.

Hofbauer, P.; and S t a t o r I ( . : Advanced Pover Systems, P a r t 2 - A D i e s e l

f o r a Subcompact Car. SAE Paper No. 770113, February 1977.

C o n t r a c t No. WIT-TSC-1193 C o n t r a c t o r ' s P r e s e n t a t i o n , September 1977:

Data Base f o r Lightweight Automative Diesel Power P l a n t s - Volkswagenuerk

AG .

W i e d e m a ~ , B.; and Hofbauer P.: Data Base f o r Light-Weight Automotive D i e s e l Power P l a n t s . U E Paper No. 780634, June 1978.

Tierney, W.T.; Jahnson, E.H.; and Crawford, N.R.: Energy Conservation of t h e Vehicle-Fuel-def i n e r y System. SAE Paper No. 75067 3, June 1975.

Tierney, W . T . : and Wilson, R.F.: Adequate F u t u r e T r a n s p o r t a t i o n Demands Vehicle-Fuel-Ref i n e r y System Optimization Today. Presented a t t h e API Michigan Meeting, January 1976.

Jones, Charles: A P r o g r e s s Report on Curtiss-Wright's Rctarg S t r a t if i e d Charge Engine Development. SAE Paper No. 741206, October 1974.

Bergcy, K a r l H . : New Technologies for General Aviation A i r c r a f t . SA5 Paper No. 790613, A p r i l 1979.

Jones, C t ~ a r l e s : A Review of Curtiss-Wright 's Rotary wine Developments w i t h Respect t o General Aviation P o t e n t i a l . SAE Paper No. 790621, A p r i l 1979.

Lamping, Harold D. ; Gal l i e r s , M.W. ; and Wolosin, S.M. : Rotary Combustion Engine Trochoid Coatings and S e a l s . SAE Paper No. 741043, October 1974.

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VARIABLE AREA T'iRBIRf i W R O C H M R RETRACTING APEX SfALS ROTOR C(IIIBUSIIOI( FLAIIK 1 N S U A T l l M A D D i T I W l II(CREASf0 ICCP M D RPN TABLE 111 .WANTAGES OF 'IWE ROTARY STRATIFIED CHARGE AIRCRAFT ENGINE POSITIVE DISPUCEMENT TYPE GENERAL-AVIAT ION ENGINES: SUMMARY AND CONCLUDING REMARKS Erwin E. Kempke, J r .

National Aeronautics and Space Administration Lewis Research Center During t h i s session, the i n t e r n a l cocnbustion engi,,e program, i t s status and accoolpl ishments have been presented. Both t h e near and longer term technical thrusts were discussed. S i g n i f i c a n t progress on both f r o n t s has been made.

I n t h e near term conventional engine area the i n t e g r a t i o n o f f o u r modifications i n a t e s t engine provides f o r a 10 percent improvement i n t h e high-performance-cruise f u e l economy while meeting t h e emissions standards.

,I new program aimed a t improved cooling and drag reduction has been establ i shed. The Phase I e f f o r t on improved-cyl i nder head-and-barrel cool i ng w i 11 be underway a t Teledyne Continental Uotors, A i r c r a f t Products Division, by the end o f t h i s year.

The grant program t o develop and v e r i f y a r e a l j s t i c Otto cycle computer model shows s i g n i f i c a n t progress a t the two-thirds completion point. Today, calculations o f two-dimensional unsteady, turbulent, compressible f l o w w i t h moving boundaries are being made. The r e l a t e d aabi t ious experimental program t o develop advanced combustior, diagnostic: techniques has been completed.

I n the f u e l - i n j e c t i o n techno1 ogy program t h e contract t o Spectron i s quanti t i v e l y c h a r a t t e r i z i ng the performance o f various i n j e c t o r nozzles. The in-house flow v i s u a l i z a t i o n work i s studying t h e same set of nozzles under motored engine cocdi t ions by means o f high-speed photography.

The e f f o r t t o define t h e benefits and requirements o f advanced but cost effective turbocharger technology has been i n i t i a t e d . The RFP f o r Phase I i s expected t o be issued i n e a r l y 1980.

For the longer term the r e s u l t s o f the on-going studies i n v o l vicg a l t e r - native engine configurations were presented. The advanced spark-ignition p i s t o n engine study, which i s 75 percent complete, shows 0.33 and 0.36 BSFC f o r t h e s t r a t i f i e d charge and lean burn concepts, respectively. The two-stroke diesel shows 0.36 BSFC and 1.07 lb/hp f o r the 250-hp cruise version.

The stratified-charge r o t a r y engine study, which has been completed t o the technology assessment task, continues t o be a viable candidate.

and Cessna w i l l soon be underway.

The airplane and mission studies by Beect Results from these studies and the on-going Lewis e f f o r t w i l l o b j e c t i v e l y evaluate, rank, dnd compare the three study engines w i t h each othzr, w i t h representative current-production engines, and w i t h a h i g h l y advanced small PEEDING PAGE GLANK NOT FILMED t ~ r b o p r o p i n t e n s relevant t o t h e industry. The r e s u l t o f these a c t i v i t i e s i n l a t e FY 80 w i l l be reconmendations as t o which concepts m e r i t serious technolcw enabl ement programs.

The supporting research and technology included both contract and in-house elements. Diesel and ~ c t a r y engine t e s t c e l l s a t Lewis are now operational w i t h baseline mapping complete. Active research programs are underway i n such areas as supercharge versus compression r a t i o tradeoffs f o r optimum performance; f u e l - i n j e c t i o n rate, timing, and other parameters; and ceramic combustion- chamber i n s u l a t i o n materials. A Curtiss-Wright program confirmed t h z t efficiency improvements i n the RC2-75 could be achieved by increasing compression r a t i o and r e l o c a t i n g spark plugs: t h e BSFC a t c r u i s e was improved from 0.54 o r i g i n a l l y t o 0.47.

The U n i v e r s i t y o f Michigan, working under a NASA grant, recently completed f e a s i b i 1 i t y t e s t s o f a diesel ized current-production spark i g n i t i o n c y l inder.

These t e s t s indicated t h a t moderate f i r i n g pressures a t low compression r a t i o resulted i n improved c r u i s e BSFC, thus i n d i c a t i n g the p o s s i b i l i t y f o r conziderably reducing t h e s p e c i f i c weight o f a diesel.

I n conclusion, t h e r e s u l t s o f the near-term conventional engine a c t i v i t i e s have shown t h a t it may soon be possible t o improve both economy and cooling efficiency. Since t h e technology i s being developed i n terms o f current p r o d ~ c t i o n engines under t h e constraint o f remaining compatible w i t h e x i s t i n g f a c i l i t i e s , processes, etc., it can be incorporated i n t o OEM production w i t h l i t t l e adverse e f f e c t . By the same token t h i s technology i s p o t e n t i a l l y a v a i l a b l e f o r incorporation by r e t r o f i t , Farther i n t o the f u t u r e it appears t h a t the i n t e r n a l combustim sngines discussed today are a l l viable candidates. They have the p o t e n t i a l of improving s i g n i f i c a n t l y on the present s i t u a t i o n i n terms o f f u e l economy, weight, a l t e r n a t i v e f u e l s c a p a b i l i t y , and other characteristics. I n a d d i t i o n t o t9ese r e a d i l y quantifiable benefits, rhere are f u r t h e r b e n e f i t s i n the areas o f safety and all-around u t i l i t y . Over-the-weather cruise a l t i t u d e c a p a b i l i t i e s could eliminate up t o 25 percent o f a l l neather-related accidents a t the same time it i s providing more e f f i c i e n t and comfortable f l y i n g conditions. A 1 though the candidate engine selection process w i l l be d i f f i c u l t , i t i s anticipated t h a t a p r e l i m i nary select i o n may be possible a f t e r comparative eng! ne/ a i rframe r e s u l t s are in hand. However, regardless o f which candidate i s selected, the p o t e n t i a l powerpl ant advancements are synergistic w i t h expected improvements i n structures, aerodynamics, materials, and avionics which together w i l l r e s u l t i n s i g n i f i c a n t l y improved airplanes f o r the 1990's and beyond.

NASA PROPELLER TECHNOLOGY PROGRAM Daniel C . Mikkelson National Aeronautics and Space Administration Lewis Research Center The v a s t m a j o r i t y of g e n e r a l - a v i a t i o n a i r c r a f t manufactured i n t h e United S t a t e s a r e p r o p e l l e r powered (approximately 98 p e r c e n t i n 1978). Most of t h e s e a i r c r a f t u s e p r o p e l l e r d e s i g n s based o n technology t h a t h a s n o t changed s i g n i f - i c a n t l y s i n c e t h e 1940's and e a r l y 1950's. T h i s o l d e r technology h a s been ade- q u a t e ; however, w i t h t h e c u r r e n t world energy s h o r t a g e and t h e p o s s i b i l i t y o f more s t r i n g e n t n o i s e r e g u l a t i o n s , improved technology i s needed. S t u d i e s con- d u c t e d by NASA and i n d u s t r y i n d i c a t e t h a t t h e r e a r e a number o f improvements i n t h e technology of g e n e r a l - a v i a t i o n p r o p e l l e r s t h a t c o u l d l e a d t o s i g n i f i c a n t energy savings. New c o n c e p t s l i k e b l a d e sweep, p r o p l e t s , add composite mate- 'rials, a l o n g w i t h advanced a n a l y s i s t e c h n i q u e s have t h e p o t e n t i a l f o r improving t h e per2ormance and lowering t h e n o i s e o f f u t u r e propeller-powered a i r c r a f t t h a t c r u i s e a t low speeds. C u r r e n t p r o p e l ler-powe r e d gene-al-aviat i o n a i r c r a f c are l i m i t e d by p r o p e l l e r c o m p r e s s i b i l i t y l o s s e s t o maximum c r u i s e s p e e d s n e a r Mach 0 5. The techcology b e i n g developed a s p a r t of NASA's -2dvaaced Turboprop Proj- e c t o f f e r s t h e p o t e n t i a l o f e x t e n d i n g t h i s l i m i t t o a t l e a s t Mach 0.8. A t t h e - ?

h i g h e r c r u i s e speeds, advanced t u r b o p r o p p r o p u l s i o n h a s t h e p o t e n t i a l o f l a r g e energy s a v i l ~ g s comparzd w i t h a i r c r a f t powered by advanced t u r b o f a n systems.

T h i s p a p e r surmnarizes NASA's program on p r o p e l l e r technology a p p l i c a b l e t o b o t h l o w and h i g h speed g e n e r a l - a v i a t i o n a i r c r a f t , and o u t l i n e s t h e o v e r a l l pro- gram o b j e c t i v e s and approach.

EFFICIENCY TRENDS The f r e e - a i r p r o p e l l e r i s t h e p r o p u l s i v e d e v i c e t h a t h a s t h e h i g h e s t l e v e l o f i n h e r e n t e f f i c i e n c y f o r s u b s o n i c a i r c r a f t . A comparison o f t h e i n s t a l l e d c r u i s e e f f i c i e n c y of propel ler-powe red and t u r b o f an-powe red p r o p u l s i o n systems is shown i n f i g u r e 1 f o r a range o f c r u i s e speeds. The i n s t a l l a t i o n l o s s e s i n - cluded w i t h t h e propeller-powered systems a r e n a c e l l e d r a g and i n t e r n a l c o o l i n g a i r f l o w l o s s e s . For t h e turbofan-powered systems t h e l o s s e s i n c l u d e f a n cowling e x t e r n a l d r a g and t h e i n t e r n a l f a n a i r f l o w l o s s e s a s s o c i a t e d w i t h i n l e t recovery and n o z z l e e f f i c i e n c y . The i n s t a l l e d e f f i c i e n c y a v a i l a b l e w i t h c u r r e n t t e c h n o l - ogy propeller-powered g e n e r a l - a v i a t i o n (GA) a i r c r a f t ranges from about 70 t o 75 p e r c e n t f o r r e c i p r o c a t i n g powered a p p l i c a t i o n s t o s l i g h t l y o v e r 8 0 p e r c e n t f o r turboprops. The. r e c i p r o c a t i n g system performance i s s l i g h t l y lower due t o high- er n a c e l l e d r a g and l a r g e i n t e r n a l c o o l i n g a i r f l o w l o s s e s ( r e f . 1 ) . The i n - s t a l l e d performance of t h e c u r r e n t lower speed t u r b o p r o p systems remains high t o about Mach 0.5; a b o u t t h i s s p e e d , e f f i c i e n c y i a l l s o f f s i g n i f i c a n t l y because o f l a r g e p r o p e l l e r c o m p r e s s i b i l i t y l o s s e s . These p r o p e l l e r s a r e g e n e r a l l y d e s i g n e d w i t h b l a u e s o f t h i c k n e s s t o chord r a t i o s ( a t 75 p e r c e n t r a d i u s ) t h a t range from about 5 t o 7 p e r c e n t . These r a t h e r t h i c k b l a d e s , when o p e r a t e d a t r e l a t i v e l y high t i p h e l i c a l Mach numbers, a r e t h e main c a u s e o f t h e s e l o s s e s .

: ! $ - s * - .

. -

The advanced, high-spekd t u r b o p r o p shown i n f i g u r e 1 is a new p r o p u l s i o n concept t h a t h a s t h e p o t e n t i a l of e l i m i n a t i n g o r minimizing c o m p r e c s i b i l i t y The l e v e l of p o t e n t i a l i n s t a l l e d e f f i - l o s s e s a t f l i g h t speeds t o Mach 0 . 8 .

c i e n c y p r o j e c t e d f o r t h e advanced t u r b o p r o p i s c o n s i d e r a b l y h i g h e r t h a n t h a t A t Mach 0.8 a v a i l a b l e w i t h comparable technology high-bypass t u r b o f a n systems.

t h e i n s t a l l e d e f f i c i e n c y o f t u r b o f a n systems would be approximately 65 p e r c e n t compared w i t h about 75 p e r c e n t f o r t h e advanced turboprop. T h i s large p e r f o r mance advantage f o r t h e advanced t u r b o p r o p may o f f e r t h e p o t e n t i a l f o r some rt- t r a c t i v e energy s a v i n g s f o r f u t u r e h i g h performance b u s i n e s s a i r c r a f t .

ADVANCED, HIGH-SPEED TURBOPROP To a c h i e v e t h e performance p o t e n t i a l of t h e advanced, high-speed turboprop, s e v e r a l new c o n c e p t s and advanced t e c h n o l o g i e s a r e r e q u i r e d . ( S e e f i g . 2.)

These new c o n c e p t s and advanced t e c h n o l o g i e s a r e d i s c u s s e d i n r e f e r e n c e s 2 and 3. The advanced p r o p e l l e r would be powered by a large, modern t u r b o s h a f t e n g i n e and gearbox t o p r o v i d e t h e maximum power t o t h e p r o p e l l e r w i t h a minimum e n g i n e f u e l consumption. P r o p e l l e r e f f i c i e n c y would b e k e p t h i g h by minimizing com- p r e s s i b i l i t y l o s s e s . I n t h e o u t b o a r d p a r t o f t h e p r o p e l l e r b l a d i n g , these l o s s e s would be minimized by u s i n g sweep and t h i n b l a d e s e c t i o n s (2.4 p e r c e n t t h i c k n e s s t o chord r a t i o a t 75 p e r c e n t r a d i u s ) . Blade sweep would a l s o reduce p r o p e l l e r s o u r c e n o i s e b o t h d u r i n g t a k e o f f and l a n d i n g and d u r i n g high-speed c r u i s e . I n t h e inboard r e g i o n a n a r e a - r u l e d s p i n n e r , i n combination w i t h a n i n t e g r a t e d n a c e l l e shape, would be u s e d t o r e d u c e t h e l o c a l v e l o c i t i e s through t h e p r o p e l l e r t o minimize l o s s e s i n t h i s region. A power l o a d i n g ( s h a f t horse- power d i v i d e d by p r o p e l l e r d i a m e t e r squared) a b o u t f i v e t i m e s h i g h e r t h a n t h a t i n c u r r e n t GA turboprops would b e used t o minimize p r o p e l l e r diameter and weight. Eight o r t e n b l a d e s would be r e q u i r e d t o maximize i d e a l e f f i c i e n c y d u r i n g h i g h - a l t i t u d e , high-speed c r u i s e . I n a d d i t i o n t o these advanced c o n c e p t s a modern blade f a b r i c a t i o n t e c h n i q u e wouid b e u s e d t o c o n s t r u c t t h e t h i n , h i g h l y swept and t w i s t e d blades.

The program t h a t N A S A h a s underway t o a d d r e s s t h e technology r e q u i r e m e n t s of t h e advanced turboprop i s shown i n f i g u r e 3. The advanced turboprop p r o j e c t , p a r t o f NASA's A i r c r a f t Energy E f f i c i e n c y (ACEE) program, h a s t h e g o a l s o f a 15 t o 30 p e r c e n t f u e l s a v i n g r e l a t i v e t o turbofan-powered a i r c r a f t , a s i g n i f i c a n t r e d u c t i o n i n turboprop propulsion-system-related o p e r a t i n g c o s t s , and a c a b i n r i d e q u a l i t y e q u i v a l e n t t o t h e b e s t turbofan-powered a i r c r a f t . The f o u r major e l e m e n t s o f t h e advanced t u r b o p r o p p r o j e c t a r e shown i n f i g u r e 3.

I n t h e f i r s t , p r o p e l l e r and n a c e l l e , technology work is c u r r e n t l y under- way i n p r o p e l l e r aerodynamics, a c o u s t i c s , and b l a d e s t r u c t u r e s . The u s e o f advanced aerodynamics c o n c e p t s i n t h e d e s i g n of high-speed p r o p e l l e r s i s d i s - cussed i n r e f e r e n c e 2, and some r e c e n t wind t u n n e l r e s u l t s a r e p r e s e n t e d i n r e f - e r e n c e 4. A photograph o f a n advanced high-speed p r o p e l l e r model is shown i n f i g u r e 4. T h i s model, a l o n g w i t h t h r e e o t h e r s , was t e s t e d i n t h e Lewis 8- by 6-foot wind tunne 1. High performance and some s i g n i f i c a n t n o i s e r e d u c t i o n s were o b t a i n e d d u r i n g high-speed c r u i s e t e s t i n g o f t h e s e models. Some d e s i g n s t u d y r e s u l t s o n advanced p r o p e l l e r a c o u s t i c s and b l a d e s t r u c t u r e s a r e c o n t a i n e d i n r e f e r e n c e s 5 and 6.

The second major p r o j e c t element, c a b i n environment ( f i g . 3 1 , c o n c e r n s t h e a i r c r a f t f u s e l a g e , which may be i n t h e d i r e c t n o i s e p a t h of t h e p r o p e l l e r .

The f u s e l a g e w i l l have t o a d e q u a t e l y a t t e n u a t e t h i s n o i s e s o u r c e i f t h e c a b i n en- vironmental g o a l s a r e t o be achieved. Some r e c e n t a n a l y t i c a l s t u d i e s o n f u s e - l a g e i n t e r i o r n o i s e c o n t r o l f o r high-speed t u r b o p r o p s a r e p r e s e n t e d i n r e f e r - e n c e s 7 and 8. Fuselage v i b r a t i o n is a l s o a n important c a b i n environmental con- s i d e r a t i o n , and f u t u r e advanced turboprop a i r c r a f t w i l l have t o be d e s i g n e d t o minimize o r c o n t r o l any u n d e s i r a b l e v i b r a t i o n s .

The t h i r d major element, i n s t a l l a t i o n aerodynamics, i s concerned w i t h t h e a c c e l e r a t i n g , s w i r l i n g , p r o p e l l e r s l i p s t r e a m p a s s i n g o v e r a wing. The technol- ogy c h a l l e n g e h e r e i s t o d e s i g n t h e o v e r a l l a i r c r a f t t o a c h i e v e t h e b e s t combin- a t i o n o f p r o p u l s i o n system performance and a i r p l a n e l i f t - t o - d r a g r a t i o , w h i l e m a i n t a i n i n g adequate a i r c r a f t s t a b i l i t y and c o n t r o l .

R e s u l t s from a r e c e n t e x p e r i m e n t a l i n v e s t i g a t i o n o f p r o p e l l e r s l i p s t r e a m wing i n t e r a c t i o n s a t s r u r s e s p e e d s n e a r Mach 0.8 a r e p r e s e n t e d i n r e f e r e n c e 9.

The f i n a l major element is t h e key mechanical components.

Advanced d e s i g n and packaging technology f o r t h e c o r e e n g i n e , gearbox, and p r o p e l l e r w i l l be r e q u i r e d i f a n advanced t u r b o p r o p i s t o reduce maintenance c o s t s and improve A s t u d y o f c u r r e n t - g e n e r a t i o n t u r b o p r o p r e l i a b i l i t y and main- r e l i a b i l i t y .

t e n a n c e c o s t s is p r e s e n t e d i n r e f e r e n c e 10 a l o n g w i t h a n e s t i m a t e of t h e poten- t i a l improvements a v a i l a b l e from advanced technology.

Because t h e f o u r major e l e m e n t s of t h e advanced turboprop p r o j e c t a r e h i g h l y i n t e r r e l a t e d , a i r c r a f t t r a d e - o f f s t u d i e s a r e b e i n g made t o i n s u r e t h a t t h e s e t e c h n o l o g i e s a r e p r o p e r l y i n t e g r a t e d and t h a t p r o g r e s s i s b e i n g made toward a c h i e v i n g t h e o v e r a l l p r o j e c t goa;s. A summary o f t h e p r o g r e s s made under NASA's advanced t u r b o p r o p p r o j e c t is c o n t a i n e d i n r e f e r e n c e 3. Also, p a p e r s t h a t were p r e s e n t e d a t t h i s conference by Gatzen, J e r a c k i , and Bober show t h e p o t e n t i a l and some of t h e r e c e n t aerodynamic advances made a s p a r t of t h i s p r o j e c t .

LOW-SPEED PROPELLERS Current-generat i o n propel ler-powe red GA a i r c r a f t o p e r a t e a t c r u i s e speeds o f Mach 0.5 and below. The technology t r e n d s :nat a r e p r o j e c t e d c ~ r t h e pro- p e l l e r s of t h e s e lower speed a i r c r a f t a r e shown i n f i g u r e . 'lne s k e t c n i n t h e lower l e f t of t h i s f i g u r e d e p i c t s c u r r e n t - t e c h n o i o g y p r o p e l i e r s . These pro- p e l l e r s a r e designed based on a trade-of f of t h e f o u r f a c t o r s e n c l o s e d i n t h e c e n t e r c i r c l e . F o r many a p p l i c a t i o n s performance i s t r a d e d o f f t o meet n o i s e and c o s t g o a l s . S o l i d aluminum b l a d e c o n s t r u c t i o n is used i n most d e s i g n s , and t h i s c a n r e s u l t i n a p o t e n t i a l weight p e n a l t y when compared w i t h some of t h e f u t u r e advanced m a t e r i a l s which a r e being s t u d i e d . The d e s i g n t r a d e - o f f on f u t u r e advanced technology l o r s p e e d p r o p e l l e r s may b e a l t e r e d because of two key f a c t o r s - o u t s i d e d r i v e r s and technology o p p o r t u n i t i e s . The o u t s i d e d r i v e r s a r e t h e h i g h c o s t and f u e l a v a i l a b i l i t y problems due t o t h e energy s h o r t a g e , t h e p o s s i b i l i t y of more s t r i n g e n t government n o i s e r e g u l a t i o n s , and t h e need t o re- t a i n o r improve a i r c r a f t s a f e t y .

I n t h e technology o p p o r t u n i t y a r e a s e v e r a l new c o n c e p t s a r e c u r r e n t l y under s t u d y t h a t show c o n s i d e r a b l e performance and n o i s e b e n e f i t s . Advanced a n a l y s i s t e c h n i q u e s w i l l make i t p o s s i b l e t o b e t t e r u n d e r s t a n d p r o p e l l e r and n a c e l l e aerodynamics and a c o u s t i c s f o r a d d i t i o n a l b e n e f i t s .

Also, l i g h t w e i g h t h i g h - s t r e n g t h composites show c o n s i d e r a b l e promise. When t h e s e o u t s i d e d r i v e r s and technology o p p o r t u n i t i e s a r e a p p l i e d t o f u t u r e ad- vanced technology p r o p e l l e r s , t h e d e s i g n may be a l t e r e d t o o p t i m i z e performance t o a lower n o i s e g o a l u s i n g l i g h t w e i g h t composite b l a d e s of advanced shape.

T h i s may r e s u l t i n a small p r o p e l l e r c o s t i n c r e a s e ; however, t h i s p e n a l t y s h o u l d be more t h a n overcome by t h e l a r g e p o t e n t i a l performance and weight advantages.

An advanced technology l o r s p e e d p r o p e l l e r may resemble t h e d e s i g n d e p i c t e d i n t h e s k e t c h on t h e upper r i g h t o f f i g u r e 5. Some o f t h e advanced f e a t u r e s i n - c o r p o r a t e d i n t h i s d e s i g n a r e b l a d e sweep, p r o p l e t s ( p r o j e l l e r t i p d e v i c e ) , ad- vanced a i r f o i l s , composite b l a d e s , and improved i n t e g r a t i o n o f t h o p r o p e l l e r and n a c e l l e .

The NASA r e s e a r c h program t h a t a d d r e s s e s t h e p r o j e c t e d technology t r e n d s o f Lou-speed GA p r o p e l l e r s i s summarized i n f i g u r e 6. A c o s t - b e n e f i t staldy t h a t e v a l u a t e s t h e e f f e c t of advanced t e c h n o l o g i e s o n l o r s p e e d p r o p e l l e r s i s b e i n g conducted by t h e HcCauley Accessory Dilvision of t h e Cessna A i r c r a f t Company.

The NASA I . m q l e y Research C e n t e r i s sponsoring two g r a n t programs o n reducing p r o p e l l e r sc~ilrce n o i s e . M L T is e v a l u a t i n g s e v e r a l approaches f o r reducing n o i s e I n a complementary program Ohio w h i l e o p t i m i z i n g performance ( r e f s . 11 and 12).

S t a t e U n i v e r s i t y is e v a l u a t i n g e n a l t e r n a t i v e n o i s e r e d u c t i o n approach i n a f l i g h t t e s t program. I n t h e p r o p e l l e r performance a r e a , Purdue U n i v e r s i t y , un- d e r a NASA g r a n t , is involved i n a program t o i n v e s t i g a t e s e v e r a l advanced con- c e p t s t h a t show c o n s i d e r a b l e p o t e n t i a l f o r improving p r o p e l l e r performance. I n a d d i t i o n , t h e c u r r e n t low-speed p r o p e l l e r a e r o a c o u s t i c d e s i g n methodology is being e v a l u a t e d and enhanced through a c o o p e r a t i v e program between Ohio S t a t e U n i v e r s i t y and Lewis.

General-aviation p r o p e l l e r s a r e u s u a l l y designed a s s e p a r a t e p r o p u l s i o n components w i t h o u t p r o p e r l y a c c o u n t i n g f o r t h e aerodynamic i n t e r a c t i o n between t h e n a c e l l e ( o r f u s e l a g e ) and t h e p r o p e l l e r .

A program is underway a t Missis- s i p p i S t a t e U n i v e r s i t y t o develop technology t h a t w i l l a l l o w t h e p r o p e l l e r and n a c e l l e t o b e designed i n a more u n i f i e d approach. T h i s program s h o u l d l e a d t o improved o v e r a l l p r o p u l s i o n system performance f o r g e n e r a l - a v i a t i o n a i r c r a f t .

F r o p e l l e r dynamics and a e r o e l a s t i c s c a n b e s e r i o u s d e s i g n l i m i t a t i o n s f o r b o t h c u r r e n t and f u t u r e advanced low-speed p r o p e l l e r s . To b e t t e r understand t h i s important a r e a , NASA is sponsoring a r e s e a r c h program a t Pennsylvania S t a t e U n i v e r s i t y ( r e f . 13). More d e t a i l e d i n f o r m a t i o n on NASA's low-speed p r o p e l l e r r e s e a r c h i s c o n t a i n e d i n i n d i v i d u a l papers g i v e n by K e i t e r , Green, Korkan, and McCormick a t t h i s conference. A summary o f t h e Purdue U n i v e r s i t y r e s e a r c h pro- gram on advanced performance concepts i s shown i n f i g u r e 7. T h i s program, under t h e d i r e c t i o n of D r . John S u l l i v a n , i n c l u d e s a n a l y t i c a l and e x p e r i m e n t a l re- s e a r c h on such new c o n c e p t s a s p r o p l e t s ( t i p d e v i c e s ) and swept b l a d e s ( r e f .

14). A new swept l i f t i n g l i n e a n a l y s i s program i s b e i n g developed a t Purdue, and t h e y a r e v e r i f y i n g t h i s a n a l y s i s and d e t e r m i n i n g t h e perf onnance p o t e n t i a l of advanced c o n c e p t s through s u b s c a l e p r o p e l l e r wind-tunnel t e s t s u s i n g some improved t e s t techniques. These new t e c h n i q u e s i n c l u d e t h e l a s e r v e l o c i m e t e r system shown i n f i g u r e 7.

The approach being used by M i s s i s s i p p i S t a t e U n i v e r s i t y t o d e v e l o p improved They have con- p r o p e l l e r - n a c e l l e i n t e g r a t i o n technology i s shown i n f i g u r e 8.

ducted a n e x t e n s i v e s e a r c h o f t h e p r o p e l l e r l i t e r a t u r e t o a s s i s t i n e v a l u a t i n g p o t e n t i a l l o s s mechanisms and t o s e e which o f t h e r e s u l t s a p p l y t o modern GA p r o p e l l e r - n a c e l l e geometries. M i s s i s s i p p i S t a t e a l s o p l a n s t o conduct f u l l - s c a l e p r o p e l l e r wind-tunnel t e s t s ( i n t h e Langley f u l l - s c a l e t u n n e l ) i n combin- a t i o n w i t h a n a l y t i c a l s t u d i e s t o d e v e l o p t h e improved p r o p e l l e r - n a c e l l e i n t e g r a - t i o n technology. T h i s r e s e a r c h should l e a d t o b e t t e r o v e r a l l p r o p u l s i o n system performance f o r g e n e r a l - a v i a t i o n propeller-powered a i r c r a f t .

I n a d d i t i o n , t o t h e advanced, high-speed p r o p e l l e r wind-tunnel t e s t progr-m d i s c u s s e d e a r l i e r , N A S A i s a l s o t e s t i n g lower speed GA p r o p e l l e r s . A 5-foot- d i a m e t e r model of one of t h e s e low-speed p r o p e l l e r s i s shown i n f i g u r e 9. T h i s p r o p e l l e r , along t h e w i t h t h r e e o t h e r d e s i g n s , was t e s t e d i n t h e Lewis 10- by 10-foot wind t u n n e l t o compare measured performance w i t h a n a l y t i c a l p r e d i c t i o n s .

R e s u l t s from t h i s comparison w i l l be used t o d e v e l o p enhanced a n a l y t i c a l pre- d i c t i o n procedures f o r t h i s c a t e g o r y of p r o p e l l e r .

CONCLUDING REMARKS The world energy s h o r t a g e h a s l e d t o t h e need f o r more f u e l e f f i c i e n t GA a i r c r a f t . P r o p e l l e r p o w e r e d p r o p u l s i o n , w i t h i t s i n h e r e n t high l e v e l of e f f i - c i e n c y , remains a n a t t r a c t i v e p r o p u l s i o n concept. C u r r e n t GA a i r c r a f t a r e l i m i t e d t o maximum c r u i s e s p e e d s n e a r Mach 0.5 because o f p r o p e l l e r compressi- b i l i t y l o s s e s . The NASA r e s e a r c h program on t h e s e lower speed p r o p e l l e r s o f f e r s t h e p o t e n t i a l o f s i g n i f i c a n t performance improvements and n o i s e r e d u c t i o n s through t h e development of advanced c o n c e p t s and new a n a l y t i c a l d e s i g n proce- dures. Extending t h e c u r r e n t c r u i s e speed l i m i t a t i o n t o a t l e a s t Mach 0.8, w i t h a l a r g e p o t e n t i a l f u e l s a v i n g compared w i t h t u r b o f a n powered a i r c r a f t , may be p o s s i b l e w i t h t h e t e c h n o l o g i e s t h a t a r e being devoloped under NASA's Advanced Turboprop P r o j e c t .

REFERENCES 1. C o r s i g l i a , V. C. ; Katz, J. ; and Kroeger, R. A. : F u l l S c a l e Wind Tunnel Study o f Nacelle Shape on Cooling Drag. A I A A Paper 79-1820, Aug. 1979.

2. Mikkelson, Daniel C . ; e t a l . : Design and Performance of Energy E f f i c i e n t P r o p e l l e r s f o r Mach C.8 Cruise. N A S A T M X-73612, 1977.

3. Dugan, James F., Jr. ; Gatzen, Bernard S. ; and Adamson, William M. : Prop-Fan Propulsion - I t s S t a t u s and P o t e n t i a l . SAE Paper 780995, Nov. 1978.

4. J e r a c k i , R. J.; Mikkelson, D. C . ; and S l a h a , B. J.: Wind 2unnel Performance of Four Energy E f f i c i e n t P r o p e l l e r s Designed f o r Mach 0.8 Cruise. SAE Paper 790573, Apr. 1979.

5. Metzger, F. B. ; and Rohrback, L.. : A e r o a c o u s t i c Design o f t h e Prop-Fan. A I A A Paper 79-0610, Mar. 1979.

6. C o r n e l l , R. W . ; and Rothman, E. A.: S t r u c t u r a l Design and A n a l y s i s of Prop-Fan. AIAA Paper 79-1116, June 1979..

7. R e v e l l , J . D.; Balena, F. J.; and Koval, L. K . : A n a l y t i c a l Study of I n t e r - i o r Noise Controi by Fuselage Design Techniques o n High Speed P r o p e l l e r - Driven A i r c r a f t . (Lockheed-California Company, NASA C o n t r a c t NASl- 15427.) NASA CR-159222, 1980.

8. Rennison, D. C.; Wilby, J. F.; and Marsh, A. H.; Wilby, E.G.: I n t e r i o r Noise C o n t r o l P r e d i c t i o n Study f o r High-Speed Propeller-Driven A i r c r a f t . ( B o l t Beranek and Neman, Inc.; NASA C o n t r a c t NAS1-15426.) N A S A CR-159200, 1980.

9. Bencze, D. P.; e t a l . : P r o p e l l e r S l i p s t r e a m Wing I n t e r a c t i o n s a t Mach No.

0.8. SAE Paper 780997, Nov. 1978.

10. S t o l p , P h i l i p C. and Baum, James A . : Advanced Turboprop P r o p u l s i o n System R e l i a b i l i t y and Maintenance Cost. SAE Paper 771009, Nov. 1977.

11. S u c c i , G. P. : Design of Q u i e t Ef f i c i e n : P r o p e l l e r s . SAE Paper 790584, A p r .

1979.

12. L a r r a b e e , E. E.: P r a c t i c a l Design of Minimum Induced Loss P r o p e l l e r s . SAE Paper 790585, Apr. 1979.

13. McCormick, B. W.; e t a l . : The A n l a y s i s of P r o p e l l e r s I n c l u d i n g I n t e r a c t i o n E f f e c t s . SAE p a p e r 790576, Apr. 1979.

14. S u l l i v a n , J.: The E f f e c t of Blade Sweep on P r o p e l l e r Performance. AIAA p a p e r 77-716, J u l y 1978.

LOW SPEED PROPELLERS - IMPACT OF ADVANCED TECHNOLOGIES

Ira D. K e i t e r McCauley Aaessory O i v i s h Ccssna Aircraft Company Studies have indicated that Zhe application of advanced technologies to General Aviation propellers c i a reduce fuel consumption in future aircraft an average of 10 percent, meeting current FAR Part 36 noise limits. Through the use of composite blade construction, up to 25 percent propeller w e i g h : reduc- tion can be achieved. This weight reduction in addition to 7 percent propeller efficiencv improvements through application of advanced technologies result in 4 percent reduction in direct operating wsts, 10 percent reduction in aircraft acquisition cost, and 7 percent lower gross weight for General Aviation aircraft.

INTRODUCTION In order to insure that USA built General Aviation aircraft reaain competitive and dominant in the world market place, support energy conservation needs, and meet the more stringent environmental controls, NASA sponsored programs are necessary to isprove propeller technology based for the most part on that developed during the World War XI era. Attention to the area of materials used and fabrication methods is the conerstone leading to the pursuit of advanced technology concepts and sophisticated computer analysis tools to evaluate those concepts.

Preliminary indications are that proper techniques could be developed with the utilization of composite mzterials in the structure of propeller assemblies. A proper blend of design and fabrication techniques will result in significant weight and cost reductions, enhanced safety through improved fatigce life, greater adaptability to a variety of design concepts and less capital requirements to produce propellers suitable to the General Aviation market.

n e use of lighter weight blades will permit both the increase in blade retention hardware safety margiiis and the reduction in weight and complexity of such hardware.

The combination of improverents in cost and weight reduction, fatigue life increases, more consistently produced airfoil sections, and more widely varied potential design selection has a significantly broadening effect on typical installation compromises which will be apparent as the potential impact of the various advanced technologies are enumerated in later sections of this paper.

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It will be shown that significant propeller weight improvements are possible for current propeller/engine/aircraft installations while achieving improveaents in performance and reduced noise, both within the aircraft and in the airport environment.

The potential propeller weight reductions and efficiency improvements can directly result in reductions in aircraft fuel consumption, direct operating costs, acquisition cost, and lower gross weight.

Iaprovements in propeller fatigue safety margins would permit correspond- ing increases in overhaul life, provide greater tolerance for field mainteriance or lack of same and thus increase safety, productivity, and economy.

The currer~t research program awarded to kCauley is needed to structure a realistic and effective technology plan for General Aviation propellers, to identify the advanced technologies and their potential costs and benefits, and to determine and identify areas of key technical risks and required research programs (Figure 1 ) . It is hoped that once the areas of greatest potential are identified, NASA research funding can then be channeled into the most appropri- ste areas. The team comprising the current study effort (contract number NAS3-21719) is highlighted in Figure 2 .

SENSITIVITY STUDIES Sensitivity studies were perforated to evaluate the potential of each technology element on propeller performance, noise, weight, and cost for the following two categories of General Aviation aircraft: I. Low speed (up to 250 knots)/low power (up to 350 HP) , 2 - 8 place.

single and twin engine, currently powered by reciprocating engines.

11. High speed (up to 400 knots)/high power (up to 650 HP) , 6 - 18 place, twin engine, turboprop.

In order to evaluate the improvement potential for the various technology elements, specific aircraft satisfying the characteristics outlined above for each category were chosen. The representative aircraft, illustrated in Figure 3, chosen for each aircraft category, are as follows: Category I : Cessna 172N Cessna 21a1 Cessna 414A Category I1 : Cessna 331 TECHNOLOGY ELEMENTS INVESTIGATED Technology elements were considered with the potential for improving airplane/mission characteristics such as fuel burned, direct operating cost, acquisition cost, and gross weight.

The original scope of technology elements considered were screened to include any element felt to have any potential whatsoever based on McCauley experience, other's experience, and available literature. Each technology element was first evaluated in terms of its effect on propeller criteria of merit including performance, noise, weight, cost, and structural considerations. The total impact of the technology elements affecting performance and weight were then investigated to determine their impact on airplane/mission characteristics.

The General Aviation propeller concept incorporating the appropriate advanced technologies is illustrated in Figure 4.

PERFORMANCE CONSIDERATIONS The following list of technology elements show potential for performance improvements. The element is specified along with the loss to be minimized.

keeping in mind the practical limitations of each of the technology elements and their impact on other important propeller criteria of merit such as noise, weight, and structure.

Technology Element Performance Loss Reference To Be Minimized dtilized Desi~n optimization Decreased power loading Axial momentum 1 Increased number of blades 1 Tip Maintain tip speed Axial r r r o m e n t u n l 1 Decreased activity factor Profile Purdue Univ.

Use of proplets Tip Use of sweep (helical tip mach

number reduction) Compressibility - 7

Advanced technoloey airfoil type Compressibility, profile 4 Decreased thickness ratio Profj le 2 Improved propeller/nacelle integration Blade profile, nacelle drag 2 . 3 Improved surface finish Profile 5,6,7 Maintainability of airfoil shape Prof i 1 e 7 The advanced technology concept of swee? is unique since it both improves performance and reduces noise. The predicted effect of sweep on performance and noise is shown in Figure 5 . Sweep is structure limited with 25-30 degrees probably feasible for future General Aviation applications. The effect of sweep in addition to the effect of other significant advanced technology ele- ments on cruise performance gains are shown in Figure 6. Power loading, number of biades, tip speed, activity factor, and proplets are grouped together and classified as design optimization parameters.

Maintainability of airfoil shape is not included since it is a parameter to prevent degradation through the use of composites and not to improve current technology.

ACOUSTIC CONSIDERATIONS The primary technology elements affecting aco~stics are number of blades, tip speed, thickness ratio, activity factor, sweep, blade loading, advanced technology airfoils, and proplets.

The following is a list of these primary technology elements: Technology Element References Utilized Design Optimization Increased number of blades 8 Decreased tip speed Decreased activity factor Use of proplets Purdue University Peak blade loading moved inboard 8 Use of sweep (helical tip mach number reduction) 8 Advanced technology airfoil types Decreased thickness ratio As with performance considerations, several elements are grouped together under design optimization. From the acoustic standpoint, all items other than advanced airfoils, sweep, and reduced thickness ratio are considered design optimization variables. The main technology elements affecting acoustics, including sweep whose effects were isolated earlier, are shown in Figure 7 .

These gains can be realized without any noticeable loss in performance. The delta d B ( A ) improvements possible are in many cases more than that required to meet noise regulations. During design tradeoff studies, the relative importance of each design parameter must be evaluated. Greater noise reduc- tions could be obtained if one were willing to sacrifice performance.

MATERIAL CONSIDERATIONS In order to reliably meet the future performance and acoustic require- ments of General Aviation propellers with weight reduction and approaching price competitiveness with aluminum, consideration of composite materials requires appropriate attention as a viable solution. Advanced filamentary composite materials combine low densities and low notch sensitivity with high strengths and stiffnesses. Adequate safety margins of current propellers can be further enhanced. Figure 8 outlines the advantages of composites and their associated propeller benefits. Because filamentary materials are only strong in the filament direction, careful consideration must be given to ply orientation to match the design requirements.

Through variations in the composite matrix, blade sections can be tailored to meet the specific radial stiffness distribution required.

The shape of primary bending and torsional modes can be altered effectively through the use of composites. Reductions in blade section size permissible with composites will result in higher blade deflections than are customary with aluminum.

Blade aeroelastic instabilities can result from large out of plane deflections and must be given careful consideration.

Appropriate blade materials, type of hub retention system, methods of construction for composite materials, and material consideration for blade leading edge erosion resistant strips are all areas which must be addressed in detail.

STRUCTURAL CONSIDERATIONS In evaluating the structural ifitegrity of advanced technology propellers, considerable attention must be given to the steady and alternating loadings experienced in service. The steady loads consist primarily of centrifugal, bending due to thrust loading, and torsion. The alternating vibratory loads are due to blade aerodynamic excitations and alternating torsional input due to reciprocating engine cylinder firing sequence and frequency. Aerodynamic inflow angles excite ixP alternating loads which are primarily evident on turboprop installations being overshadowed by engine alternating torsionals in reciprocating installations. In all installations one should assure that 1xP resonance does not occur in the normal operating RPbl range.

It is a relatively easy job to evaluate the steady loads on a propeller blade using conventional techniques. To determine the vibratory effects, however, with incorporation of advanced technologies such as sweep and proplets, may require the use of three dimensional finite element analysis rather than two dimensional beam analysis or lumped parameter matrix manipulation techni- ques currently utilized. With regard to vibratory analysis, available analytical techniques applicable to General Aviation propellers determine mode frequencies with good accuracy and vibratory loads and resultant stresses within 25-30 percent on turbine installations. The effect of alternating torsionals from reciprocating engines is currently not included in existing models. Experimental testing of strain gaged propellers is still relied upon heavily. Experience and experimental data will dictate allowable alternating stress levels with composites as is the case kith aluminum alloys.

Using the torsional mode results from three dimensional finite element analysis (3-D FEA), the possibilities of stall flutter can be addressed. The stall flutter parameter is based on static torsional frequency, and semichord, velocity, and mach number at the SO percent blade radius location. Through extensive experimental pl-cgrams, a stall flutter boundary has been determined for conventional blade shapes. A similar boundary must be determined for blade shapes incorporating the advanced technology concepts outlined in Figure 4.

S t a l l f l u t t e r occurs under conditions of blade angle o f attack and inflow velocity where a major portion of t h e blade is s t a l l e d .

S t a l l f l u t t e r o s c i l l a t i o n s occur a t t h e first torsional mode when t h e spanwise damping i n t e g r a t i o n along a blade becomes zero o r less.

The cocditions conducive t o s t a l l f l u t t e r a r e during s t a t i c , takeoff, and reverse t h r u s t operation.

By using t h e bending and t o r s i o n a l mode data from 3-D FEA and customarily presented i n terms of a Campbell diagram, c l a s s i c a l f l u t t e r can be addressed.

Classical f l u t t e r can occur a t high a i r c r a f t v e l o c i t i e s where t h e mod2 spacirig over t.he operating RPM range is i n s u f f i e n t and a coupling of t o r s i o n a l and bending modes occur (reference 9) .

Because of t h e limited composite f a t i g u e strength d a t a a v a i l a b l e and t h e lack of a n a l y t i c a l techniques t o predict vibratory loads, t h e evaluation of f a t i g u e l i f e is highly q u a l i t a t i v e .

Only through extensive test programs and f i e l d experience, can t h e required data base of information be compiled from which t h e appropriate f a t i g u e limits can be determined.

This same process occurred many years ago t o e s t a b l i s h t h e current baselines u t i l i z e d f o r aluminum.

BLADE WEIGHT, COST, AND AIRCRAFT MISSION CONSIDERATIONS Preliminary screening of candidate materials indicates configurations of E-Glass, S-Glass, Kevlar, and Craphite with medium and high density epoxy cores t o meet t h e mean load, a l t e r n a t i n g load, fatigue, and weight requirements o f Generxl Aviation ~ r - p e l l e r s . Relative blade weight and c o s t comparison against aluminum a r e shown i n Figure 9. In determining t h e impact of weight reducticils through t h e use of composites it must be emphasized t h a t t h e blade weight savings e x i s t only with a d i r e c t replacement of aluminum blades. This does not take i n t o account t h e blade r e t e n t i o n area. Also, i n order t o achleve a desired compromise of advanced technologies between performance and nuise, t h e p o t e n t i a l weight savings may be reduced. In o t h e r words, t h e trends of decreased power loading through diameter increases, increased nmber of blades, sweep, and proplets w i l l tend t o increase weight while being o i f s e t through lower blade a c t i - ~ i t y f a c t o r s and lower thickness r a t i o s ( f e a s i b l e because of composites).

The performance gains indicated e a r l i e r i n addition t o weight reductions possible through t h e use of composites have a d i r e c t e f f e c t on a i r c r a f t / mission c h a r a c t e r i s t i c s such a s fuel burned, operating c o s t , acquisition c o s t , and gross weight. In addressing mission analysis, payload, range, speed and Potential t r i p f u e l savings a i r c r a f t l i f t t o drag r a t i o a r e kept constant.

versus a i r c r a f t c r u i s e speed a r e shown i n Figure 10. This assumes two hours a t c r u i s e , f e l t t o be f a i r l y representative. Studies i n d i c a t e t h a t average t r i p fuel reductions of about 10 percent r e s u l t i n 4 percent reductions i n d i r e c t operating costs. Included in DOC determination a r e engine and airframe periodic maintenance, fuel and o i l burned, reserves f o r engine and propeller A s f u e l p r i c e s overhaul, reserves f o r avionics, systems and miscellaneous.

r a i s e i n t h e future, t h e i r e f f e c t on increases i n DOC w i l l be a s indicated i n Eigure 11.

Aircraft acquisition cos i reduct ions average about iO percent.

Reductions as affected by aircraft cruise speed are indicated in Figure 12.

A twenty-five percent increase in propeller cost has been taken l ~ t o account but does not alter the results since the propeller cost is so low in relation to aircraft cost.

Studies also indicate average potential aircraft gross weight reductions of seven percent.

FUTURE RECOWENDATIONS It is apparent from the sensitivity studies performed on the various technology elements that NASA funding directed primarily into the areas of composite materials research, and the advanced technology concepts outlined in this paper can provide the data base required to achieve the stated airplane/ miss ion improvement s .

Since many technology elements improving performance have an adverse effect on acoustics and future government regulations controlling noise limits will probably be more stringent, it is imperative that research funding be expended in this area.

171is should include careful evaluation of current methodology of propeller noise prediction techniques and the unification into a common, recognized procedure for utilization by the General Aviation community. Experimental verification of resulting theories through wind tunnel testing and flight substantiation is necessary.

Through the use of composites, some of the more promising technology requirements will become possible such as thickness ratio reductions, sweep, lower activity factors, reduced power loading, more blades, advanced airfoils with complex curvature, smoother airfoil surface, and maintainability of air- foil shape in service. These factors lead to optimized designs meeting appropriate strength requirements. At a certain blade load level, the This can lead to aeroelastic instabilities composite blade will deflect more.

Although considerable which is an important area requiring NASA support.

research has taken place in the composite materials area, the product applica- tions have not included propellers. The aircraft propeller is one of the most critically stressed aircraft components. It operates in a severe environment It is and is a major structural component with complex stress distribution.

exposed to the wide range of variables created by power plants in a most intimate manner. The successful use of composite materials in General Aviation propellers will provide information on fatigue to establish limit lines of mean stress versus alternating stress for lo5 and 108 cycles as typically represented on a Goodman diagram. Such information is not readily obtainable in any other application.

NASA sponsored research will help fill the gap in the application of composite technology (design and fabrication methods) between current applica- tions and their potential use with propellers. Advanced composites technology has progressed to a point where reliable application as aircraft secondary structure is accepted and the application for primary structure is relatively close but the confidence level for commercial application has not been established. It is, therefor,:, highly desirable and appropriate to explore this General Aviation application.

Specifically, the total cost of composite blades must be nearly competi- tive with aluminum blades in order to experience wide use in General Aviation.

Research into low cost fabrication techniques is the key to achieving cost competitiveness. Wind tunnel testing with follow-on programs for flight test verification is necessary.

In the advanced airfoil design area, NASA has continually made efforts in improving communication with the General Aviation community over the past five years through workshops, symposiums, conferences, etc., and from these have come airfoil design implementation schedules satisfying the needs of wing designers. What is needed now is an airfoil technology plan to design airfoils specifically tailored for the widely varying fluid flow conditions which prevail along a propeller blade.

The area of propeller/nacelle integration is already receiving some attention by NASA with Grant NSG1402 to Mississipi State University. The first phase involving the collection of baseline data in the NASA-Langley full scale tunnel is just getting underway.

Current program calls for investigation of nacelle shapes characteristic of those used in twin reciprocating engine installations. There should be future testing including a wide variety of propeller/nacellc configurations covering the broad range of aircraft/engine combinations tnisal of the General Aviation fleet.

NASA has supported analytical studies to provide special purpose user oriented programs to calculate propeller inflow velocity fields, steady and unsteady aerodynamic loads and mode shapes and frequencies. The next step should be to concentrate on analytical procedures to predict the vibrational loads and stresses the propeller is subjected to in service. The model must include the coupling effects of the propeller-engine system. Accuracy of existing prediction techniques is not acceptable except for 1xP analysis.

Higher order stresses A malent with reciprocating engine installations are not predicted with adequate accuracy. A good, reliable prediction technique would eliminate much of the uncertai~:~ which exists prior to vibration survey certification testing. Wasted time and cost associated with experimental test- ing of a configuration exceeding vibratory load limits could be nearly eliminated .

With the very competitive market in General Aviation which lir,~its funds in research and development, it is very evident that NASA sponsored support is necessary to enhance the state of the art in the areas mentioned above.

The key areas requiring future research enumerated above are highlighted in Figure 13.

CONCLUDING REMARKS The study of a wide range of propeller design variables and advanced technologies has indicated that the potential exists for propeller performance improvements and weight reductions meeting consistently more stringent regulatory noise levels. Advanced technological development of propellers has a direct impact on the fuel burned, direct operating costs, acquisition cost, and gross weight of General Aviation aircraft. NASA can assure that these goals are met by allocating appropriate funds in the areas where the greatest potential exists.

REFERENCES 1. Anonymous: SBAC-Standard Method of Propeller Performance Ectimation, Society of British Aircraft Constructors, Ltd. April i950.

2 . Goldstein, S . : On The Vortex Theory of Screw Propellers. Royal Society (London) Proc . , 1929.

3 . Hess, J.L.; and Smith, A.M.O.: Calculation of ?otential Flow About Arbitrary Bodies. Progress in Aeronautical Sciences, vol. 8, 1967, pp. 1-138.

4. Bocci, A.J. : A New Series of Aerofoil Sections Suitable For Aircraft Propellers. Aeronautical Quarterly, vol. 28, Part 1, Feb. 1977, pp.

59-73.

5. Hoerner, S.F.: Fluid-Dynamic Drag. 1965.

6 . Hoerner, S.F.; and Borst, H.V.: Fluid-Dynamic Lift. 1975.

7 . Stevens, W.A.; Goradia, S.H.; and Braclen, J.A.: Mathematical Fiodel For Two- Dimensional Multi-Component Airfoils In Viscous Flow. NASA CR-1843, July 1971.

8. Succi, George P.: Design of Quiet Efficient Propelless, SAE Payer 790584, SAE Business Aircraft Meeting, Wichita, April 1979.

9. Dugan, J.F. Jr.; Gatzen, B.S.; and Adamson, W.M.: Prop-Fan Propulsion-Its Status alrd Potential. SAE Paper 780995, SAE Aerospace Meeting, San L'iego, November 1978.

AWANCED TECHNOLOGY PROPELLER STUDY

IDENTIFY ADVANCED TECHNOLOGIES

ASSESS BENEFITS . COSTS & RISKS

DEFINE OPTIMUM CONFIG., MISSION ANALYSIS

FORMULATE RESEARCH PROGRAM

FIG. I

TEAM COMPRISING STUDY EFFORT

McCAULEY

PROGRAM MANAGEMENT, PERFORMANCE , COST, STRUCTURES

CESSFA

AIRPLANE MISSION ANALYSIS

OHIO STATE

ACOUSTICS

MATERIALS SCIENCES b S A I

COMPOSl T E S

FIG. 2

33 7 ADVANCED TURBOPROP POTENTIAL FOR HIGH SPEED

Bernard S . Gatzen

Hamilton Standard Division, United Technologies Corporation Over t h e last f o u r y e a r s , t h e r e h a s been a s i g n i f i c a n t amount o f t e c h n i c a l p r o g r e s s on an advanced p r o p u l s i o n concept c a l l e d t h e Prop-Fan. I want to d i s c u s s t h i s p r o g r e s s and what it might mean r e g a r d i n g high speed b u s i n e s s The Prop-Fan p r o p u l s i o n system is a n advanced p r o p e l l e r d r i v e n a i r c r a f t .

The r o t o r technology is being pursued i n N A S A ' s by a t u r b o s h a f t e ~ g i n e .

ACEE program. Advancements i n aerodynamics, a c o u s t i c s , s t r u c t u r e s , and mechanical 2xecution a r e involved and I ' l l touch on each o f t h e s e s u b j e c t s .

The major b e n e f i t a t t r i b u t e d to t h e Prop-Fan is f u e l s a v i n g s r e s u l t i n g from improved perfor~i'ance. T h i s p r o p u l s i o n d e v i c e is aimed a t a i r s p e e d s above 0.55 Mach where t h e t u r b o f a n is t o d a y ' s s t a n d a r d . T h e r e f o r e , t h e key o b j e c t i v e is h i g n performance a t high s u b s o n i c Mach number which r e s u l t s i n f u e l saved and reduced o p e r a t i n g costs. However, i n a c h i e v i n g t h i s g o a l , c e r t a i n o t h e r c h a i a c t e r i s t i c s must n o t be s a c r i f i c e d . It is n e c e s s a r y to remain a good neighbor around t h e a i r p o r t , it is n e c e s s a r y to have a c o m f o r t a b l e i n t e r i o r environment, and t h e a i r c r a f t s a f e t y must be uncomprmised. A l l o f t h e s e c o n s t i t u t e t h e g o a l s being worked towards i n t h e Prop-Fan programs being conducted by NASA w i t h Hamilton Standard p a r t i c i p a t i o n .

T y p i c a l l y t h e turboprop on a b u s i n e s s a i r c r a f t o p e r a t e s below 0.55 kkch, more i n t h e area o f 0.45 Mach, w i t h v e r y h i g h e f f i c i e n c i e s . The E l e c t r a turboprop had high e f f i c i e n c i e s up t o 0.6 Mach. Beyond t h e s e a i r s p e e d s , t h e e f f i c i e n c y d r o p s o f f and t h e t u r b o f a n is king. The p r o d u c t i v i t y o f a 0.8 Mach t u r b o f a n v e r s u s a 0.5 Mach turboprop h a s c l e a r l y k e n demonstrated i n t h e commercial passenger c a r r y i n g market. With t h e Prop-Fan, t h e i n h e r e n t l y h i g h p r o p e l l e r type e f f i c i e n c i e s are extended o u t to t h e t u r b o f a n o p e r a t i n g regime o f 0.8 Mach. With core e n g i n e s o f comparable technology, t h e Prop-Fan o f f e r s s i g n i f i c a n t f u e l s t v i n g s o v e r t h e e n t i r e Mach number range up to 0.85. i have g e n e r a l i z e d t h e p r o j e c t e d f u e l s a v i n g s w i t h Prop-Fans i n p l a c e o f t u r b o f a n s based on a v a r i e t y o f a i r c r a f t s t u d i e s conducted by Boeing, Douglas, both Lockheeds, P r a t t & Whitney, General E l e c t r i c , NASA and United Technologies. I t shows maximum f u e l s a v i n g s a t s h o r t and v e r y l o n g o p e r a t i n g ranges. For s h o r t range a i r c r a f t , t h e mission is dominated by c l i m b and d e s c e n t , g e n e r a l l y keeping a i r s p e e d below d e s i g n c a p a b i l i t y . A t v e r y s h o r t range, it shows t h a t a high d e s i g n speed is n o t r e l e v a n t t o f u e l usage. There is a bucket a t 1500 t o 2500 ~ a u t i c a l m i l e s with i n c r e a s i n g f u e l s a v i n g s a t l o n g e r ranges. Here, t h e f u e l s a v i n g is reducing g r o s s weight and a compounding e f f e c t t a k e s p l a c e , Another a t t r i b u t e of t h e Prop-Fan is t h a t f u r t h e r f u e l s a v i n g s a r e achieved by j u s t reducing a i r s p e e d ; t h i s is n o t t h e c a s e w i t h t h e t u r b o f a n .

F'fKC€?'?!C 7 ' : X 2:A:iK NOT FILMED Let's compare the general characteristics of this advanced Prop-Fan with the typical business aircraft turboprop. I have already mentioned the significant difference with intended airspeed range. Coupled with this is operating altitude; the higher the Mach, the higher the altitude. Typically propellers are operated in cruise at maximum efficiency which translates into lower power loadings. This generally works out well for low speed aircraft which end up being sized for climb conditions.

As design Mach is increased, the propeller becomes sized by cruise and selecting a power loading for peak efficiency would result in very large diameters.

So, some efficiency is sacrificed for size and a higher power loading is selected.

The blade count is increased to maintain high cruise efficiency levels and good low speed performance as well. The rpm's are similar but Prop-Fan favors the lower side to minimize the transonic effects as airspeed is increased.

In both aerodynamics and acoustics, it is necessary to design for operation ~ . f a large portion of the rotor area in the transonic regime. 800 fps at 0 . 2 Elach results in a tip helical Mach of 1.14. The key to maximizing ie-dynamic efficiency is to eliminate compressibi!ity losses and this is done by using the thinnest airfoils possible, consistent with structural integrity. Blade sweep is used in a manner similar to wing sweep to reduce the effective Mach number the airfoil section operates at.

Nacelle shaping is accomplished to reduce the velocity through the disk with the primary emphasis to controlling inboard root choke, Area ruling is used in conjunction with nacelle shape to minimize drag losses in the root where Lastly, an advanced airfoil shape may allow thicker the solidity is high.

airfoils or less sweep by raisi~g the section critical Mach number above the point where compressibility losses rise rapidly. Three generations of model Prop-Fans have been designed, fabricated, and tested. Advanced aerodynamic analyses have been developed from existing propeller and fan ex,pertise and applied to the design tasks. The aero analysis treats the inboard blade area as a cascade like a turbofan and the outboard area like a propeller. Only compressible 2D airfoil data are used. Both the effects of the nacelle and blade sweep are accounted for. with advanced design methods, achieving aerodynamic efficiency improvement at high Mach is accomplished with a methodical approach and a high confidence of success in place of the emperical, that is, cut and try approach.

The advanced aero design methods are allowing a fairly rapid focussing towards designs which improve upon Electra turboprop efficiencies. Of course, you can see that even t i . s 1950's turboprop technology used in the Electra, which is the sane as that used on the P3, is far superior to the typical general aviation turboprop. This is a direct result of the oper- ating Mach number. The general aviation turboprops are gesigned to be highly efficient at 0.4 to 0.5 Mach cruise. The Electra technology offers no improvement there, but holds efficiency up high out to 0.6 Mach where it begins to drop off rapidly. Prop-Fan technology offers small gains at 0.6 Mach but holds efficiency up high out to 0.85 Mach. The Electra achieves efficiency improvement over general aviation turboprops through lower blade thickness ratios, both at the blade tips and roots. The Thick- Prop-Fan makes use of all the aerodynamic concepts discussed earlier: ness ratio, blade sweep and nacelle shaping. Wind tunnel test results on two-foot model Prop-Fans have demonstrated 80% efficiency which is the .Additional aero/acous tic designs are expected to program objective.

em on strati on of this should improve upon these results by a few percent.

occur next year. Further improvements in efficiency can b e obtained by As much as eight efficiency points are lost in the recovering swirl.

slipstream of a single rotation rotor. Portions of this are recoverable by having the wing act as stator vanes or by using a counter-rotating propeller.

As with aerodynamics, controlling the acoustic levels required advacced concepts and design methodologies. Both far field noise around the air- port and interior noise in cruise must be controlled. Concepts which have been considered in acoustic analysis conducted to date include reduced thickness ratio, reduced tip speed with increased blade count, optimum blade planform including swept shapes, and advanced airfoils.

Let's consider their impact on far field noise. Noise reductions in the far field have typically been achieved by reducing tip speed; this is a very powerful noise reducing means, but in the past has been accompanied by performance reductions as well. This traditional means of noise reduction can be accomplished without performance decrement by increasing rotor diameter, but this is usually an unattractive alternate. Studies with new acoustic design methodology indicates that significant noise reductions can be achieved with improved airfoils, higher blades count, and a more optimum blade shape. Both airfoil optimization and increasing the blade count have a compound impact on noise by reducing noise in themselves but also improving efficiency so that for constant thrust, a smaller diameter can do the same job. Improved blade shapes, including sweep, reduce noise significantly without diameter changes.

The advanced acoustic method mentioned above is a procedure developed specifically for Prop-Fan cruise near fieid noise control but is generally applicable for any turboprop noise analysis. The noise method recognizes the components of tone noise associated with thickness, loading, and quadruple. Thickness and loading aro linear components and are directly related to the blade surface pressure and geometric definition.

Quadruple is a non-linear component related to the velocity derivatives in air around the airfoil. Non-linear effects are important when the air- foils are operating near their section critical Mach number. This method which recognizes the details of the blade allows much more effective noise reduction designs.

As mentioned earlier, blade sweep can be a very effective means for performance improvement, that is, increasing efficiency and reducing noise. The advantages of sweep for Prop-Fan are quite significant for high subsonic airsped where the blade tip helical Mach number is supersonic.

The character of the aero&J-namic and acoustic improvements differ in that efficiency peaks at about 40 degrees of tip sweep while noise continues to decrease as sweep increases. The reason is that noise reduction is attributed to eliminating compressibility effects as with aero and also to a cancellation phenomenon. In fact, the cancellation of source noise by sweep can be accomplished for subsonic tip helicals as well, once compressibility has been eliminated. For the latest Prop-Fan blade design efficleacy is up two percentage points and near field noise is down 18 dB compared to a straight blade. For the Prop-Fan, near field source noise at 0 . 8 Mach cruise is only one-third of the story associated with a comfortable interior of about 80 dB ' ) .

In order to achieve this very quiet aircraft interior noise level, source noise reductions are being pursued and the source noise objectives are considered achieveable. Recent interior noise data indicates that further reductions are possible by correctly handling phasing and rotation effects. Synchrophasing can reduce the level in the peak noise area substantially. Finally, the fuselage is designed for attenuation and there is about 20 dB noise reduction for a standard turbo- fan type fuselage. Increased noise attenuation fuselage designs are under study and are considered practical.

So, at 0 . 8 Mach it is possible to achieve a quiet interior. At lower Mach number it becomes easier.

The blade concept for improved efficiency and reduced noise must be structurally sound. Blade construction is key to a propeller/Prop-Fan design. It establishes the structural dynamics, is the major weight contributor, sizes the mechanical components, and establishes the maintenance and reliability philosophy The spar-shell blade construction concept, where the spar is metal structure and the shell is lightweight

fiberglass , offers a large weight L eduction over the traditional solid

aluminum hlade and allows the shaping of the blade for enhanced performance with fewer constraints. This spar-shell concept is safer and offers improved reliability and enhanced maintenance. Reliability is imprwed in several ways such an integral (buried) blade heater and individually Failure probability is reduced by eliminating surface replaceable blades.

Hamilton Standard has damage as a source of structural degradation.

accomplished blade designs for swept blades using traditional beam analysis and the more sophisticated finite element analysis technique. The accuracy and the level. of information resulting from finite element analysis is quite superior to beam analysis. The stresses and deflections are defined everywhere on the spar, in the bond, and on the shell; and the mode shapes and natural frequencies are more precisely defined. The use of this advanced analysis technique will provide a hlade design with high structural confidence and light weight. The Prop-Fan blade construction concept i s an extension of the very successful current production fiberglass blade configuration. The highlights of the Hamilton Standard metal spar-composite shell experience die 5000 blades manufactured of 22 different designs.

Thirteen achieved flight test status and four of these achieved production.

The estimated biade flight time is 1.7 million hours.

I have reviewed the last 20 plus years of Hamilton Standard blade safety experience and the data shows about 35 million flight hours between in- flight blade fractures on reciprocating engines and no in-flight fractures on turbine engines in 60 million hours. This compares with general aviation data indicating about one million flight hours between fractures. A review of our blade fractures indicates that all were due to damaged exterior surfaces operating in the high engine vibratory environment of a recip.

Elimination of the recip environment has eliminated blade fractures. It that elimination of the environmental damage to the blade is projected structure will reduce the probability of failure. Coupling this with improved str~ctural analysis techniques shouldvirtually eliminate blade failures.

A study recently done for NASA determined the reliability and maintenance costs for a typical turboprop propulsion system such as the Electra. Spe- cific problem areas were isolated and improved upon in a preliminary design of a new Prop-Fan propulsion system.

It was found that significant reli- ability improvements and mainte:.ance cost reductions could be achieved for both the propeller and the turboprop related portions of the gearbox. In both cases, adoption of a modern on-condition maintenance philosophy, thereby eliminating scheduled major maintenance and overhaul, reduced costs in It was found on the Electra gearbox that the non-turboprop functions half.

of engine and airframe accessory drives accounted for about 258 of the gearbox unscheduled maintenance costs. Such drives are necessary for any propulsion system and should be accounted with the core engine for turb3props as they are for turbofans. Lastly, as mentioned earlier on advanced blades, improvements in cost can be made as a result of improved reliability and modularity. Simplified component hardware, individually removed with simple procedures finally results in dollars per flight hour which are very low. The propulsion system maintenance cost is dominated by the turbine core engine, not the propulsive device.

Advanced design techniques when applied to advanced turboprops or Prop- Fans operating at high subsonic Mach number have improved performance over both conventional turboprops and high bypass turbofans. They additionally offer uncompromised safety and high reliability/low maintenance characteristics. Application of this propulsion concept extends the traditional turboprop utilization in General Aviation to high Mach where significant gains can be achieved over turbofans.

ADVANCED TURBOPROP POTENTIAL

FOR HIGH SPEED

Prop- Fan

Figure 1

ADVANCED TURBOPROP OBJECTIVES

Improved Performan.:e Above M = 0.55

Low Airport Noise

a Turbofan ,nterior Comfort

Enhanced Structural Integrity

Reduced Operating Costs

Figure 2

VARIATION OF INSTALLED EFFICIENCY

0.5 0.6 0.7 0.8 0.9

Cruise Mach Number

Figure 3

FUEL SAVINGS

Prop-Fan Vs. Turbohn

A

2000 4000 6000

Operating Range (nm)

Rgure 4 HIGH-SPEED-PROPELLER WIND-TUNNEL AEROACOUSTIC RESULTS

Robert J . Jeracki and James H . Dittmar

National Aeronautics and Space Administration Lewis Research Center The energy s a v i n g p o t e n t i a l o f t h o high-speed t u r b o p r o p h a s been d i s c u s s e d 1 t o 9; w i t h many a d d i t i o n a l r e f - w i t h i n c r e a s e d i n t e r e s t i n r e c e n t y e a r s ( r e f s .

e r e n c e s i n r e f . 1). Gatzen's p a p e r a t t h i s c o n f e r e n c e i n d i c a t e d t h e b e n e f i t s from and t h e approach t o a p p l y i n g t h e high-speed p r o p e l l e r t o b u i s n e s s - j e t t y p e of a i r c r a f t . These a i r c r a f t f l y i n t h e Mach 0.5 t o 0.8 range, above t h e speeds of p r e s e n t turboprop-powered e x e c u t i v e a i r c r a f t , and a t a l t i t u d e s above 9.144 km (30 000 f t ) where t u r b o j e t and t u r b o f a n e n g i n e s a r e p r e s e n t l y used. An advanced high-speed turboprop, however, h a s t h e p o t e n t i a l f o r s i g n i f i c a n t f u e l s a v i n g compared w i t h t h e s e two p r o p u l s i o n systems. The h i g h f l i g h t speed and a l t i t u d e and o t h e r d e s i g n c o n s t r a i n t s make t h e high-speed t u r b o p r o p a unique p r o p u l s i o n system.

T h i s p a p e r w i l l p r e s e n t some aerodynamic c o n c e p t s , e x p l a i n how t h e s e a r e a p p l i e d t o advanced p r o p e l l e r d e s i g n , and t h e n s k w r e , .ts from r e c e n t wind- t u n n e l t e s t s a t Lewis. T h i s s h o u l d g i v e a f e e l i n g t o iy t h e c o n c e p t s were used and t h e i r importance i n o b t a i n i n g good aerodynamic and a c o u s t i c p e r f o r - mance. F i g u r e 1 shows how u n i q u e t h i s p r o p u l s i o n system r e a l l y would be, based on t h e d e s i g n c o n c e p t s b e i n g c o n s i d e r e d f o r t h e high-speed turboprop. Most ob- v i o u s a r e t h e b l a d e sweep, long b l a d e c h o r d s , and, o f c o u r s e , t h e l a r g e number of blades. O t h e r d e t a i l s n o t e a s i l y s e e n i n t h i s photograph w i l l be d e s c r i b e d l a t e r . These unique f e a t u r e s come from t h e need t o keep a r e a s o n a b l e p r o p e l l e r s i z e and t o f l y e f f i c i e n t l y and q u i e t l y a t h i g h speed and a l t i t u d e . The l a r g e f u e l s a v i n g p o t e n t i a l and t h e l a c k of a n a d e q u a t e d a t a b a s e f o r t h i s new propul- s i o n concept prompted NASA t o b e g i n a t e s t program t o v e r i f y t h e high-speed t u r - boprop p o t e n t i a l f o r s a v i n g 2 n e G y .

AERODYNAMIC CONCEPTS The high f l i g h t Mach number r e q u i r e s t h e d e s i g n e r t o minimize c o d p r e s s i - Some aerodynamic c o n c e p t s t h a t c o u l d be used a r e shown i n f i g u r e b i l i t y l o s s e s .

I n t h e b l a d e t i p r e g i o n c o m p r e s s i b i l i t y l o s s i s reduced by u s i n g t h i n a i r 2.

f o i l s e c t i o n s and by sweeping t h e b l a d e t i p back, a s i l l u s t r a t e d by t h e two s k e t c h e s a t t h e t o p of t h e f i g u r e . I n t h e hub r e g i o n t h e blockage o f t h e na- c e l l e behind t h e p r o p e l l e r and a r e a - r u l l i n g , o r s c u l p t i n g - o u t o f t h e s p i n n e r between b l a d e s , a r e used t o reduce l o s s e s . These a r e i l l u s t r a t e d by t h e n e x t two sketches. Advanced a i r f o i l s designed f o r h i g h performance and low n o i s e s i g n a t u r e were n o t p a r t o f t h e s e model d e s i g n s , b u t t h e s e c o u l d be i n c l u d e d *'hk.CEUING PAGE E.! A ! # & PJO! F i t MEU l a t e r a s f u t u r e improvements.

The e f f e c t s of t h e s e c o n c e p t s i n c o r p o r a t e d i n t o a p r o p e l l e r d e s i g n are shown i n f i g u r e 3. T h i s f i g u r e i s based on t h e c r u i s e c o n d i t i o n a t Hach 0.8 and p r e s e n t s t h e Mach number approaching the b l a d e from t h e hub t o t h e t i p . The t o t a l Mach number, which i n c l u d e s b o t h t h e f r e e - s t r e a m component and t h e pro- p e l l e r r o t a t i c n a l component, i s t h e t o p curve. The Mach number s t a r t s just above t h e c r u i s e Mach number a t t h e hub, i n c r e a s e s a s t h e r o t a t i o n a l v e l o c i t y becomes l a r g e r , and r e a c h e s Mach 1.14 a t t h e t i p f o r t h e d e s i g n c o n d i t i o n s .

T h i s l o c a l approach Mach number must be compared w i t h t h e Hach number where each b l a d e a i r f o i l s e c t i o n e n t e r s i n t o d r a g r i s e t o e v a l u a t e how t h e p r o p e i l e r w i l l perform. T h e r e f o r e , f o r a t h i n b l a d e w i t h a thicknzss-to-blade-chord r a t i o of about 15 p e r c e n t a t t h e hub and down t o 2 p e r c e n t a t t h e t i p , i s o l a t e d two- dimensional a i r f o i l d a t a were used t o p r e d i c t t h e Mach nuzaber a t which e a c h a i r - f o i l s e c t i o n would go i n t o d r a g rise. That d r a g rise Mach number is t h e second c u r v e i n f i g u r e 3. Note t h a t t h e l o c a l Mach number is above t h e d r a g rise Mach number from t h e hub t o t h e t i p and t h a t a c r o s s t h e e n t i r e b l a d e i s a l a r g e po- t e n t i a l c o m p r e s s i b i l i t y l o s s region. T h i s l o s s r e g i o n is d e p i c t e d i n f i g u r e 3 by t h e cross-hatched region.

The aerodynamic c o n c e p t s shown i n f i g u r e 2 were used t o reduce t h e s e l o s s e s . I n t h e t i p r e g i o n sweep reduces t h e component of v e l o c i t y normal t o t h e b l a d e a i r f o i l s e c t i o n , a s i s done f o r swept wings. So t h e Hach number is re- duced from t h e l o c a l t o t h e e f f e c t i v e Mach number shown i n f i g u r e 3. With t h e e f f e c t i v e Mach number below t h e d r a g d i v e r g e n c e Mach number i n t h e t i p r e g i o n , I n t h e hub r e g i o n n a c e l l e b- xckage behind t h e l o s s is s i g n i f i c a n t l y reduced.

t h e p r o p e l l e r reduces t h e l o c a l Mach number through t h e p r o p e l l e r plane. That i s p l o t t e d a s t h e e f f e c t i v e Mach number c u r v e n e a r t h e hub. A d d i t i o n a l s u p p r e s s i o n i s used h e r e because, w i t h t h e l a r g e number o f b l a d e s , t h e hub b l a d e s e c t i o n s o p e r a t e a s e s s e n t i a l l y a c a s c a d e o r f a n where blade-to-blade choking c o u l d b e a problem. Area-ruling t h e s p i n n e r between b l a d e s g i v e s f u r t h e r p r o t e c t i o n from choking by opening t h e f l o w a r e a between t h e b l a d e s a t t h e F u r t h e r d i s c u s s i o n of t h e a p p l i c a t i o n o f t h e s e c o n c e p t s t o high-speed spinner.

p r o p e l l e r d e s i g n is g i v e n i n r e f e r e n c e s P, 4, and 5, and b l a d e s t r u c t u r a l d e s i g n i s covered i n r e f e r e n c e s 3 and 10.

PROPELLER MODEL DESIGNS The c o n c e p t s d e s c r i b e d above were used t o d e s i g n a s e r i e s o f p r o p e l l e r models f o r wind-tunnel t e s t i n g i n a c o o p e r a t i v e program betwee;: L e w i s and Hamilton Standard. The t h r e e b a s i c b l a d e planforms p i c t u r e d i n f i g u r e 4 r e p r e - s e n t t h e f o u r p r o p e l l e r d e s i g n s . -n common a r e t h e b l a d e t i speed of 244 mlsec

S

(800 f t l s e c ) , c r u i s e power l o a d i n g o f 301 kw/m2 (37.5 s h p l f t ) (which is about f o u r t i m e s t h a t of a c o n v e n t i o n a l p r o p e l l e r such a s on t h e E l e c t r a ) , and e i g h t blades. The planforms a r e i d e n t i f i e d by t h e i r sweeps o f 0, 300, and 4s0, Here, t h e t i p sweep i s approximately t h e a n g l e o f t h e t i p o f t h e b l a d e measured back from a r a d i a l l i n e normal t o t h e a x i s of r o t a t i o n through t h e b l a d e r o o t .

The o r i g i n a l 0 and 300 swept b l a d e s were d e s i g n e d u s i n g e x i s t i n g e s t a b - l i s h e d a n a l y s e s ( r e f . 11) t h a t lacked a r e f i n e d methodology t o d e s i g n t h e t w i s t of a swept blade. i n d i c a t e d a I n i t i a l t e s t s of t h e 30° swept d e s i g n (SR-1) retwist was r e q u i l e d ( t h a t is, r e d i s t r i b u t i o n of t h e b l a d e load from hub t o t i p ) . The r e t w i s t e d b l . ~ d e became t h e second 300 swept d e s i g n (SR-1M). The 450 swept b l a d e was swept, and t h e planform shaped f o r a c o u s t i c s u p p r e s s i o n d s well a s improved aerodynaeic performance. More d e t a i l e d d i s c u s s i o n s of t h e a e r o a c o u s t i c d e s i g n methodology a r e p r e s e n t e d i n r e f e r e n c e s 3, 12, and 13.

E f f i c i e n c y and n o i s e l e v e l were p r e d i c t e d when t h e s e b l a d e s were designed.

Those p r e d i c t e d e f f i c i e n c i e s ( l i s t e d i n f i g . 4) i n d i c a t e d improved performance w i t h i n c r e a s e d sweep. Those n o i s e p r e d i c t i o n s i n d i c a t e d some r e d u c t i o n f o r 30° of sweep and s i g n i f i c a n t r e d u c t i o n f o r t h e a e c o a c o u s t i c 4S0 swept design.

The photographs i n f i g u r e 5 show t h e 0, 300, and 45O swept, 62.2-cm (24.5-in.) d i a m e t e r p r o p e l l e r models i n s t a l l e d on t h e P r o p e l l e r T e s t Rig (PTR) i n t h e Lewis 8- by 6-foot wind t u n n e l . The t u n n e l ( r e f . 13) h a s a porous w a l l t e s t s e c t i o n t o minimize any w a l l i n t e r a c t i o n s . The PTR i s powered by a 746-kW (1000-hp) a i r t u r b i n e u s i n g a c o n t i n u o u s flow, 3. 1x106-~/m2 (450-psi) a i r system routed through t h e s u p p o r t s t r u t . Force and t o r q u e on t h e p r o p e l l e r a r e measured on a r o t a t i n g balance l o c a t e d i n s i d e o f k-. a x i s y w e t r i c n a c e l l e behind t h e p r o p e l l e r .

PROPELLER IERODY NAMIC PERFORMANCE T y p i c a l t e s t r e s u l t s from t h e 45O swept d e s i g n a r e shown i n f i g u r e 6 t o g i v e a n u n d e r s t a n d i n g of t h e way d a t a were t a k e n and used. T h i s i s t h e b a s i c p r o p e l l e r d a t a p l o t where n e t t h r u s t e f f i c i e n c y and i . d i m e n s i o n l e s s power coef- f i c i e n t a r e p l o t t e d as o r d i n a t e s . The a b s c i s s a i s t h t advance r a t i o , which is p r o p o r t i o n a l t o t h e r a t i o o f f l i g h t o r advance speed t o b l a d e t i p speed.

A s t i p speed i n c r e a s e s from windmill ( n o power), t h e advance r a t i o d e c r e a s e s a s shown by t h e two h o r i z o n t a l s c a l e s . Blade a n g l e i s s e t and d a t a a r e t a k e n from wind- m i l l t o h i g h e r power a s shown by t h e d a t a symbols on t h e power c o e f f i c i e n t p l o t .

The b l a d e a n g l e ( $ 3 / 4 ) , measured a t 3 / 4 o f t h e p r o p e l l e r r a d i u s , be- comes 90° when t h e chord o f t h a t a i r f o i l s e c t i o n i s a l i g n e d d i r e c t l y w i t h t h e f l i g h t d i r e c t i o n . A s power i s i n c r e a s e d t h e t h r u s t i n c r e a s e s a n d , a s s e e n i n t h e upper d a t a c u r v e s , t h e n e t t h r u s t e f f i c i e n c y i n c r e a s e s , r e a c h e s a peak, and t h e n begins t o d r o p o f f . O t h e r b l a d e a n g l e s y i e l d s i m i l a r power and e f f i c i e , ~ c y curves.

A t t h e d e s i g n Mach number of 0.8, t h e d e s i g n power l o a d i n g and d e s i g n t i p speed g i v e a power c o e f f i c i e n t of 1. 7 and advance r a t i o of 3.0b. A s o l i d l i n e i s drawn through t h i s p o i n t on t h e power c o e f f i c i e n t p l o t , i n t e r s e c t i n g t h e t w .

1 i n e s of d a t a shown. T h i s s o l i ~ ' l i n e r e p r e s e n t s t h e d e s i g n power a t d i - i e r e n t p r o p e l l e r t i p speeds.

The e f f i c i e n c y a t t h e d e s i g n power c a n be found f o r each b l a d e a n g l e , i n d i c a t e d by each v e r t i c a l l i n e . Then t h e v a r i a t i o n of n e t e f f i - c i e n c y w i t h advanced r a t i o ( i . e., t i p speed) a t d e s i g n power c a n be p l o t t e d a s shown i n f i g u r e 7. T h i s p l o t i s f o r models w i t h a r e a - r u l e d s p i n n e r s a t t h e d e s i g n power l o a d i n g a t Mach 0.8. Curves of n e t e f f i c i e n c y v e r s u s advance r a t i o #.re compared f o r d i f f e r e n t sweep a n g l e s . S i g n i f i c a n t improvement c a n be s e e n i.n going from 0 t o 30° of sweep.

The 4 5 O swept b l . a shows s t ill more improve- ment, e s p e c i a l l y a t low advance r a t i o s ( c o r r e : ~ . :ding t o h i g h t i p speeds). The o v e r a l l improvement a t t h e d e s i g n advance r a t . 3 3.06 i s a b o u t 3 p e r c e n t .

O t h e r important d e s i g n v a r i a t i o n s were i n v e s t i g a t e d ..sing t h e 300 swept d e s i g n s . A s noted i n t h e d e s c r i p t i o n of t h e b l a d e d e s ~ ., t h e r e were two d i f - f e r e n t t w i s t , o r l o a d i n g d i s t r i b u t i o n s , w i t h t h e same hV swept planform. The b l a d e d e s i g n w i t h t h e r e v i s e d ( r ~ a u c e d ) t w i s t was t e s t e d w i t h b o t h a c o n i c and a n a r e a - r u l e d s p i n n e r . The performance cou,parison i s shovn i n f i g l l r e 8. A s i n t h e p r e v i o u s f i g u r e , t h i s one p r e s e n t s d a t a f o r t h e d e s i g n Mach n ~ m b z r and power loading.

The o r i g i n a l d e s i g n ( b a s e l i n e t w i s t ) yes t e s t e d w i t h a c o n i c s p i n n e r and is t h e lowest of t h e t h r e e d a t a curves. R . z t w i s t i ~ g t o i n c r e a s e t h e l o a d a t t h e t i p improved t h e performance n e a r t h e d e s i g n advance r a t i o , where t h e high- e s t e f f i c i e n c y f o r t h a t blade-spinner combination t h e n o c c u r r e d . That reduced- t w i s t d e s i g n was a l s o t e s t e d w i t h a n a r e a - r u l e d s p i n n e r . That change improved t h e performance a b o u ~ 1 p e r c e n t o v e r t h e f u l l rarGe o f t i p s p e e d s t e s t e d . T h i s f i g u r e i n d i c a t e s t h e b e n e f i t o f a r e a - r u l i n g and t h a t th;. p r o p e r t w i s t o r l o a d i n g is r e q u i r e d t o o b c a i n h i g h performance.

Because figv.res 7 and 8 summzrize d a t a a t t h e d e s i g n power, t h e perf o r mance a t t h e a c t u a l d e s i g n p o i n t a t Mach 0.8 c a n b e o b t a i n e d a s t h e n e t e f f i - cierlcy a t t h e advance r a t i o (3.06) corresponding t o t h e d e s i g n t i p speed. Siml- l a r l y , n e t e f f i c i e n c y a t o t h e r f ree-stream Mach numbers c a n be ~ b t a i n e d a t t h e same power c o s f f i c i e n t and advance r a t i o a s t h e Mach . 8 d e s i g n p o i n t . The v a r i a t i o n of performance w i t h f r e e - s t ree-.L Mach number i s jhown i n f i g u r e 9 a t con- s t a n t power c o e f f i c i e n t and advanrc r a t i o , f o r t h e 0, 330, and 450 swept blade d e s i g c s w i t h a r e a - r u l e d s p i n n e r s T h i s p l o t i s i n t e r e s t i n g because t h e i d e a l e f f i c i e n c y , p r e s e n t e d a s t h e u p p e r dashed l i n e , i s c o n s t a n t a c r o s s t h e Mach num- b e r range f o r a g i v e n v a l u e of power c o e f f i c i e n t and advance r a t i o . T5e i d e a l e f f i c f e n c y is t h e performance of a n optimlim p r o p e l l s r w i t h n o b l a d e d r a g and, s o , r e p r e s e n t s o n l y a x i a l momentum, s w i r l , and t i p l o s s e s . Selow t h a t l i n e i s t h e r e a l world where v i s c o u s and c o m p r e s s i b i l i t y l o s s e s occur. A s t h e d a t a show, t h o s e l o s s e s i n c r e a s e a s Mach number i s i n c r e a s e d .

Again, t h e b e n e f i t from sweep o f about 3 p e r c e n t - s e e n a t Mach 0.8. The e f f i c i e n c y of t h e 45O swept model approached t h e \;ah: ,sed i n t h e s t u d i e s which showed t h e l a r g e f u e l s a v i n g p o t e n t i a l f o r t h e h..dh-speed turboprop.

A c t u a l l y , a t power l o a d i n g s lower t h a n d e s i g n , t h e e f f i c i m c y s l i g h t l y exceeded t h e s t u d y value. However, t h e lower power l o a d i n k would r e q u i r e a l a r g e r and h e a v i e r p r o p e l l e r f o r t h e same a i r c r a f t i n s t a l l a t i o n .

The 450 swept b l a d e which achieved t h i s h i g h performance a l s o r e t a i n s f a i r l y h i g h e f f i c i e n c y o u t t o Mach 0.85. More perfonliance d e t a i l s a r e g i v e n i n r e f e r e n c e s 1, 4, 5, and 6.

PROPELLER ACOUSTIC PERFORMANCE Acoustic d a t a were a l s o t a k e n i n t h e t u n n e l , s i n c e t h e c a b i n n o i s e a t c r u i s e c o n d i t i o n s i s of concern. The t r a n s o n i c p r o p e l l e r r e l a t i v e t i p speed of t h e s e b l a d e d e s i g n s h a s t h e p o t e n t i a l f o r g e n e r a t i n g high n o i s e l e v e l s . T h i s n o i s e rgeeds t o be minimized and f u s e l a g e attenuation needs some imprcvement.

Wall-mounted p r e s s u r e t r a n s d u c e r s were used t o o b t a i n n e a r f i e l d a c o u s t i c d a t a f o r t h e p r o p e l l e r models. ( F u r t h e r d e t a i l s a r e g i v e n i n r e f s . 1 5 and 16. i Wave shapes o f t h e n e a r f i e l d p r e s s u r e s i g n a l d u r i n g b l a a e p a s s a g e s a r e shour. i n f i g u r e 10. These a r e enhanced pressure-time t r a c e s f o r b o t h t h e s t r a i g t t a l a d e and t h e a e r o a c o u s t i c a 1 l ; r designed 650 swept blade. The t r a n s d u c e r c l o s e s t t o t h e p r o p e l l e r was used ( o n t h e t u n n e l w a l l i n t h e p l a n e of t h e p r o p e l l e r ) . A s caz be seen, t h e t r a c e f o r t h e s t r a i g h t b l a d e shows a h i g h amplitude, s t e e p wave However, i n t n e p i o t shape, which approaches t h e c l a s s i c N dave shock p a t t e r n .

f o r t h e q u i e t e r 6 p swept b l a d e , a n almost s i n u s o i d a l wave was observed which i s a l s o of c o n s i d e r a b l y l e s s amplitude. These d i f f e r e n c e s i n t h e c n a r a c t e r of t h e t r a c e s i n d i c a t e d t h a t t h e a e r o a c o u s t i c a l l y design& p l a n f o m o f t h e 45O swept b l a d e h a s been s u c c e s s f u l i n r e d w i n g t h e s h a r p p r e s s u r e r i s e t h a t would normally be a s s o c i a t e d w i t h t r a n s o n i c h e l i c a l - t i p - s p e e d p r o p e l l e r s .

Another comparison of t h e b e n e f i t s of s e e p makes t h e magnitude of t h e n o i s e r e d u c t i o n more a p p a r e n t . F i g u r e 11 is a p l o t of t h e maximum b l a d e passage t o n e on t h o t u n n e l c e i l i n g v e r s u s t h e h e l i c a l - t i p ( t o t a l , i n c l u d i n g f l i g h t and r o t a t i o n a l ) Hach number. The advance r a t i o and power c o e f f i c i e n t f o r a l l o f t h e V a r i a t i o n i n h e l i c a i - t i p 3 a c h d a t a p o i n t s are approximately t h e d e s i g n values.

number is o b t a i n e d by t a k i n g d a t a a t d i f f e r e x : f r e e - s t r e a m nacn numbers. Tne p l o t s f o r b o t h 0 and 45O s v + p t b l a d e s e x n i b i t a r e g i o c o f s h a r p n o i s e i n c r e a s e w i t h i n c r e a s i n g h e l i c a l - t i p Mach number, which is t h e n followed by a r e g i o n where n o i s e l e v e l s o f f . The c a j l o r e d sweep o f t h e 4 5 O desigr. p r o v i d e s n o i s e r e d u c t i o n o v e r t h e complete range of t i p speeds.

Near t h e c r u i s e d e s i g n t i p Hach number o f I. 14, t h e r e d u c t i o n i s about 5 t o 6 dB and a p p e a r s t o be even l a r g e r a t t h e lower t i p s p e e d s t e s t e d . Daca i n r e f e r e n c e 1b from a 54' swept b l a d e , t o g e t h e r w i t h t h e d a t a shown h e r e , i n d i c a t e t h a t i n c r z a s i n g t i p sweep d e l a j ; t h e s h a q i n c r e a s e i n n o i s e t o a h i g h e r t i p Mach rrumber. T h i s d e l a y i n n o i s e rise vould be s x p e c t e d j u s t a s swee? d e l a y s t h e s h a r p d r a g r i s e of t h e b l a d e s by p o s t p o n i ~ g t h e o n s e t of shocks i n t h e b l a d e t i v region. The o v e r a l l n o i s e r e d u c t i o n a t high and low t i p speeds i n d i c a t e s t h e beaef i t of t h e a e r o - acoust:c methodology of t h e 450 swept design.

Noise d a t a *-ere t a k e n a t p o s i t i o n s i n f r o n t of and behind t h e p r o p e l l e r p l a n e on t h e t u n n e l c e i l i n g . F i g u r e 12 i s a p l o t of t h e a x i a l ariat ti an o f t h e b l a d e passage t o n e v e r s u s t h e a x i a l p o s i t i o n from t n e p r o p e l l e r p l a n e p i o t t e d i n p r ~ p e l l e r d i a n e t e r s . The c e i l i n g i t s e l f i s abo:-t 1.5 p l a y e l l e r d i a m e t e r s from t h e p r o p e l l e r i i p . The jame s i g n i f i c a n t r e d u c t i o n i n peak n o i s e l e v e l f o r t h e 45O swept b l a d e r a n be ooserved h e r e a s i~ f i g u r e il. A t Mach 0.8 c r u i s e t h e n o i s e from t h e s e p r o p e l l e r s d i f f e r s o n l y s l i g h t l y ah+-d of t h e p r o p e l l e r and t e n d s t o have t h e i a r g e s t d i f f e r e n c e f a r t h e r h e h i r i t h e p r o p e l l e r . Notice t h e d i r e c t i v i t j of t h e n o i s e p a t t e r n . The peak n o i s e l e v e l h a s dropped o f f s i g n i f - i c a n t l y f r ~ m t h e peak w i t h i n a t o t a l a i s t c i c e o f aboui 2 p r o p e l l e r d i a m e t e r s .

T h i s i n d i c a t e s t h a t any r e q u i r e d f u t e l a g e t r e a t m e r t would be l i m i t e d i n a r e a .

Hare d c t a i l s and d a t a a r e g i v e n i n r e f e r e n c e s 15 ;.nd I t .

SUETlARY OF RESJLTS The n c i s e r e d u c t i o n and t i n e h i g h measured pe?*ormance show t h e lrodyanmic and acous: ic benef i r -. * c a d v a n c e i p r o p e l l e r d e s i g n concept: Hifh aerodynamic ~ e r f o ~ r n a n c e was o b c a 1 . 1 d a t Mach 0.8; w i t h i n 1 p e r c e n t of t h e s t g d y v a l u e used P e r f o m - t o p r e d i c t l a r g e p o t e n t i a l f u e l s a v i n g f o r t h e high-speed turboprop.

ance n e a r 80 p e r c e n t was o b t a i n e d a t l o v e r than-design power l o r d i q s . Nacelle blockage w a s a o important p a r t of t h e d e s i g n s , and a r e a - r u l i n g was shown t o k important i n improving measured performance. i3lade t i p sweep improved a e m d y - n a n i c performance about 3 p e r c e n t , w h i l e t h e a e r o a c o u s t i c d e s i g n o f t h e 450 swept p r o p e l l e r reduced c m i s e n e a r - f i e l d n o i s e about 6 dB.

The performance r e s u l t s shown a r e a t t r a c t i v e , and f u r t h e r r e f i n e m e n t s c a n be made. There is -GW a b a s i s f o r improvements i n d e s i g n and a n a l y s i s . F i g u r e 1 3 i n d i c a t e s t h e f JL -re of high-s?eed turboprop improvewnts. The l o w e r - l e f t c i r c l e r e p r e s e n t s t b r e c u r r e n t d e s i g n procedures and model r e s u l t s d e s c r i b e d i n t h i s paper. The p r o p e l l e r s v e r e e i g h t bladed, h i g h l y loaded, and designed using e s t a b l i s h e d a n a l y s e s , af though a n m b e r o f advance.. c o n c e p t s were incorporated.

Two new f u r t h e r advanced models a r e being designed and should 3e t e s t e d I n 1980. These models a r e advanced, 10-bladed d e s i g n s which have g r e a t e r t i p sweep and l o v e r t i p speed t o improve t h e a c o ~ a s t i c as w e l l as t h e aerodynamic p e r f o r Sorne r e f i n e d a n a l y s e s were a v a i l a b l e f o r t h e d e s i g n o f t h e s e b l a d e s l a n c e .

( r e f . 12).

The r e s u l t s of b ~ t h t h e p r e s e n t t e s t s and chose planned i n 19NJ w i l l be used tr, a c h i e v e a n i n i c i a l o p t i m m d e s i g n , c a i l e d SR-7. When t h a t p r o p e l l e r i s being designed, a d d i t i o n a l advanced anakyses ( d e s c r i b e d i n Bober's paper a t t h i s conference and r e f . i l l w i l l be a v a i l a b l e t o f u r t h e r enhznce t h e d e s i g n prac- T e s t i n g t h a t d e s i g n w i l l conclude t h e p r e s e n t vind t u n n e l program on pro- ess.

p e l l e r performance and noise. Another approach which is under s t u d y by NASA as a f u t u r e r e s e a r c h a r e a t o f u r t h e r improve perfr~rmance i s t o r e c o v e r t h e t h r u s t l o s t i n t h e s w i r l of t h e p r o p e l l e r s l i p s t r e a m . The s w i r l l o s s f o r t h e s e h i g h l y loaded p r o p e l l e r s can be a s mcch a s 6 t o 8 p e r c e n t i n e f f i c i e n c y . Methods being c o n s i d e r e d f o r s w i r l recovery a r e c o a x i a l c o u n t e r r o t a t i o n , wing c o n t c u r i n g be- hind t h e p r o p e l l e r ( t o a c t l i k e a s t a t o r ) , and t h e i n t r o d u c t i o n of s t a t o r s be- hind t h e p r o p e l l e r . T h i s continued e f f o r t , shown i n f i g u r e 13, is expected t o a i l o w f u t u r e p r o p e l l e r s t o be designed f o r high-speed f l i g h t w i t h both h i g h e r lower c r u i s e noise.

e f f i c i e n c y and s i g n i f i c a n t l y REFERENCES 1. J e r a c k i , Robert J. ; Hikkelson, Daniel C. ; and Blaha, Bernard 2 . : Wind ?urine1 Performance o f Four Energy E i f i c i e n t P r o p e l l e r s Designed f o r P x n 0.8 Cruise. NASA TM-79124, 1979 o r SAE Paper 790573, Apr. 1979.

2. Nored. Donald L. : Fuel Conservative A i r c r a f t Engine Technology. NASA TM- 78962, 1978.

3. h g a n , James F. Jr. ; Gat t e n , Bernard S. ; and Adamson, William M. : Prop-Fan Propulsion - Its S t a t u s and P o t e n t i a l . SAE Paper 780995. Nov. 1978.

4 . Xikkelson, D a n i e l C. ; e t a l . : Design a n d Performance o f Energy E i f i c i e n t P r o p e l l e r s f o r Mach 0.8 Cruise. USA T M X-73612, 1977 o r SAE P a p e r 770658, Mar. 1977.

5 . Black, D. M . ; Menthe, R. W . ; and Wainauski, H. S.: A e r o d y n r i c D e s i g n and Performance T e s t i n g o f a n Advancea 300 Swept, E i g h t Bladed P r o p e l l e r a t n a c h Numbers from 0.2 t o 0.85. NASA CR-3047, 1978.

6. Rohrbach, C a r l : Prop Fan O f f e r s Lou F u e l Consumptioa a t Today's F l i g n t Speeds and A l t i t u d e s . A I M P a p e r 76-667, J u l y 1976.

I . J a c k s o n , A. H . , Jr. ; and Gatzen, B. S. : ? ¶ u l t i - t ¶ i s s i o n Uses for Prop-Fan V a r i a b l e Geometry and nu1 t i c y c l e E n g i n e s , A G A W C P - ~ ~ ~ , F ,-opulsion.

A d v i s o ~ Group F o r Aerospace R e s e a r c h a n d Development, P a r i s , 1977, pp.

15-1 t o is-13.

8. K r a f t , C. A. ; and S t r a c k , W. C. : P r e l i m i n a r y S t u d y o f Advanced T u r b o p r o p s f o r Lou Energy Canswuption. NASA TH X-71740, 1975.

9. Rohrback, C . ; and S e e r y , H . E. : A New Look st t h e Turboprop. A I M 3aper 75-1208, S e p t . 1975.

LO. C o r n e l l , R. W. ; and R o t b a n , E. A. : C e s i g n a n d A n a l y s i s o f Prop-Fan Blades. AIAA P a p e r '9-1116, Gune 1979.

11. Bober, L. A.; and M i t c k r l l , G. A. : Suarmary o f Advanced Methods f o r Pre- d i c t i r i g High Speed P r o p e l l e r Performance. NASA TM-81*09, 1980, o r A I M Paper 80-0225, Jan. iYBi.

:2. Metzger, F . B. - Rohrbach, C.: A e r o a c o u s t i c Design o f t h e Prop-Fan.

A I A A P a p e r 79-0610, Mar. 1979.

13. Hanson, n . B. : Near F i e l d Noise o f S u p e r s o n i c P r o p e l l e r s i n Fcnaard F l i g h t . AXE-% P a p e r 76-565, J u l y 1976.

16. Swallow, Robert J.; and A i z l l o , Robert A.: NASA Lewis 8- by 6-Foot S u p e r s o n i c Wind Tunnel. N A S A T M X-71512, 1974.

15. D i t t m a r , James H. ; Blaha, Bernard J. ; and J e r a c k i , Robert J. : Tone Noise o f T h r e e S u p e r s o n i c H e l i c a l T i p Speed P r o p e l l e r s i n a Wiqd Tunnel rt 0.8 n a c h Number. NASA TM-79046, 1978.

16. D i t t m a r , James H. ; J e r a c k i , Robert J. ; and B l a h a , Bernard J. : Tcrle N o i s e o f NASA Three S u p e r s o n i c H e l i c a l T i p Specd P r o p e l l e r s i n a Wind Tunnel.

TM-79167, 1979.

ADVANCED PROPELLER AERODYNAMIC ANALYSES Lawrence 1. Bober National Aeronautics and Space Administration Lewis Reasarch Center SUMMARY Three advanced a n a l y s i s methods f o r p r e d i c t i n g t h e aerodynamic performance of p r o p e l l e r s a r e presented. Two of t h e s e a n a l y s e s a r e l i f t i n g - l i n e methods, and t h e t h i r d i s a l i f t i n g - s u r f a c e method. The approach used i n each of t h e methods is d e s c r i b e d , and t h e c a p a b i l i t i e s a r e presented.

INTRODUCTION I n c r e a s e d concern o v e r f u e l c o s t and a v a i l a b i l i t y have f o s t e r e d renewed i n t e r e s t i n p r o p e l l e r s f o r a i r c r a f t p r o p u l s i o n because of t h e p r o p e l l e r ' s i n h e r e n t h i g h e f f i c i e n c y compared w i t h a t u r b o f a n . The q u e s t f o r improved e f f i c i e n c y and lower n o i s e a t h i g h s u b s o n i c f l i g h t s p e e d s a t 30 000 f e e t o r more h a s f o r c e d p r o p e l l e r s away from c o n v e n t i o n 1 d e s i g n s toward t h e t y p e shown i n t h e p r e v i o u s paper. The combination of h i g < f l i g h t speed and h i g h r o t a t i o n a l speed h a s r e - s u l t e d i n p r o p e l l e r s w i t h a l a r g e number of h i g h l y swept b l a d e s having s i g n i f i - c a n t c a s c a d e e f f e c t s i n t h e inboard r e g i o n of t h e blades. C a r e f u l l y c o n t o u r e d n a c e l l e s c o n t r i b u t e t o t h e a l r e a d y complex flow f i e l d a s s o c i a t e d w i t h t h e pro- p e l l e r . These r a d i c a l l y d i f f e r e n t g e o m e t r i e s and t h e complex f l o w f i e l d s t h e y c a u s e c a n n o t be a d e q u a t e l y analyzed u s i n g c o n v e n t i o n a l aerodynamic performance analyses. To overcome t h e shortcomings of e s t a b l i s h e d approaches t o p r o p e l l e r performance p r e d i c t i o n s , advanced p r o p e l l e r aerodynamic a n a l y s e s a r e being de- veloped a s p a r t of NASA's Advanced Turboprop Program.

T h i s p a p e r w i l l d i s c u s s t h r e e advanced a n a l y s e s c u r r e n t l y under d+ve;op- mrnt. Two a r e l i f t i n g - l i n e a n a l y s e s i n which e a c h b l a d e i s r e p r e s e n t e d by a s i n g l e l i n e of v o r t i c i t y . The t h i r d i s a l i f t i n g - s u r f a c e a n a l y s i s i n which e a c h b l a d e i s r e p r e s e n t e d a s a s o l i d s u r f ace.

Although t h e s e a n a l y s e s were developed f o r t h e kinds of p r o p e l l e r s d e s c r i b e d i n t h e p r e v i o u s paper, t h e y have f e a t u r e s t h a t a r e a p p l i c a b l e t o b o t h low and high-speed g e n e r a l - a v i a t i o : ~ p r o p e l l e r s .

Be- f o r e d i s c u s s i n g t h e advanced a n a l y s e s , an e s t a b l i s h e d approac!~ t o p r o p e l l e r per- formance p r e d i c t ion w i l l be d i s c u s s e d .

ESTABLlSHED APPROACH I n a v e l o c i t y p o t e n t i a l s o l u t i o n f o r t h e flow a r o ~ i n d a p r o p e l l e r , t h e non- uniform spanwise l o a d i n g on t h e b l a d e s c a u s e s a s h e e t of v o r t i c i t y t o extend downstream t o i n f i n i t y a s shbwn i n f i g u r e 1. T h i s v o r t e x wake is shown a s a f i - r e p r e s e n t e d a s a n i n f i -

n i t e number of f i l a ... e n t s , b u t c o u l d more a c c u r a t e l y be

n i t e number o f v o r t e x f i l a m e n t s . The v o r t e x wake i s important i n p r o p e l l e r performance p r e d i c t i o n s i n c e it c a u s e s a n induced v e l o c i t y a t t h e p r o p e l l e r , thereby changing t h e l o c a l b l a d e a n g l e of a t t a c k .

The important f e a t u r e s of t h i s e s t a b l i s h e d approach a r e summarized i n f i g - u r e 1. t h e induced f l o w i n t h i s approach is based on t h e work C a l c u l a t i o n of done by G o l d s t e i n ( r e f . 1 ) about 50 y e a r s ago. Because of t h e l i m i t e d computing c a p a b i l i t y a t t h a t time, G o l d s t e i n used a very s i m p l i f i e d model s o t h a t he c o u l d o b t a i n a n a n a l y t i c a l s o l u t i o n f o r t h e induced v e l o c i t y a t t h e p r o p e l l e r d u e t o t h e wake. He assumed t h a t t h e shape of t h e wake was a r i g i d h e l i x which was known t o correspond t o a l i g h t l y loaded p r o p e l l e r w i t h optimum d i s t r i b u t i o n of loading. To v i s u a l i z e a r i g i d h e l i c a l wake, c o n s i d e r a p l a n e normal t o t h e a x i s of r e v o l u t i o n of t h e p r o p e l l e r . The i n t e r s e c t i o n o f t h i s p l a n e w i t h t h e r i g i d h e l i c a l wake i s a s t r a i g h t l i n e . Another f e a t u r e of t h i s wake shape is t h a t t h e p i t c h of t h e h e l i x d o e s n o t change w i t h a x i a l l o c a t i o n . ( T h i s wake i s shown i n f i g . 1.) G o l d s t e i n p u b l i s h e d r e s u l t s f o r s i n g l e r o t a t i o n p r o p e l l e r s ( r e f . 11, and Theodorsen o b t a i n e d r e s u l t s f o r c o a x i a l c o u n t e r r o t a t i n g p r o p e l l e r s u s i n g an e l e c t r i c a l analogy ( r e f . 2 ) . The induced v e l o c i t i e s o b t a i n e d from t h e G o l d s t e i n and Theodorsen r e s u l t s a r e s t r i c t l y c o r r e c t o n l y f o r s t r a i g h t propeLler blades.

Also t h e i r o r i g i n a l words c o n t a i n e d no p r o v i s i o n s f o r a n a c e l l e s i n c e t h e vor- t i c i t y extended t o t h e a x i s of t h e p r o p e l l e r . These r e s u l t s form t h e b a s i s of a procedure which h a s been r e f i n e d o v e r t h e y e a r s and h a s become a n e s t a b l i s h e d approach t o p r o p e l l e r performance a n a l y s i s .

T h i s procedure i s implemented u s i n g a s t r i p a n a l y s i s i n which t h e f l o w con- d i t i o n s a r e determined a t one r a d i a l l o c a t i o n a t a time.

For e a c h s t r i p t h e induced v e l o c i t y i s determined from G o l d s t e i n ' s o r Theodorsen's r e s u l t s . The e f f e c t of t h e n a c e l l e is t a k e n i n t o account i n an appro imate manner by assuming t h a t a t each s t r i p t h e induced v e l o c i t y i s t h e same a s f o r t h e e n t i r e p r o p e l l e r o p e r a t i n g a t t h e same v e l o c i t y t h a t e x i s t s a t t h a t s t r i p . The t o t a l v e l ~ c i t y is t h e n t h e v e c t o r sum of t h e induced v e l o c i t y , t h e l o c a l v e l o c i t y f o r t h e i s o l a t e d n a c e l l e , and t h e r o t a t i o n a l v e l o c i t y . The flow v e l o c i t y and t h e b l a d e geometry determine t h e l o c a l b l a d e a n g l e of a t t a c k , which a l l o w s t h e d e t e r m i n a t i o n of t h e l i f t and d r a g c o e f f i c i e n t s from i s o l a t e d a i r f o i l d a t a . Sweep i s t a k e n i n t o account through t h e c o s i n e r u l e ( r e f . 3). These f o r c e s a r e r e s o l v e d i n t o t h r u s t and t o r q u e components t h a t c a n be i n t e g r a t e d r a d i a l l y t o g e t t h e p r o p e l l e r t h r u s t , t o r q u e , power, and e f f i c i e n c y .

I t i s important t o n o t e t h a t f o r any o p e r a t i n g c o n d i t i o n , t h e e f f e c t of t h e wake is assumed t o be t h e same a s f o r an o p t i m a l l y loaded p r o p e l l e r . even i f t h e r e i s a n a c e l l e and s p i n n e r p r e s e n t .

ADVANCED ANALYSZS Curved L i f t i n g - L i n e A n a l y s i s The important f e a t u r e s of t h i s a n a l y s i s a r e shown i,i f i g ~ l t - e 2. The wake is r e p r e s e n t e d by a f i n i t e number of h e l i c a l v o r t e x f i l a m e n t s i n s t e a d of t h e con- t i n u o u s s h e e t of v o r t i c i t y used by Goldstein. Each f i l a m e n t h a s c o n a t a n t p i t c h , but i t s l o c a t i o n r e l a t i v e t o a n o t h e r i s a r b i t r a r y a s strown i n f i g u r e 2. A t any p o i n t on t h e b l a d e , t h e induced f l o w d u e t o e a c h w a k e f i l a m e n t and t h e l i f t i n g l i n e i s c c l c u l a t e d u s i n g t h e law of Biot-Savart ( r e f . 3). The t o t a l induced flow a t any p o i n ~ is t h e n t h e sum from a l l t h e s e v o r t i c e s . T h i s a n a l y s i s is c u r r e n t l y r e s t r i c t e d t o s i n g l e - r o t a t i o n p r o p e l l e r s . The p r o p e l l e r b l a d e s a r e r e p r e s e n t e d by curved l i f t i n g l i n e s of a r b i t r a r y shape. The u a c e l l e i s r e - s t r i c t e d t o b ~g a n i n f i n i t e c y l i n d e r s i n c e t h e wakes c a n n o t c o n t r a c t r a d i a l l y .

The s t r e n g t h s of t h e wake v o r t e x f i l a m e n t s a r e r e l a t e d t o t h e spanwise var- i a t i o n of l i f t on t h e blade. Thus i t is n e c e s s a r y t o s o l v e f o r t h e b l a d e and An i m ~ o r t a n t a s p e c t of t h ? s o l u t i o n pro- wake v o r t e x s t r e n g t h s s i m u l t a n e o u s l y .

c e d u r e is t h e placement o f t h e bound v o r t e x a t t h e q u a r t e r chord l i n e and t h e requirement t h a t t h e f l o w be t a n g e n t t o t h e mean camber Line a t t h e t h r e e - q u a r t e r chord l i n e . Thus no i s o l a t e d a i r f o i l d a t a a r e needed s i n c e t h e l i f t a t any r a d i u s c a n be determined from t h e v o r t e x s t r e n g t h a t t h e same r a d i u s . How e v e r , t h i s approach c a n n o t p r e d i c t t h e b l a d e drag. T h i s a n a l y s i s h a s been de- veloped by S u l l i v a n ( r e f . 4) a t Purdue U n i v e r s i t y under a g r a n t from NASA Lewis Research Center.

An i n t e r e s t i n g a p p l i c a t i o n o f t h i s a n a l y s i s i s shown i n f i g u r e 3 , which shows t h e e f f e c t of p r o p l e t s on p r o p e l l e r performance. A p r o p l e t i s a n a e m d y - namic d e v i c e a t t h e t i p o f t h e p r o p e l l e r b l a d e s i m i l a r t o w i n g l e t s , which have been shown t o i n c r e a s e t h e l i f t - t o - d r a g r a t i o of wings. I f t h e p r o p l e t i s prop- e r l y i n t e g r a t e d i n t o t h e t i p f l o w f i e l d , a n improvement i n performance a s shown i n f i g u r e 3 c a n be o b t a i n e d . The p l o t shows t h e p r e d i c t e d e f f i c i e n c y a s ii func- :ion o f power c o e f f i c i e n t f o r a p r o p e l l e r w i t h and w i t h o u t p r o p l e t s . The r e - s u l t s show a n i n c r e a s e i n e f f i c i e n c y due t o t h e p r o p l e t s of about 1 ps-tcn: a t low power and about 3 p e r c e n t a t h i g h power.

The development of t h i s a n a l y s i s i s continuing b o t h a t Purdue and a t Lewis.

R a d i a l l y v a r y i n g i n f l o w v e l o c i t i e s w i l l be i a c l u d e d t o b e t t e r account f o r t h e n a c e l l e . A d r a g p r e d i c t i o n procedure u s i n g i s o l a t e d a i r f o i l d s e a w i l l be implemented s u t h a t t h e e f f e c t of b l a d e d r a g c a n be i w l u d e d .

P r o p e l l e r Nacelle I n t e r , c t i o n A n a l y s i s The second advanced l i f t i n g - l i n e a n a l y s i s h a s narc e x t e n s i v e c a p a b i l i t i e s , which a r e summarized i n f i g u r e 4. "..is a n a l y s i s was developed by United Tech- n o l o g i e s Research Center ( r e f . 5 ) ~ , , : e r c o n t r a c t t o Lewis. The wake i s r e p r e - s e n t e d by a f ' n i t e number of v o r t e x f i l a m e n t s t h a t a r e l o c a t e d on s t r e a w s u r - f a c e s s o t h a t t h e y conform t o t h e shape of t h e n a c e l l e . The p i t c h o f t h e s e f i l a m e n t s i s n u t c o n s t a n t , and they can c o n t r a c t i n b o t h t h e a x i a l and r a d i a l d i r e c t i o n s . T h i s c a p a b i l i t y i s c l e a r l y shown i n f i g u l 4 j u s t downstream of t h e p r o p e l l e r where t h e wake f i l a m e n t s a t e d i s p l a c e d r a d i a l l y because o f t h e i n - c r e a s i n n n a c e l l e diameter. T h i s a n a l y s i s is c a p a b l e of a n a l y z i n g both s i n g l e an. coa; i a l c o u n t e r r o t a t i n g p r o p e l l e r s . The b l a d e s a r e r e p r e s e n t e d by l i f t i n g l i n e s and :an have ar?y a r b i t r a r y shape. rhe q a c e l l e c a n be any axisymmetr'c shape, The s o l u t i o n p r o c e d u r e i n t h i s a n a l y s i s is as follows. F i r s t a n i n v i s c i d s o l u t i o n f o r t h e n a c e l l e a l o n e is o b t a i n e d . The r e s u l t s are u s e d t o locate t h e wake v o r t e x f i l a m e n t s a l o n g stream s u r f a c e s and t o d e t e r m i n e t h e r a d i a l l y vary- i n g i n £ low f o r t h e p r o p e l l e r . The induced f l o w a t t h e p r o p e l l e r d u e t o e a c h wake f i l a m e n t is c a l c u l a t e d u s i n g t h e l a w o f B i o t - S a v a r t ( r e f . 3). The t o t a l induced v e l o c i t y at any r a d i a l l o c a t i o n o n a b l a d e is o b t a i n e d by s u a i n g t h e The induced f l o v from t h e i n d i v i d u a l wake f i l a m e n t s and from t h e l i f t i n g l i n e s .

b l a d e l i f t and d r a g are d e t e r m i n e d from two-dimensional a i r f o i l a n d c a s c a d e d a t a . An i t e r a t i o n p r o c e d u r is r e q u i r e d t o i n s u r e t h a t t h e wake v o r t e x A s t r e n g t h s are c o n s i s t e n t ~ 5 t h t h e spanwise l o a d d i s t r i b u t i o n o n t h e b l a d e .

f i n a l o p t i o n a l s t e p i s tci t -- ,he b l a d e f o r c e s i n a c i r c u m f e r e n t i a l l y a v e r a g e d , v i s c o u s , c o m p r e s s i b l e f l o b e i c u i a t i o n .

F o r a p r o p e l l e r o p e r a ; a t h i g h f l i g h t s p e e d o r h i g h r o t a t i o n a l s p e e d , p o r t i o n s o f t h e b l a d e s may be w v i n g a t s u p e r s o n i c s p e e d s relative t o t h e un- d L s t u r b e d flow. When t h i s o c c u r s some a d d i t i o n a l e f f e c t s (shown i n f i g . 5) u s t I n a s u p e r s o i ~ i c f l o w a d i s t u r b a n c e i n t h e f l o w is f e l t o n l y i n a be c o n s i d e r e d .

c o n i c a l r e g i o n downstream o f t h e d i s t u r b a n c e known as t h e r e g i o n o f i n f l u e n c e .

F o r t h e p r o p e l l e r shown i n f i g u r e 5 o n l y t h e shaded p o r t i o n o f t h e u p p e r r i g h t hand b l a d e is a f f e c t e d by what happens a t t h e t i p o f t h e u p p e r l e f t hand blade.

Thus, when t h e induced v e l o c i t y d u e t o t h e wake is c a l c u l a t e d , it is n e c e s s a r y t o l i m i t t h e r e g i o n o v e r which e a c h wake f i l a m e n t h a s a n e f f e c t . A second con- s i d e r a t i o n is t h a t , when t h e t i p is s u p e r s o n i c , t h e £ 1 3 w becomes h i g h l y t h r e e - d i m e n s i o n a l n e a r t h e t i p d u e t o t h e t i p Mach cone. T h i s e f f e c t is t a k e n i n t o a c c o u n t by a p p l y i n g a c o r r e c t i o n t o t h e l i f t d e t e r m i n e d from i s o l a t e d a i r f o i l d a t a .

The r e l a t i v e importance o f some o f t h e s e e f f e c t s is shown i n f i g u r e 6 where p r e d i c t e d power c o e f f i c i e n t i s s h o r n a s a f u n c t i o n o f advance r a t i o . The d i f - f e r e n t c u r v e s w e r e o b t a i n e d u s i n g t h e same computer program b u t w i t h d i f f e r e n t o p t i o n s f o r e a c h c u r v e t o i s o l a t e c e r t a i n e f f e c t s . The c o n f i g u r a t i o n a n t l y z e d was a n e i g h t b l a d e d p r o p e l l e r w i t h 300 o f t i p sweep ( d e n o t e d as SR1 i n t h e The c u r v e l a b e l e d " r i g i d wake" w a s o b t a i n e d u s i n g a r i g i d h e l - p r e v i o u s paper).

i c a l wake w i t h o u t t h e high-speed e f f e c ~ s and i s e s s e n t i a l l y t h e same as t h e re- s u l t s t h a t would be o b t a i n e d w i t h t h e e s t a b l i s h e d appro act^. F o r t h e c u r v e l a b e l e d "wake model," t h e wake f i l a m e n t s w e r e d i s t o r t e d by t h e n a c e l l e . (The The d i f f e r e n c e between t h e two high-speed e f f e c t s a g a i n w e r e n o t included.)

c u r v e s is s o l e l y d u e t o t h e d i f f e r e n t assumed wake shapes. The c u r v e l a b e l e d "wake model and high-speed e f f e c t s " i n c l u d e s t h e d i s t o r t e d wake and t h e high- speed e f f e c t s d e s c r i b e d i n f i g u r e s. The d i f f e r e n c e between t h i s c u r v e and t h e "wake model" c u r v e is s o l e l y d u e t o t h e high-speed e f f e c t s . A s e x p e c t e d t h e l a r g e s t d i f f e r e n c e between t h e s e two c u r v e s o c c u r s a t t h e l o w e s t a d v a n c e r a t i o , which c o r r e s p o n d s t o t h e h i g h e s t t i p speed. The high-speed e f f e c t s c a u s e t h e l a r g e s t c h a n g e s t o t h e s h a p e o f t h e c u r v e s . From c o m p a r i s o n s n o t shown h e r e , it was found t h a t t h e s h a p e of t h e wake model and h i g h speed e f f e c t s c u r v e a g r e e s w i t h t h e e x p e r i m e n t a l r e s u l t s , i n d i c a t i n g t h e t r e a t m e n t o f t h e s e e f f e c t s i n t h e a n a l y s i s is q u a i i t a t i v e l y c o r r e c t .

Shown i n f i g u r e 7 a r e r e s u l t s from t h e c i r c u m f e r e n t i a l l y a v e r a g e d , v i s c o u s , c o m p r e s s i b l e f l o w c a l c u l a t i o n f o r t h e c o n f i g u r a t i o n shovn i n f i g u r e 6.

Each c u r v e i n f i g u r e 7 r e p r e s e n t s t h e r a d i a l d i s t r i b u t i o n o f c i r c u m f e r e n t i a l l y a v e r - wed r w i r l v e l o c i t y a t t h a t a x i a l l o c a t i o n . The uniform s p a c i n g o f t h e c u r v e s ahead o f t h e p r o p e l l e r i n d i c a t e s n o swirl i s p r e s e n t i n t h e flow. The d i s t o r t e d curves i n t h e v i c i n i t y of t h e b l a d e s i n d i c a t e t h a t s w i r l is b e i n g i n t r o d u c e d i n t o t h e flow. The uniform s p a c i n g o f t h e c u r v e s downstream o f t h e b l a d e s i n d i - cates t h a t t h e s w i r l p e r s i s t s i n t h e flow. These r e s u l t s fror t h e v i s c o u s cal- c u l a t i o n are used t o check f l u i d v e l o c i t i e s between t h e b l a d e s and downstream o f the p r o p e l l e r . I f t h e v e l o c i t i e s are t o o high, l a r g e losses d u e t o shock waves These r e s u l t s a l s o g i v e t h e p r e s s u r e and v i s c o u s d r a g o n t h e n a c e l l e can occur.

i n t h e presence o f t h e p r o p e l l e r . United Technologies Research C e n t e r w i l l be doing m o d i f i c a t i o n s and a p p l i c a t i o n s o f t h i s a n a l y s i s under a f o l l o r o n c o n t r a c t .

Lif ting-Surface A n a l y s i s The key f e a t u r e s of t h e three-dimensional, compressible l i f t i n g - s u r f a c e a n a l y s i s are shovn i n f i g u r e 8. Also shouo are p a r t i a l f r o n t and s i d e views of t h e g r i d o n which t h e f l o w c a l c u l a t i o n s are performed.

The g r i d a c t u a l l y ex- t e n d s much f u r t h e r i n t h e ' r a d i a l d i r e c t i o : ~ t h a n is shown. The n a c e l l e is re- q u i r e d t o be a x i s y u m e t r i c s o t h a t t h e flow b e t w e n e a c h two a d j a c e n t b l a d e s is t h e same. i h u s it is o n l y n e c e s s a r y t o s o l v e f o r t h e flow between two blades.

The f l o v is r e q u i r e d t o be t a n g e n t t o a l l s o l i d s u r f a c e s end beyond t h e b l a d e t i p s is assumed t o be periodic. The e q u a t i o n s of motion i n f i n i t e - d i f f e r e n c e The e q u a t i o n s t h a t are s o l v e d form, are s o l v e d a t d i s c r e t e p o i n t s i n t h e g r i d .

are t h e three-dimensiona 1, unsteady, E u l e r e q u a t i o n s , which govern t h e i n v i s c i d f l o v o f n compressible f l u i d and c a n a c c u r a t e l y r e p r e s e n t t h e p r e s s u r e v a r i a t i o n caused by shock waves and t h e work done by t h e p r o p e l l e r on t h e f l u i d . The e q u a t i o n s a r e s o l v e d by marching i n time u s i n g a n i m p l i c i t f i n i t e - d i f f e r e n c e method u n t i l a s t e a d y state is reached. No wake modeling o r two-dimensional a i r f o i l d a t a are r e q u i r e d . Viscous e f f e c t s , however, a r e not included. T h i s a n a l y s i s was developed by K u t l e r of N A S A Ames Research C e n t e r and Chaussee of Nielsen Engineering and Research and i s d e s c r i b e d i n r e f e r e n c e 6.

R e s u l t s from t h i s l i f t i n g s u r f a c e a n a l y s i s a r e shown i n f i g u r e 9 f o r a n e i g h t bladed p r o p e l l e r having 30° of sweep a t t h e b l a d e t i p s . The p l o t s show t h e d i s t r i b u t i o n o f s t a t i c p r e s s u r e c o e f f i c i e n t on t h e s u c t i o n and p r e s s u r e s u r f a c e s o f t h e b l a d e s a t t h r e e spanwise l o c a t i o n s from n e a r t h e hub t o n e a r t h e t i p . The most s i g n i f i c a n t f e a t u r e of these r e s u l t s is t h e p r e d i c t e d shock wave along t h e e n t i r e span o f t h e blade. A t t h e c o n d i t i o n s f o r which t h e s e r e s u l t s were o b t a i n e d , experimental d a t a a l s o i n d i c a t e c o m p r e s s i b i l i t y l o s s e s . The de- t a i l e d spanwise and chordwise d i s t r i b u t i o n of l o a d i n g p r e d i c t e d by t h i s code i s important f o r improved p r o p e l l e r d e s i g n s from a c o u s t i c and s t r u c t u r a l stand- points. Development of t h i s code is c o n t i n u i n g a t Lewis.

FUTURE PLANS The developmert of t h e s e advanced a n a l y s e s w i 11 continue. I n i t i a l compar- i s o n s of t h e a n a l y t i c a l r e s u l t s from a l l t h e s e advanced methods w i t h performance d a t a have shown q u a l i t a t i v e agreement ( ref. 7). However, performance d a t a can- n o t s u b s t a n t i a t e t h e d e t a i l s of t h e flow a s p r e d i c t e d by t h e a n a l y s e s . Thus a n experimental program i s planned f o r t h e Lewis 8x6 f o o t wind t u n n e l i n 1980 t o provide d e t a i l e d d a t a f o r v e r i f y i n g t h e s e a n a l y s e s . A l a s e r v e l o c i m e t e r system ( f i g . 10) w i l l be used t o make t h e s e measurements s i n c e t h i s t y p e of system d o e s T h i s e x p e r i m e n t a l progrsm n o t i n t r o d u c e hardware which might d i s t u r b t h e flow.

w i l l d e f i n e t h e d e t a i l s of t h e f l o w around t h e b l a d e s and upstream and down- These r e s u l t s w i l l p i n p o i n t any d e f i c i e n c i e s i n t h e s t r e a m of t h e p r o p e l l e r .

a n a l y s e s s o t h a t q u a n t i t a t i v e , as w e l l a s qualitative, agreement c a n be o b t a i n e d .

Three advanced a n a l y s e s f o r p r e d i c t i n g p r o p e l l e r aerodynamic performance have been presented. The a n a l y t i c a l approaches as w e l l a s t h e c a p a b i l i t i e s o f t h e s e a n a l y s e s have been d e s c r i b e d . Two o f t h e s e a n a l y s e s u s e a l i f t i n g - l i n e r e p r e s e n t a t i o n f o r t h e p r o p e l l e r b l a d e s , and t h e t h i r d u s e s a l i f t i n g - s u r f a c e r e p r e s e n t a t i o n . The d e t a i l e d f l o r £ i e l d measurements t o be made i n t h e n e a r fu- t u r e w i l l provide d a t a f o r v a l i d a t i n g t h e a n a l y s e s , making them a v a i l a b l e as a n a l y t i c a l t o o l s f o r d e s i g n i n g improved p r o p e l l e r s .

REFERENCES 1. G o l d s t e i n , S. : On t h e Vortex Theory of Screw P r o p e l l e r s . Proc. Roy. Soc., (London), vol. 123, no. 792, Apr. 6, 1929, pp. 440-465.

2. Theodorsen, Theodore: Theory of P r o p e l l e r s . ~ c G r a r H i l 1 Book Co. , Inc. , 1948.

3. Kuethe, Arnold M and S c h e t z e r , J. G. : Foundat i o n s o f Aerodynamics. John Wiley 6 Sons, Inc., 1950.

4. S u l l i v a n , J. P. : The E f f e c t o f Blade Sweep on P r o p e l l e r Performance A I A A Paper 77-716, June 1977.

5. Egolf, T. A. ; e t a l . : An A n a l y s i s f o r High Speed P r o p e l l e r - N a c e l l e Aero- dynamic Performance P r e d i c t i o n . Vol. 1 - Theory and I n i t i a l Application.

Vol. 2 - U s e r ' s Manual f o r t h e Computer Program. R79-912949-19, Dnited

Technologies Research Center.

6. Chaussee, D. S.: Computation of Three-Dimensional F l o v Through Prop Fans.

NEAR-TR-169, Nielsen Engineering and Research, Inc. , 2979.

7. Bober, L. J. and M i t c h e l l , G. A. : Summary o f Advanced n e t h o d s f o r P r e d i c t i n g High Speed P r o p e l l e r Performance. A I A A Paper 80-0225, Jan. i980; a l s o NASA T M 81409, 1980.

ESTABLISHED APPROACH

M O M L

WAKE - RIGID HELICAL VORTEX SHEET

SINGLE OR CWNTER ROTATION

BLADES - STRAIGHT LIFTING LlNE

SOLUTION TECHNIQUE STRIP ANALYSIS Figure 1

CURVED LIFTING LlNE ANALYSIS

M O M L

WAKE - RIGID HELICAL VORTEX FILAMENTS

S ING E ROTAT ION BLADES - CURVED LIFTING LlNE

NACELLE - INFINITE CYLINDER

SOLUTION TECHNIQUE SIMULTANEOUS SOLUTION FOR BLADE AND WAKE VORTEX STREXTHS Figure 2 PROPELLER AEROACOUSTIC METHODOLOGIES* Kenneth D. Korkan and Gerald M. Gregorek The Aeronautic;al and Astronautical Research Laboratory Department of Aeronautical and Astronautical Engineering The Ohio State University This paper briefly covers aspects related t o propeller p e r f c m n c e by means of a review of propeller methodologies: ?resentation of preliminary wind tunnel propeller perfomar.ce date taken i n t'le N A S A Lewis Research Center 10 x 10 wind tunnel; discussion of the predominent limitations of existing propeller performance methodologies; and a brief review of a i r f o i l developments appropriate for propeller applications. This paper is intended a s a s t a t u s report with the complete study t o be documented a t a l a t e r date.

INTRODUCTION Because of the increased emphasis on fuel efficiency f o r general aviation aircraft, thcre has been a renewed interest i n the use of propellers. It has been estimated that in the use of the prop fan concept (Ref. I ) , a fuel savings of approximately 36% can be realized over the turbofan through proper propeller design. Also, recent studies have shown a 5 fo 7% savings in f u e l efficiency can be obtained (Ref. 2 ) through proper propeller design and c r f t i c a l examina- tions of propeller-nacelle interactions. As a result, a study supported by the National Aeronautics and Space Administration Lewis Research Center was i n i t i a - ted involving the Ohio State University, Borst and Associates, Hartzell Propeller, Inc., and Rockwell Corporation of Bethany, Oklahcma t o evaluate and eriance current analytical prediction methods f o r propellers designed specifi- cally for the twin engine Rockwell Aerocmander 690B. This three year study has and w i l l involve computer prediction studies i n the theoretical evaluation of 2ropeiler performance; wind tunnel model t e s t s conducted a t the NASA Lewis 3esearch Center; f l i g h t t e s t comparisons; and enhancement of the theoretical methods by means of ccmparison with wind tunnel and flight t e s t s . The intent of t h i s paper is t o briefly cover aspects related t o propeller performance and t o i l l u s t r a t e preliminary data resulting f r m the wind tunnel t e s t s of two propellers i n this study. It i s intended that a report w i l l be made atailable on the canparisons between the theoretical predictions and the complete experi- mental data s e t resulting f r m the wind tunnel t e s t s .

*This study was funded by NASA Lewis Research Center under NASA Grant NSG 3247.

PROPELLER PERFORMANCE bdETHODS A brief review of the methodologies ( ~ e f . 3 ) used i n predicting propeller performance ( ~ i g . 1 ) has been included in this work for completeness. Pro- peller theories have proceeded from the simple Rankine-Froude momentum disc theory (Refs. 4 , 5 ) uhich assumes that the propeller disc is replaced by a disc with an i n f i n i t e number of blades producing a uniform change i n velocity of the stream passing through the disc. This th2ory is useful i n calculatin- theore- t i c a l maximum efficiencies but does not Ceal in the details of t?. propeller configuration such a s number of blades snd blade thickness. These factors are considered i n the blade element analysis (Refs. 6,7) the next degree of sophistication, which deals i n the forward and rotational velocity components t o determine the resultant velocity or the effective pitch angle and hence the angle-of-attack a s seen by each a i r f o i l sention making up the propeller blade.

Here the angle-of-attack i s taken a s the difference between the geometric pitch angle and the effective pitch angle (Fig 2 ) and a s m s that the induced flow past the blade element is the same a s past a wing with an aspect r a t i o cf six.

The simple blade element theory has been used f o r preliminary calculai:',nr,s a d in some eases gives accurate answers within 10% of the measured thrust and torque values.

More precise results may be obtahed i n the prediction of thrust and torque by calculaling the local induced velocities a t each radial station

(Fig. 2 ) by means of vortex theory ( ~ e f . 8). Here, the combination of simple

momentum theory and blade element analysis r e s u l t s i n a theory t h a t also accounts for rotation of t h e s l i p stream. However t h i s approach, although pro- still does not account f a r t i , p losses, blade viding an accurate approximation, t o blade interference, and nonuniform flow i n the disc plane resulting from the presence of 2 nacelle.

The next order of development and accuracy cane with the Goldstein liftirg- l i n e model ( ~ e f . 9) where the blade is replaced by a s e r i e s of horseshoe vor- t i c e s as sham i n Figure 1. The approximation of blade replacernen-:, by vortices is acceptable since most generc.1 aviation propellers have a relatively high aspec+, ratio. Also, the l i f t i n g l i n e approach can u t i l i z e corrections f o r viscosity and canpressibility but is accepted a s an "approximate method" using the Goldstefn factor. The Goldstein factor mthod is usually taken f o r l i g h t l y loaded propellers where the Betz condition holds, =d does not apply t o other than constant pitch propellers i n uniform flow (Ref. 3 ) . The l i f t i n g l i n e problem can also be solved by the "rigorous method!' using Lerbfs induction

factor method (Ref . 9 ) which is based on the velocity potential of helical

vortex lines applied t o ai3y moderately loaded optimum or non-optimum propeller operating i n a uniform or nonuniform free stream (Ref. 3 ) . This is the method of analysis that has been used i n the performance comparisons t o be shown i n a l a t e r section of t h i s paper.

A s the propeller configurations change t o relatively small apsect r a t i o and/or large surface areas (Fig. I ) as i n the prop-fan concept (Ref. I ) , advsnced analytical methods must be used such as the Ludwig-Ginzel l i f t i n g sur- face model (Ref. 10) t o model the propeller f low f i e l d accurately. These advanced methods and the current state-of-the-art have been discussed by Bober and Mitchell (Ref. 11) in addition t c the importance of wake modeling.

It is the purpose of the present effort t o compare directly with experi- mental data the theoretical predicticm r e s u l t s of vortex theory and l i f t i n g line theory t o determine the ranges of applicability and levels of accuracy.

The current wind tunnel t e s t s cover a broad range i n advance ratios, blade angle settings, and f l i g h t conditions for four general aviation propellers, each having different activity factors and propeller blade sections. In so doing, the current methods may be enhanced t o pro-ride increased accuracy i n the predic tian of propeller performance.

PRfiPELLER AIRFOIL DEVELOPMENT Airfoil development f o r propeller applications has been limited with the continual use of the Clark Y and RAF 6 series a i r f o i l s . The l a s t major develop ment i n t h i s area occurred with the development of the NACA 16 s e r i e s a i r f o i l s (Ref. 1 2 ) and a s shown i n Figure 3 does have relatively good prformance i n terms of the metric CL/CD a s a function of CI,. This a i r f o i l has +,he character- i s t i c " f l a t bottom1', m a x i r m thickness occurring a t approximateiy tke 50% point, and a small leading edge radius with many of the design characteristics dicta- ted by manufacturing constraints. Therefore many propellers of today incorpor- a t e the Clark Y or RAF 6 a i r f o i l series during the i n i t i a l 50% of the blade transition- t o the NACA 16 s e r i e s which has a high drag divergence Mach number in the outer segment of the propeller where the resultant Mach numbers can approach upity.

Bocci ( ~ e f . 13) i n a paper published i n 1977 described a new series of propeller a i r f o i l sections entitled the A9.A-D series. Here, the manufacturing constraints have been relaxed a s shown i n Fi-gure 3 resulting i n a section incorporating increased camber on the underside of the a i r f o i l ; drooped leading edge t o prevent leading edge s t a l l a t high angle-of-attack; and an increased leading edge radius. The r e s u l t s of t h i s design approach can be seen in Figure 3 with an improvement over the performance of the NACA 16 series a t the high l i f t coefficients.

The importance of the a i r f o i l section t o propeller performance is indicated i n Figure 4 where it can be seen that the a i r f o i l pressure distributions which evolve i n t o the aerodynamic coefficients determine the load distribution and also allows an acoustic evaluation by the s t r i p method.

In a l a t e r study, the authors ( ~ e f . 14) have compared the aerodynamic performance and acoustic estimates of the ARA-D, Clark Y and NACA 16 series a i r f o i l s .

In the discussions gf propeller a i r f o i l development, the w i l d tunnel t e s t s of the propellers i n t h i s study incorporate a variety of a i r f o i l sections, i.e.,

Clark Y - NACA 16 a i r f o i l s

( a ) (b ) ARA-D a i r f o i l s ( c ) G A ( w ) a i r f o i l s ( d ) 6 series a i r f o i l s Since a l l have been designed for the Aerocomander 690 B, a comparison of the propeller performence can be interpreted as a cmpariaon of these a i r f o i l s in terms of efficiency ( s ! , thrust coefficient (CT), and power coefficient (Cp) which t r e discussed i n the following sections.

EXPERIMENTAL PROPELLER PERFOWCE ~ ~ N T S Propeller performance experimental values were obtained i n the presmt

program through use of the Propeller Test R i g ( PTFt ) ( ~ e f . 15) installed i n the

subsonic leg of the 10 foot x 10 foot supersonic wind tunnel located a t the NASA Lewis Research Center. The configuration tested also incorporated equiva- lent body of revolution representatin@-the actual nacelle of the Rockwell Aeroc-der 690 B including a scaled representation of the spinner (Fig. 5 ).

Pressure orifices were located along the periph,-y of the nacelle a t two azimu- thal locations to aid in evaluating the drag of the nacelle and its effect on the performance of the propeller.

Three f l i g h t conditions were examined f o r the approximately 0.5 scale pro- pellers, i.e., take-off (M = O.11), climb (M = 0.23), and the cruise condition The advance r a t i o (J) was varied f o r a fixed blade angle setting (M = 0.39).

by fixing the t e s t section Mach nuniber through manipulation of the wind tunnel second throat and changing RPM. Values of propeller thrust and torque rere de&dced from the experimental measurements and the thrust coefficient ( CT ), torque coefficient (CQ), power coefficient (Cp), and efficiency ( rl ) were deter- mined by t h i s method.

An appropriate range i n J values w a s examined with respect t o the actual operating conditions o r u n t i l s t a l l - f l u t t e r was en- countered.

The f i r s t propeller tested on the FTR i n the configuration s h m in Figure 5 consisted of Clark Y-NACA 16 a i r f o i l s with an activity factor of 101. The preliminary results are shorn i n Figure 6 f o r the cruise condition (M = 0.39, B = 4e0) in terms of efficiency ( q ) a s a function of the advance r a t i o (J).

Also shown i n t h i s figure are the theoretical estimates using vortex theory* and l i f t i n g line theory previously discussed. -9s can be seen, at the lower J values the l i f t i n g l i n e prediction coincides with the experimental data with a resulting overprediction for J values i n excess of 2.3. T h i s may be compared directly with the vortex theory results which overpredicts the experimental data over the enti.re range of J values. A similar r e s ~ l t is also found f o r an off-design condit~on as shown i n Figwe 7. Consideration of the climb condi- tion (M = 0.23, 8 = 3 2 ' ) f o r t h i s propeller, shown i n Figure 8, indicates acceptable agreement between experiment and l i f t i n g line theory ctrer the range in J valu,?s. Here vortex theory m e e s well with the experimental data a t the low J values with disagreement occurring a t J values in excess of 1.2. A s found previously, an investigation of the off-design condition as shown i n Figure 9 also produces similar results.

B ~ h e vortex theory i s presently under examination t o include the influence of the blade-spinner interference which could result i n better correlation with experiment.

The acond propeller tested having an activity factor of 83 utilized the ARA-D a i r f o i l sections previously discussed. The i n i t i a l canparison8 are sbaar in Figure 10 and k d i c a t e t h a t the lifting lice prediction pmvides a reasar- able correlation w i t h experiaental data f o r t % cruise condition a t a 8 of @.

Here agafn, the vortex theory overpredicts that of the experimental data, i camplete s e t of data including camparisom with theory f o r CT, CQ, n , and Cp k i l l be published f o r all four m l l e r s tested, Determination of the range of vafiditg of these theories i n conparisan with experimental data can then 3e investigated. Also, u t i l i z , , t i m of a rdte rake probe (Fig. 11) is presently in use to obtain meas~~e=i.ts: of t o t a l pressure deficit, flow aagu- larity, arrd s t a t i c pressure meamrattents behind the disc plane of each pro- peller tested as a function of radial lacation. These data w i l l result i n indepeudent thrust measurements as w e l l a s d e t a i l s of the prapeller wake rhich can be c q m - e d directly to the current theoretical rake model being used.

=%se results rill be included in the reports previwslg mentioned a t a iat-er date.

LIMZTATIrn OF CURRENT AmLYSEs The theorezical analyses that have been utilized in the comparisons with experimental data previousiy discussed a r e analytical models which contain limitations. For example, the importance of an accurate rake model and pro- peller/nacelle interactions has teen enphasized by Bober (Ref. 11) in the prediction of high speed propeller performance pivdictions. Further, the re- sult of a f i n i t e blade length, i.e., recognition of t i p flow is necessary f o r an evaluatian of three-dimensional effects. This e f f e c t has been treated by Cooper (Ref. 16) by obtaining a correction factor t o the lift-curve slope a s a ftincticn of the radial location but is valid f o r propellers using only NACA 16 and 6 series a i r f o i l s .

Alm rhen considering l h i t a t i a n s , the area of centrifugal viscous effects on the l i f t coefficient should be considered. In an expx=imental investigation by ) r ! l s k a m p (Ref. 17), he had found that there is a significant relationship between the magnitude of CL and the radial loczi$ion of tlle propeller blade. In a series of t e s t s with a propeller made ap of GO625 a i r f o i l s , fiinunelsbamp fixed the wgle-of-attack a t each radial location and measured the section l i f t co- efficient. These values of CL were then canpared t o the two-dimensional l i f t a = 5 ' case and found t o be co-:- coefficient, as given in F - a r e 12, f o r the siderably higher with the greatest difference occuring a t the root and decreas- ing a s the propeller radius increased. These differences may be attributed t o centrifugal viscous effects which obviously are not accounted f o r in two- dimensional theory. Since a l l propeller performance analyses u t i l i z e an a i r - f o i l data bank based upon two-dimensional experimental and analytical data, the differences indicated in Figure 12 i f properly modeled could have a significant infiuenw in the prediction aczuracy of propeller performance theoretical values and resulting caparisons with experimental data.

As previously indicated, propeller performance analyses u t i l i z e a i r f o i l data banks c ~ n s i s t i n g of wind tunnel t e s t and anblytical computer codes.

Advancements made i n t b e t h e o r e t i c a l analpsis of a i r f s i l s has been considerable r l t h t h e a v a i l a b i l i t y of such s u b c r i t i e a l computer codes a s Smetana, et. al.

(Ref 1 8 ) and Eppler (Ref. 19). The more extreme case of both s u b c r i t i c a l and s u p e r c r i t i c a l flow over an a i r f o i l can a l s o be t r e a t e d and,lytically a s given by Ciabedian, et. a l . (Ref. 20) and Carlson (Ref. 2 i ) . Also : h e results of massive separation M an a i r f o i l . i.e., t h e o r e t i c a l investigations of air- f o i l maximum l i f t ccefficietlt has been under study by Barnwell ( ~ e f . 22), Carlson (Ref. 233, and Dvorak (Ref. 24) .md a r e being used on a limited basfs.

and resulting limitations, To i l l u s t r a t e t h e a p p l i c a b i l i t y of these codes t h e IS( 1 ) - a 1 3 airf .::I t h e o r e t i c a l and experimental ( Ref. 25 ) pressure' d i s t r i - butioxl is s h o m i n Figure 13 f c r hl = 0.755. a = 0°, and Reynolds number of 5.11 x lo6 c a n d i t i m .

The canparison between experiment and theory is reason- a b l e on both the upper and lower surface of t h e a i r f o i l with respect to t h e maximum negatf ce and positfve C magnitude. locaticm of t h e shock rave, and the

T base pressure value. Hcnever, f t h e k c h number is increased to M = 0.802 for

the same condition of a and iie+vnolds number a s shom i n Figwe 14, t h e misaaatch between t l e o r y and experiment is evident indicating d e f i c i e n c i e s i n t h e theore- t i c a l analyses and/or experiment.

PROFELLEI3 AiTiSTIC ANALYSIS WEEL The empliasis has been sn t h e propulsion perfomance of general a v i a t i o n propellers, however recent e f f o r t has resulted i n t h e design of e f f i c i e n t as w e l l a s quiet propellers. T * : : t h i s end an accustic a n a ~ v s i s ( ~ e f . 26) has been derived t h a t , provided t h e pressure fiistributions st severs1 r a d i b l l o m t i o n s along t h e blade a r c speciffed. t h e r e s u l t i n g t.ots1 noise due t~ loading and thickness can be predict.ed. The c h a m c t e r i s t i c acoustic pressure s i g m t u r e s are sham i n Figure 15 f o r the near f i e l d condition. f r a n which t'ne sound pressure l e v e l (dB) a s s - k c t i o n of hamcrlic number o r multiples of t h e fundamental can be calculated (Fig. 16). The sccuncy of t h i s t h e o r e t i c a l approach is s h o m i n Figure 16, which shows the comparison between messured and predicted noise for a series of s t a t i c tests conducted by Huhbard (Ref. 2 7 ) f o r two n e m f i e l d 1 cations. As can be seen, the compnrisl:n is reasonable as f x m d i n .levera1 other applications (Ref. 2 8 ) c 3 f t h i s t h e o r e t i c a l s p p r ~ ~ a c h .

It was intended ir. the current st.uCy t o obtain ncsr f i e l d acoustic measure- ments of t h e propellers tested (Fig. 1"). ilowever. there a r e a series of pro- blems associated with tunnel w,s:l condit-ions that. are curre!lt.ly under stukv before acoustfc data can taken with t.he desired accuracy.

A preliminary summnry of t h e ::t.utiy to d # ~ ' e has irldizntczd that:

-

l i f t i n p 1 i n e snnl,vsis t ~ i v e s d-wersl l bet. t,er agrec.ment, with c s v r i m e n t a l resul t,s:

-

at design "lib. l i f t i ~ g l i n e agrseu v e i l ritb me:~summnt.s but over- predicts crsise perf~rmsnce:

- vortex %fiecry overpredicts e r p e r i r n e n t s l results a t both climb and

z r u i s e w n d i tions:

-

present prediction methods r e q u i r e i m p r ~ r ~ . n e n t It is i5ie:rdod 2hzt ihese wi.ri.3 t;ml &ta he empared tc Full scale flight test during 1QSt;. Also, en'larcenmt cf the present theoretical madeb rill be required 3s irrdicatel? $n t h i s phase of the stIL3,v result,ing in better campari- s.Jn between zxperimental data and arsf ,vticai ?rec;ictions.

1, V!&elson. 3. C.. 9laiia, 3. J . , %!itchell, C. A . , 3nd V i k e t e . J, E., "bs5gn sr?d Ferfzrx;;i.- ~f * e e y Efficimt Fropeliers fcr &zh 3.5 Cxisem. S A E Paper 7'8:c<>;.:. :,+sra:: 1 4 7 (A7,s;. x ! S-f :h>Z).

2 . L e i tsr. I . . "Lzr Sree2 Potential". F's~er yrfise:lted a : . : . ' , h e Gneral Xvktion Wo?ulsi ? : : L'o~t'cren,-z. Xis4 ' k r ; s Zesesrch Center. Xovmber 3-29. L 4 9 .

I P . r~sr,ki?re. 'df.c'.l!. . ",'n t3s Yezh%x-,icsl F ~ i n e i p l e s 2 ' t l k Actics of hopellern.

Trans. Ins:. Xsv. Ar.?h., V21, 6 , 1305.

5 . Vmude, R . Z.. "ch the Part P l s y e d In Frapuision bb- Differezce of F l u i d Presswe'', T r m s . Inst. NRV. Arcfi.. ' 1 ' 3 1 . . id. IES9.

6 . Froude. W . . "On t h e f l m e a t a r y Belstisn %tween Pitch S i i y , m d Prapufsive

-

SffIcEcncy". Lrans. f n s t . N3v. Arch.. l;.~l. 19. IS"$.

5. h i ! ! e s t e r , E . W.. Aer.o&..~~?rnic::. i,>nstablc Sr Carn?m;~. Lzil., 'London, 19~37.

3. 1 . 5.. ' e V i Y e . f ::k*~~,r f rc7~~lle~::". f raceedings of :.he 9 o ~ - a l .q.?,.iei.:: ( Londcn). S e r i e s :I. Val. c 3 . l Q , " ? .

9. krSs. H. W., "?!oderately Lmded P r o r e l l c r s W i t h a F i n i t e Nw.bor of Blades and sn :lrbi:.rsry l7istribution of C i r c u i s t i c ! ? " . Trans. Tht. .Sc?ciet,v of Naval A r c h i t e 2 t . s rind %farin;: Engineers ( ShlUE). Vo: . G Q . 1052.

10. Ludwiet:. H . . :1r;;2 Ginzel. I.. "(Tn the T h e o w (3:' Screws %L'c.h Wide Blades", :!eroi:mmisr'!?e Verszc1:senstalt. Goet tiW:e!l, Report .!;/:I 435, la.;4.

I' C 11. Fober. 2 . .r. . .inti '.'it ? h e l l . ~ 2 . A . . .-t~rn:1ry cf Aiiv.;nL3c,d !,!etiiL~d:: f o r Fre- diL:t.lnc 3 i::!~ Syced Propel! i.r !'crfnnrwi;.t:", ?,!AA .'.'0-0,':5. AIAII E i ghteenth : I e rcs!~ace : ; : i t.nl.e:: ? & : > t,tnq, i':\::~,ic;i:~, (::I 1 i Ckl r ! ~ i 3 , <?311\ig r~ ~ Q S C ? .

32. Lindsey, V . F . , Stwensun, D . B . , and Ihley, B . , N., mAemdgnemic Characteristics d 24 NACA l6-Series Airfoils a t h c h Nunbers Between 0.3 and 0.8", UCA TW 1546, lrecmber 1 9 4 7 .

for Aircraft

Bocci, A. J. , "A New Series of firfoil Sectiaos Suitable

Ropellers", A e r a m u t i d Quarterly, February lQn, pp. 5+73.

Korkan, K . D . , ban, C . J., andGregorek, G. M . , "Effect of Airfoil Sections c m Acoustic Perfcamsrice of Pmpellers", Paper presented t o the Advauced Technology firfoil Research Conference, NASA Laagley Research Center, Yarcb 1978.

Jeracld, R . , "High Speed Results", Paper presented a t the General Aviation

Prapuldan Catferena, H A S Lewis Research Center, November 28-29, 1479.

Cooper, J. P. , " l 7 e Linearized hr'lorl Pmpeller Strip Analysis", Wright

M r -lopreat Center, ldDC TR 56415, February 19W.

M s k M p , E . , " ~ f ' i l u n t e ~ h u o g e n an einemaPlanfenden Propeller", Tbesis Gttingen 1x5. Report of the Mm-Planck-Inst?tut f & S t ~ s ~ 3 r i 3 c h u n g , Giittiagen, Wo. 2, 1950.

ataaa, P . O., -, D . C., mth, N, So, andcarden, R . K O , "Light

Aircraft Lift, m, and Yr#nt Prediction - A Review and Analysisn, NASA CR-2523, 1375.

Eppler, R., "Private -cationn, 1976.

Bawr, F., Gsrabedian, P., Korn, D., and Jameson, A., Supercritical Wing Sections 1 1 , A Hsndbook, Iactuw Notes in Econdcs axx3 Mathematical Systems, V. 1 0 8 , Springer-Verlag, New York ( 1975 1.

Carlscm, L. A., "Transonic Airfoil Analysis and Design Usirg Car3isia.n Coordinates", J. of Aircraft, Vol. 13, No. 5, May 1976.

Barnwell, R., "Private Camnmicatia", 1978.

Carlson, L. A., "TRBWIFS: A FC#lTRAN Program for Transonic Airfoil Analysis o r Designw, NASA CR-2821, June 1977.

Dvorak, F., "Rivate CoPm\micatimn, 1979.

Lee, J. D., Gregomk, G. M., end Korkan, 8 . D., "Testing Techniques and Interference Evaluation i n the OcSU Transonic Zdrfoil Facility", AIAA 78-118, AIAA l l t h Fluid and Plasm Dynamics Conference, July 1978.

W m n , C. J., and Gregorek, G. H . , "The Exact Numerical Calculation of Propeller Noiseg, AIAA Paper No. 78-1122, AIAA l l t h Fluid and Plasma Dynamics Conference, Seattle, Washington (July 1978).

2 7 . Hubbard, H . H . , and Regier, A. A., "Free Space Oscil3~~35rg Pressies Kear tbe T i p s of EZotating Propellersn, NACA TN 1870, 1949, 28. Korkan, K . D . , Gregorek, G. M . , and Keiter, I., "General Aviatim Pro- Paper presented to the Sixth Annual peller Study - Acoustic Analysis", General Aviation Technologyfest , Wichita, Kansas, November 1979.

M # M DISC THEDRY Figure 1 INUUCOO VRDCllY f i - VORTEX TSORY

- LlFTlA L l l L

AIRFOlL CHARACTERISTICS

- TEST

- ANALYTICAL CODES

ROTAT IOML VEL

- DATA BANK

Figure 2 AIRFOIL PERJRMNCE tWPP.RISOI( tN = 0 . 4 T/C = lor>' 0 I L 1 0 a25 ~ ‘ 9 Q 7 5 LOO L25 L SO

- LIFT COEFFICIEWT

Figure 3 M A D E AIRFOIL ANALYSIS

- SECTION AERODYWHIC PROPERTIES

PRESSURE DISTRIBUTIONS

- ACIXiSTiC PRESSURE SIGNATURE

I

Figure 4 OFF-DESIGN CRUISE f'ERFORfWNCE CMPARISOII

H = 0.39/ A F = 101/ CLARK Y - W C A 16 PiPFOlLS

VORTEX TnEORV PREDICTION LIFTING LINE PRED!CTION WIND TUNNEL TEST

-

-

6 0

0 DESIGN POINT

a

-

-

4 0 _P = 36*

I

J - ADVANCE RATIO Figure 7 C L I m PERFOW.NCE COWARISON

N = 0.23/ AF = 101/ CLARK Y - NACA 1 6 A I R F O I B

100 I I I C-) DESIGN POINT VORTEX THEORY PREDICT104 LIFTING LINE PREDlCTIOlY WIND TUNNEL TEST F-60 = P* 0. 4 a 8 L 2 L 6 2 . 0 2.4 2.8 J - ADVANCE RATIO Figure 8 NASA MOPELLER NOISE RESEARCH

George C . Greene

National Aeronautics a n d Space Administration Lang ley Research Center General A v i a t i o n A i r c r a f t represent a cost e f f e c t i v e s o l u t i o n t o many o f the pub1 i c ' s t r a n s p o r t a t i o n problems. Due t o t h e i r f u e l e f f i c i e n c y , ~ r o p e l l e r d r i v e n commuter a i r c r a f t are appearing i n ever increasing numbers, replacing j e t CTOL's on s h o r t block time, low passenger d e n s i t y routes. Business a i r - c r a f t have experienced a steady growth i n recent years. Add these a i r c r a f t t c the l a r g e number of p r i v a t e l y owned general a v i a t i o n a i r c r a f t which are already i n operation and the r e s u l t i n g propel l e r noise represents a growing n a t i o n a l problem.

The purpose of NASA's p r o p e l l e r noise research program i s t o provide a technology base f o r reducing p r o p e l l e r noise w i t h a minimum o f perfonance, weight, and economic penal t i e s . The t h r u s t s o f t h i s program are shown sche- m a t i c a l l y i n Figure 1. Noise p r e d i c t i o n technology represents the most basic p a r t of the orogram. The emphasis o f t h i s a c t i v i t y i s on t h e understanding of and p r e d i c t i o n o f p r o p e l l e r noise vsing basic p r i n c i p l e s o f physics.

Deficiencies i n the p r e d i c t i o n process i d e n t i f y areas where f u r t h e r research i s needed. New research r e s u l t s are incorporated i n the noise p r e d i c t i o n process u n t i 1 predicted w s u l t s are s a t i s f a c t o r y . Engineering noise p r e d i c t i o n methods can then be developed.

P r o p e l l e r noice/performance o p t i m i z a t i o n studies emphasize the development c f p r a c t i c a l p r o p e l l e r design techriology. The c u r r e n t program i n t h i s area u t i l i z e s single-engine a i r c r a f t . Future e f f o r t s w i l l include l a r g e r t w i n engine a i r c r a f t . New design techno1 ogy wi; 1 be demonstrated w i t h f 1 i g h t programs as required.

The t h i r d program area i s i n t e r i o r noise reduction. Research t o p i c s include d e f i n i t i o n o f the source iriput t o the fuselage sidewall, evaluation o f sidewall transmission characteristic; f o r d i f f e r e n t types o f s t r u c t u r e s , and development and evaluation o f advanced noise c o n t r o l treatments. This research i s e s p e c i a l l y important i n view o f the high predicted noise l e v e l s f o r advanced high speed p r o p e l l e r s .

This paper w i l l describe the currel?t research program i n p r o p e l l e r noise prediction, noise/performance optimization, and i n t e r i o r noise r<duction.

Selected r e s u l t s w i l l be presented t o i l l u s t r a t e the status o f c u r r e n t tech- nology and the d i r e c t i o n of f u t u r e research.

Propel l e r Noise P r e d i c t i o n Technology Some c h a r a c t e r i s t i c s o f the propel l e r noise p r e d i c t i o n e f f o r t are shown i n Figure 2. The technology being developed i s applicable t o low and high speed p r o p e l l e r s . It i s based on the basic physics o f the noise generation process, r a t h e r than empirical methods. The techriology i s r e l a t i v e l y sophisticated t o permit analysis of complex c o n f i g u r a t i o n s such as t h a t proposed f o r a h i g h speed turboprop. Noise p r e d i c t i o n r e q u i r e s a knawledge of the propel l e r aeometry and a d e s c r i p t i o n of the aerodynamic c h a r a c t e r i s t i c : of the p r o p e l l e r . Examples of n o i s e c a l c u l a t i o n s using t h i s technoloa,~ a r e shown i n Figures 3 and 4. Figure 3 shows a comparison o f m e a s u ~ d and c a l - culated noise f o r a Twin O t t e r a i r c r a f t . Sound pressure ?sue1 i s shown as a function of frequency expressed i n mu1 t i p l e s o f the blade passage frequency.

The acoustic data were taken i n the plane of the p r o p e l l e r w i t h a microphone mounted on a boom on the a i r c r a f t wing. The measured data i n c l u d e n o i s e from sources o t h e r than the p r o p e l l e r , b u t i n general t h e agreement i s very good.

Typical r e s u l t s f o r an advanced high speed p r o p e l l e r c o n f i g u r a t i o n are shown i n Fiaure 4 . Again sound pressure l e v e l i s presented as a function o f blade passage hnrmonic number. The measured data were taken i n an acoustic wind tunnel ilsing a four-bladed model o f the p r o p e l l e r c o n f i g u r a t i o n s h a m i n the r h o t o i n s e r t . This p r c p e l l e r c o n f i g u r a t i o n i s known as the SR-3.

l e v e l i s very good with The sgreement between theory and data f o r the o v e r a l l some e r r o r s occuring a t the h i g h frequencies. The causes of t h i s e r r o r a r e under i n v e s t i g a t i o n and w i l l probably r e s u l t i n refinements t o the predictior.

technique.

P r o p e l l e r Koise/Performance Optimization C h a r a c t e r i s t i c s o f the propel l e r noise/performance o p t i m i z a t i o n program are show1 i n Figure 5. This i s a j o i n t NASA/EPA program t o demonstrate t h a t p r o p e l l e r noise can be reduced i n 3n economical?y reasonable manner.

The goal o f t h i s e f f o r t i s t o reduce l i g h t a i r c r a f t p r o p e l l e r noise by 5 dSA w h i l e maintaininq o r imprcving propeller- performance. The e f f o r t c o n s i s t s o f 1) o p t i m i z a t i c n studies t o assess the p o t e n t i a l noise and performance benefits o f various p r o p e l l e r parameters, 2 ) wind tunnel t e s t s t o v e r i f y desian concepts. and 3) f l i g h t t e s t s t o demonstrate the noise r e d u c t i o n tecptlo l ogy. Para1 1el e f f o r t s are being conducted a t Massachusetts I n s t i t u t e o f Technology and Ohio State U n i v e r s i t y . Some r e s u l t s from t h i s program a r e shown i n Figures 6-13.

Figure 6 shows the e f f e c t o f varying the p r o p e l l e r diameter.

This assumes a constant s h a f t rpm so t h a t t h i s i s e s s e n t i a l l y the e f f e c t o f varyin9 pro- p e l l e r t i p speed. Noise i n ternis of dBA and e f f i c i e n c y are shown s u b j e c t t o the assumptions 1 i s t e d on the f i g u r e . For each c a l c u l a t e d p o i n t , the p r o p e l l e r was optimized f o r t h a t p a r t i c u l a r diameter. As can be seen both the noise and perfomance are q u i t e s e n s i t i v e t o t h i s parameter. A small percentage reduction i n propel l e r diameter can r e s u l t i n a verv s u b s t a n t i a l noise reduction. E f f i c i e n c y i s a l s o compr-oniised b u t n o t t o the same e x t e n t t h a t the noise i s reduced. A reduction o f pt-ope1 l e r diameter must be acconipanied by other parameter chnnqes i t the p r o p e l l e r e f f i c i e n c y i s t o remain constant.

The c a l c u l a t e d e f f e c t o f varyinq the number o f p r o p e l l e r blades i s shown i n Figure /. Subject t o the l i s t e d assumption, t h i s i n d i c a t e s t h a t noise can be reduced by increasing the number o f blades. P r o p e l l e r e f f i c i e n c y i s not ckci~~ged s i g n i f i c a n t l y by changing the number o f blades. It should be remembered t h a t these are c a l c u l a t e d r e s u l t s and do n o t contain the e f f e c t s o f blade interference a t the l a r g e r blade numbers. Figure 8 shows the calculated effect of varying r a d i a l load d i s t r i b u t i o n on the blade.

Subject t o the assumptions l i s t e d , i t i s shown t h a t noise can be s u b s t a n t i a l l y reduced by moving the peak o f t h i s load d i s t r i b u t i o n inboard. There i s an optimum l o c a t i o n which r e s u l t s i n maximum p r o p e l l e r e f f i c i e i t c y , however the e f f i c i e n c y i s n o t very s e n s i t i v e t o small changes i n the p o s i t i o n o f t h i s peak loading.

Figure 9 shows the c a l c u l a t e d e f f e c t o f blade sweep on p r o p e l l e r noise.

The c a l c u l a t e d p o i n t s are f o r sweep angles from zero degress, which represents a s t r a i g h t blade, t o the extreme case o f a p r o p e l l e r which i s completely wrapped around i t s e l f . For p r a c t i c a l sweep angles, which are r e l a t i v e l y small, there i s a s l i g h t noise reduction. The e f f e c t o f sweep o f t h i s magnitude on performance has n o t been evaluated.

I n order t o t e s t some o f the concepts which were developed during the parametric studies, model p r o p e l l e r s were constructed f o r t e s t i n g i n the MIT wind tunnel. Figure 10 shows two model propellers, a "quiet" p r o p e l l e r and a standard Cessna 172 p r o p e l l e r . A1 though n o t obvious due t o the angle a t which the photograph was taken, the modified p r o p e l l e r has the same I t has a wider cord which was designed diameter as the standard p r o p e l l e r .

t o move the load d i s t r i b u t i o n inboard on t h e p r o p e l l e r blade.

These p r o p e l l e r s were t e s t e d over a wide range o f conditions on a p r o p e l i e r spinning r i g w i t h and w i t h o u t an after-body t o simulate an a i r c r a f t Figure 11 shows the t e s t c o n f i g u r a t i o n i n the MIT acoustic wind fuselage.

tunnel w i t h a fuselage afterbody. Figure 12 shows a sample comparison of measured and predicted noise data. A schematic o f the runnel c o n f i g u r a t i o n i s shown on the r i g h t p a r t o f tiie f i g u r e . The data are f o r the standard Cessna p r o p e l l e r model w i t h no afterbody. Noise data were measured w i t h a microphone mounted i n the ~ i r s t r e a m 1 diameter from the center of p r o p e l l e r r o t a t i o n . The data presented i s a pressure time h i s t o r y f o r approximately 2 r e v o l u t i o n s o f the p r o p e l l e r . These data correspond t o c r u i s e conditions f o r an actual a i r c r a f t . As can be seen, the agreement between the p r e d i c t e d and measured noise i s e x c e l l e n t . S i m i l a r r e s u l t s were obtained f o r other configurations.

A f t e r demonstrating the noice p r e d i c t i o n techniques i n the wind tunnel, f u l l - s c a l e p r o p e l l e r s were designed f o r f l i g h t t e s t s a t both Massachusetts I n s t i t u t e of Technoiogy and Ohio State U n i v e r s i t y . Figure 13 shows the f l i g h t I t i s a Beech Sundowner a i r c r z f t t e s t a i r c r a f t which w i l l be used by the OSU.

and i s equipped w i t h a microphone born which can be extended t o measure noise i n and behind the plane o f the p r o p e l l e r . Ground noise measurements w i l l be made f o r 500-feet and 1000-feet flyovers. Noise measurements w i l l be made w i t h and without an engine exhaust m u f f l e r t o deternine the r e l a t i v e l e v e l s A s i m i l a r f l i g h t demonstration w i l l be o f p r o p e l l e r and exhaust noise.

conducted by M I T using a Cessna 172 a i r c r a f t ; howcver, the HIT a i r c r a f t w i l l n o t be equipped f o r nearq-field i n f l i g h t noise me, urements and w i l l n o t have an engine exhaust m u f f l e r .

A f i n a l purpose o f t h i s p r 9 r a m i s t o e s t a b l i s h a center f o r e f f e c t i v e d i s t r i b u t i o n o f p r o p e l l e r optimization technology. Because of i t s c u r r e n t involvement i n the NASA program and i t s ready access t o a i r c r a f t manufacturers, the Ohio State University A i r f o i l Design and Analysis Center has been chosen t o serve t h i s function.

I n t e r i o r Noise Reduction I n t e r i o r Noise Reduction involves a l t e r i n g the c h a r a c t e r i s t i c s of the swnd path from the soorce t o the observer, as well as a l t e r i n g the charac- t e r i s t i c s of the noise source i t s e l f . The major elements o f the i n t e r i o r noise reduction program are l i s t e d i n Figure 14. The d e f i n i t i o n o f the i n p u t o r source f o r transmission studies i s obviously important. The under- standing o f s i dewall noise transmission mechanisms and the evaluation o f p o t e n t i a l noise c o n t r o l treatment are dlso key elements of the program.

Structureborne noise i s also o f i n t e r e s t due t o the p r o b l e m encountered by small p i s t o n engine a i r c r a f t . One source o f i n t e r i o r noise i n l i g h t a i r c r a f t i s vibration. This o r i g i n a t e s i n the engine, i s transmitted through the support structure, and i s radiated i n t o the cabin.

An example of structureborne noise research i s shown i n Figure 15. The research was d i r e c t e d toward determining tk r e l a t i v e magnitudes o f structure- borne noise and noise from other sources such as the p r o p e l l e r which might be transmitted through the a i r and through the fuselage sidewall i n t o the a i r c r a f t cabin. The p r i n c i p l e feature of the setup shown i s the use ~f stanchions located a t the f i r e w a l l on each side o f the a i r c r a f t t o support the engine weight and t h r u s t loads so t h a t the engine can be operated without any mechanical attachment t o the fuselage. The fuselage i s located i n the c o r r e c t geometry r e l a t i v e t o the engine so t h a t other noise sources are the same. The engine can also be attached t o the fuselage i n a normal configu- The engine attached configuration provides the t o t a l i n t e r i o r noise r a t i o n .

from a l l sources and paths wnile the engine detached configuration provides a l l sources except the stnictureborne noise, so the difference provides the structureborne contribution.

The bar chart a t the r i g h t o f the f i g u r e indicates t y p i c a l r e s u l t s . The t o t a l bar height indicates the t o t a l i n t e r i o r noise as measured i n the engine The shaded p o r t i o n o f the bar indicates attached condition. the structure- borne contribution. As indicated by the o v e r a l l l e v e l bar a t the r i g h t o f the figure, detaching the engine reduced the l e v e l by 3 dB, i n d i c a t i n g t h a t the structureborne c o n t r i b u t i o n i s about equal t o the c o n t r i b u t i o n from a l l Examination o f the spectrum indicates t h a t other source/path sombinatjons.

the structureborne c o n f r i b u t i o n i s s i g n i f i c a n t over a r e l a t i v e l y wide frequency range, up t o about 2000 !iertz.

Current research e f f o r t s are directed toward p r e d i c t i o n o f the s t r u c t u r a l l y transmitted noise and development o f noise control methods i n v o l v i n g c o n t r o l o f both noise radiated from panels t o the a i r c r a f t i n t e r i o r as well as noise transmitted through the engine mounting v i b r a t i o n i s o l a t o r s .

Fuselage sidewall transmission i s very important f o r those a i r c r a f t which have wing-mounted propel l e r s operating close t o the fuselage sidewall.

Research i s c u r r e n t l y underway t o evaluate p o s s i b l e s t r u c t u r a l t r e d t m n t s t o impr )ve the fuselage sidewall noise attenuation. An a i r c r a f t used i n one such study i s shown i n the upper l e f t photograph on Figure 16. This a i r c r a f t i s an Aero Commander 680, modified f o r e v a l u a t i n g i n t e r i o r noise c o n t r o l treatments. To provide a baseline f o r s t r u c t u r a l modifications being investigated, the i n t e r i o r t r i m and i n s u l a t i o n were removed i n t h e area o f i n t e r e s t and the windows were replaced w i t h s t i f f e n e d aluminum panels s i m i l a r t o the fuselage construction. These m o d i f i c a t i o n s are shown i n the lower r i g h t photograph o f the a i r c r a f t i n t e r i o r . The area i n v e s t i g a t e d i s s h m as the shaded area i n the sketch i n the lower l e f t o f t h e f i g u r e .

Sidewall noise a t t e n u a t i o n c h a r a c t e r i s t i c s were measured f o r p r o p e l l e r noise inputs and f o r a r t i f i c i a l noise inputs from the l a r g e horn shown i n the photograph. Attenuation provided by the sidewall f o r the horn i n p u t i s shown as noise reduction i n the upper r i g h t o f the f i g u r e . Noise r e d u c t i o n i s the d i f f e r e n c e between the i n s i d e and o u t s i d e sound l e v e l s as a function o f t h e frequency. The two curves shown are f o r the bare s i d e w a l l and f o r t h e side- w a l l w i t h 15 pounds of asphalt type, glue on mass added t o the a i r c r a f t .

These r e s u l t s i n d i c a t e t h a t even a modest amount o f a p p r o p r i a t e l y added mass may reduce i n t e r i o r noise by 4-15 dB depending on the frequency o f noise.

Conventional treatment w i l l n o t be s u f f i c i e n t f o r the new generation o f high-speed p r o p e l l e r driven a i r c r a f t .

Figure 17 shows the r e l a t i o n s h i p between desired cabin noise l e v e l s and c u r r e n t l y p r e d i c t e d pro^ e l l e r noise l e v e l s f o r c u r r e n t designs o f high-speed p r o p e l l e r s . The bar on the l e f t i n d i c a t e s the range of noise l e v e l s experienced i n t e s t i n g o f s u r r e n t pro- pel l e r s and the p r o j e c t e d improvement due t o advanced p r o p e l l e r design.

These l e v e l s are on the order o f 140 dB w i t h possible improverlents below t h a t .

Predicted i n t e r i o r noise l e v e l s and the i n t e r i o r noise goal a r e shown by the bars on the r i g h t . As can be seen, there i s a gap of approximately 25 dB i n the cabin a t t e n u a t i o n which must be obtained from n2w technology.

This problem i s being addressed i n two ways. F i r s t there a r e c o n t i n u i n g e f f o r t s t o reduce the noise o f high-speed p r o p e l l e r s througn c a r e f u l desiqn o f advanced configurations. I n a d d i t i o n , improved estimates o f p r o p e l l e r noise w i l l be obtained i n the sumner o f 1986 when a propfan model i s flown on a J e t Star a i r c r a f t .

A n a l y t i c a l studies are a l s o being pl~rsued t o d e f i n e low weight, low-noise- t r a n s m i t t i n g sidewalls. Pre! iminary r e s u l t s from two studies a r e shown i n Figure 18. The primary conclusion o f t h i s study i s t h a t acceptable cabin i n t e r i o r noise l e v e l s can be achieved using conventional technology. Both studies employed a double-wall design using an optimum combination o f added mass, s t r u c t u r a l dampicg, and t u n i n g o f the s t r u c t u r e . These s t u d i e s estimated the acoustic weight p e n a l t i e s which would accrue f o r the types o f a i r c r a f t l i s t e d i n Fsgure 18. The weights l i s t e d are p e n a l t i e s i n a d d i t i o n t o the acoustic treatment weights c u r r e n t l y c a r r i e d . A1 though these weights are high, the ~ o t e n t i a l o f the propfan as a f u e l e f f i c i e n t propulsion system i s s t i l l v i a b l e .

Future Research Trends of future NASA research are shown i n Figure 19. There w i l l be a continued e f f o r t i n the development and refinement o f noise prediction methods As these methods mture, simp1 i fied design techniques w i 11 be developed t o permit t h e i r practical application. The emphasis o f prediction and design technology w i l l s h i f t t o twin and c m u t e r size a i r c r a f t t o r e f l e c t t h e i r growing importance. I n t e r i o r noise research w i l l continue for a1 1 classes o f a i r c r a f t w i t h a special emphasis on developing the technology necessary f o r the timely development o f high-speed propeller-driven a i r c r a f t .

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FUTURE RESEARCH CONTINUED OEVELOPMEM AND REFI-T OF PREDICTION MOHOM EMPHAS I S TO INCLUDE TWIN AND COMMUTER SIZE A IRCRAR CONTINUING RESEARCH IN INTERIOR NOISE REDllCTlOFl F i r e 19 PROPELLER DYNAMIC AND AEROELASTIC EFFECTS* Barnes W . McComrick The Pennsylvania State University Various a s p e c t s of p r o p e l l e r blade dynamics are considered including those f a c t o r s which are e x c i t i n g t h e blades and t h e dynamic response of t h e blades t o t h e e x c i t a t i o n s . lCethods f o r t r e a t i n g t h i s dynamic system are described and problens discussed which rag arise w i t h advanced turboprop designs employing t h i n , swept blades.

INTRODUCTION A p r o p e l l e r on a s h a f t driven by an engine attached t o an a i r f r a r e repre- s e n t s a dynamic system. This system responds t o excitations from t h e power p l a n t as well a s t o unsteady aerodynamic f o r c e s on t h e blades. These unsteady f o r c e s r e s u l t from t h e non-uniform inflow i n t o t h e p r o p e l l e r as produced by an angle of a t t a c k o r by i n t e r f e r e n c e from t h e fuselage, wing and nacelles. They can a l s o be caused by a e r o e h s t i c phenomena such as classical o r s t a l l f l u t t e r .

The propeller-engine-airframe conbination, a s a continuous system, has an i n f i n i t e number of degrees of freedom and hence, an i n f i n i t e number of normal modes. Generally, a disturbance w i l l excite a l l of t h e modes, but it is only a few of t h e lowest modes which a r e of importance.

For purposes of determining p r o p e l l e r blade v i b r a t i o n s one can treat t h e p r o p e l l e r a s an i s o l a t e d dynamic system e x c i t e d by t h e unsteady torque at t h e hub from t h e engine anJ by t h e unsteady a i r l o a d s d i s t r i b u t e d along t h e blade surfaces. The dynamic response of t h e p r o p e l l e r t o t h e s e e x c i t a t i o n s determines t h e v i b r a t o r y blade stress l e v e l s , a knowledge of which is essen- t i a l t o a s s u r i n g an acceptable f a t i g u e l i f e of t h e blades. Noise and fuselage v i b r a t i o n l e v e l s a r e a l s o dependent t o same degree on t h e p r o p e l l e r dynamics.

This presentation w i l l d i s c u s s b r i e f l y methods of c a l c u l a t i n g t h e dynamic behavior of a p r o p e l l e r and w i l l present some r e s u l t s obtained t o d a t e on a N A S A research grant t o The Pennsylvania S t a t e University. This

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supported under N A S A Grants NSG-1308 and NSG-3304 grant involves not only t h e dynamics of t h e propeller but t h e unsteady aero- dynamics as w e l l ; i n p a r t i c u l a r , t h e interference v i t h t h e fuselage, ving and racelles .

The area of propeller blade dynamics promises to become even mre imqortant i n t h e f u t u r e with increasing a p p l i c a t i o n of f u e l - e f f i c i e n t turbo- prop i n s t a l l a t i o n s designed t o c r u i s e at high Mach numbere, These propellers, employing innovations such as composite materials, t h i n blades and sweep, w i l l present challenges i n t h e i r design and analysis which are not found with current all-metal blades. The enviromaent f o r a propeller can be lore severe than f o r t h e compressor of a turbofan engine. As s h o w i n f i g u r e 1, at an angle of a t t a c k , t h e i n l e t duct serves t o r e d i r e c t t h e i n f l w i n t o t h e ccm- pressor blades; whereas, t h e propeller blades experfence an unsteady f l w %re s p e c i f i c a l l y , because of t h e angle of a t t a c k of t h e propeller's axis.

t h e s e c t i o n angle of a t t a c k at a given radius v a r i e s approximately sinwoi- d a l l y v i t h t h e azimuth angle v i t h an amplitude proportional t o t h e propeller angle of a t t a c k and t h e square of t h e advance ratio.

The brief discussion t o f o l l w of propeller dynamics is perhaps best summarized by reference t o f i g u r e 2.

I n t r e a t i n g propeller dynamics, one is concerned v i t h those f a c t o r s which are exciting t h e system and with t h e response of t h e systen t o t h e excitation. These e x c i t a t i o n s may not depend on t h e dynamic response of t h e propeller o r they may as i n t h e case of stall f l u t t e r . A n accurate c a l c u l a t i o n of t h e unsteady forces and propeller dynamic response is e s s e n t i a l t o assuring an adequate f a t i g u e l i f e f o r t h e propeller.

Such calculations promise t o become more challenging i n t h e f u t u r e as pro- p e l l e r s are operated a t high c r u i s e Mach numbers. The introduction of new materials and t h e departure from today's conventional planform and a i r f o i l shapes may a l s o give rise t o new problems associated with propeller dyuamics, UNSTEADY KlRCES In addition t o t h e e f f e c t of angle of a t t a c k , non-uniformities i n t h e inflow t o a propeller r e s u l t from v e l o c i t i e s induced by t h e presence of t h e fuselage, wing and/or nacelles. These can cause both t h e magnitude and direction of the velocity vector i n t h e plane of t h e propeller t o vary with position (ref. 1). A computer code has been developed t o predict t h i s non- uniform velocity f i e l d . A s shown i n figure 3, t h e wing is replaced by a s i n g l e horseshoe vortex having a span equal t o t h a t f o r a t r a i l i n g rolled-up vortex sheet f o r an e l l i p t i c spanwise loading d i s t r i b u t i o n . The fuselage and nacelles a r e panelled with sources being placed on each panel. The strengths of these sources a r e adjusted t o assure t h a t the velocity normal t o the surface vanishes. This boundary condition can be relaxed t o allow f o r cooling airflow through i n l e t s .

Figure 4 presents t h e calculated v a r i a t i o n of a x i a l v e l o c i t y f o r a t y p i c a l single-engine l i g h t a i r p l a n e a s a function of azimuth position f o r 302 and 752 r a d i a l s t a t i o n s . The inboard section of t h e propeller is seen t o experience a s i g n i f i c a n t v a r i a t i o n i n the inflow equal t o approximately 60% of the ac.vance velocity a s the propeller rotates. Also of i n t e r e s t is t h e r e s u l t shown on t h i s f i g u r e t h a t one need not model t h e complete fuselage in o r d e r t o o b t a i n an a c c u r a t e d e s c r i p t i o n of t h e p r o p e l l e r inflow. I n t h i s case, t h e cowling, closed a t t h e rear by a simple f a i r e d shape, r e s u l t s in a predicted inflow which is c l o s e t o t h a t obtained v i t h t h e complete fuselage.

These curves are n o t synsnetrical about a v a l u e o f 180° because t h e p r o p e l l e r is yawed r e l a t i v e t o t h e fuselage.

A s a result o f non-uniform inflow, t h e aerodynamic loads on a p r o p e l l e r blade can vary s i g n i f i c a n t l y with azimuth position. However, t h e unsteady aerodynamic l o a d s p a l e by comparison t o t h e unsteady torque of a p i s t o n engine. Figure 5 p r e s e n t s some unpublished measurements obtained r e c e n t l y on

a four-cylinder , horizontally-opposed engine o p e r a t i n g a t 1300 rpm . From t h e

f i g u r e it is obvious why one should avoid o p e r a t i n g an engine continuously a t a speed corresponding t o a normal p r o p e l l e r mode. The amplitude o f t h e unsteady torque is of t h e o r d e r of 300% of t h e average value.

For turboprop a p p l i c a t i o n s , t h e engine torque is e s s e n t i a l l y constant so, here, one is more concerned about t h e unsteady a i r l o a d s .

BLADE DYNAMICS The dynamic response of a continuous system t o an e x c i t a t i o n can be c a l c u l a t e d by a s o l u t i o n of t h e d i f f e r e n t i a l equations governing t h e system o r by a lumped-parameter method which approximates t h e continuous system by Both of t h e s e approaches are being t r i e d under d i s c r e t e masses and springs.

t h e research g r a n t previously mentioned.

Because of t h e complex p r o p e l l e r geometry and t h e n a t u r e of t h e e x c i t i n g f o r c e s , a closed-form s o l u t i o n f o r t h e equations of p r o p e l l e r blade motion is highly unlikely, i f n o t impossible. Instead, one r e s o r t s t o c l a s s i c a l energy methods t o determine t h e normal modes of t h e propeller. These normal modes can then be applied t o t h e method of generalized coordinates and f o r c e s t o o b t a i n t h e dynamic response. To accomplish t h e foregoing, one must r e s o r t t o t h e use of l a r g e computer codes.

A t y p i c a l p r o p e l l e r f o r a s i n g l e , piston-engine, l i g h t a i r p l a n e is shown i n f i g u r e 6. I n t h i s l a b o r a t o r y study, t h e p r o p e l l e r is clamped i n a uni- v e r s a l t e s t i n g machine. An electromagnetic shaker e x c i t e s t h e blade at t h e and a p i e z o e l e c t r i c accelerometer measures t h e blade response a t various t i p , p o i n t s on t h e blade surface. Restrained a t t h e hub by t h e l a r g e mass of t h e t e s t i n g machine, t h e blade responds a s i f i t is c a n t i l e v e r e d from t h e hub with no e l a s t i c coupling t o t h e o t h e r blade.

Figure 7 i l l u s t r a t e s a d i f f e r e n t t e s t s e t up f o r a shaker test. Here, supported on a s o f t rubber innertube, t h e p r o p e l l e r responds a s a f r e e beam.

The e l e c t r o n i c equipment is shown i n t h i s f i g u r e c o n s i s t i n g of a m p l i f i e r s , power supply, frequency generator, o s c i l l o s c o p e and a ubiquitous spectrum analyzer and averager. By sweeping t h e frequency and noting resonances, one A manu- can quickly determine t h e frequencies of t h e lower normal modes.

f a c t u r e r w i l l test each p r o p e l l e r model i n t h e manner of t h i s f i g u r e t o a s s u r e t h a t none of tile lower modes correspond with e x c i t i n g frequencies from t h e engine. Since half of t h e c y l i n d e r s of a four-cycle engine f i r e during each revolution, t h i s impulse frequency i n Hertz is given by t h e product of t h e rpm and t h e number of c y l i n d e r s divided by 120.

Generally, t h e v i b r a t o r y motion of a p r o p e l l e r blade w i l l c o n s i s t of a bending o u t of its plane of rotatioil coupled with a bending in t h e plane and a t o r s i o n a l displacement along t h e blade.

Based on energy methods and t h e concept of a transmission matrix, a computer code h a s been developed (ref. 2) vhich p r e d i c t s t h e normal modes f o r coupled bending-bending o r coupled t o r s i o n with or-t-of-plane bending. The modelling of t h e complete coupling of all t h r e e motions has not been accomplished thus f a r . However, s i n c e present p r o p e l l e r blades a r e very s t i f f t o r s i o n a l l y and in-plane, t h e lower modes of the coupled bending-bending and bending-torsion models have approximately the same frequencies vhich are determined prir.cipally by t h e r e l a t i v e l y s o f t out- of-plane bending s t i f f n e s s . Thus, t h e l a c k of a completely coupled numerical m d e l is not too r e s t r i c t i v e f o r t h e present. However, t h i s may not be t h e case f o r f u t u r e turboprop designs. For t h i s reason a lumped-parameter model is being developed which w i l l allow f o r complete coupling, sweep and, possibly, a n i s o t r o p i c materials. This model w i l l be discussed b r i e f l y later.

The Campbell diagram f o r t h e p r o p e l l e r i n t h e previous f i g u r e s is pre- sented in f i g u r e 8. Here, t h e fundamental e x c i t i n g frequencies and harmonics f o r a four-cylinder, horizontally-opposed engine a r e superimposed on t h e n a t u r a l frequencies of t h e f i r s t t h r e e normal modes. The p t e d i c t i o n s are based on t h e combined bending-bending model f o r t h e clamped hub. Observe t h a t t h e n a t u r a l frequencies increase with rpm due t o c e n t r i f u g a l s t i f f e n i n g .

Data p o i n t s f o r zero r p m a r e included i n t h e f i g u r e and agree f a i r l y w e l l with t h e predictions. In c r u i s e , t h i s p a r t i c u l a r p r o p e l l e r operates a t around 2500 rpm. A t t h i s r o t a t i o n a l speed, t h e fundamental e x c i t i n g frequency of t h e engine and its harmonics do not coincide with any of t h e n a t u r a l frequencies of t h e f i r s t three modes. It would not be w e l l t o o p e r a t e t h i s engine- p r o p e l l e r combination continuously a c approxinately 2200 r p m s i n c e t h e n a t u r a l frequency of t h e f i r s t mode of t h e p r o p e l l e r , e i t h e r clamped o r f r e e , c o i n c i d e s with t h e e x c i t i n g frequency a t t h i s r o t a t i o n a l speed.

STALL FLLTTER In a d d i t i o n t o responding t o an unsteady inflow o r engine torque, a p r o p e l l e r blade can experience t h e aeroelastFc phenomena of s t a l l f l u t t e r .

This f a c t is not a new one, but is mentioned i n t h e l i t e r a t u r e a s e a r l y as 1941 ( r e f . 3). Unlike c l a s s i c a l f l u t t ~ r which r e q u i r e s a combined bending and t o r s i o n motion together with a phase s h i f t between t h e aerodynamic f o r c e and t h e angle of a t t a c k , s t a l l f l u t t e r can occur a s a pure bending o r t o r s i o n a l o s c i l l a t i o n . Because of negative damping provided by aerodynamic l i f t and moment beyond t h e s t a l l , s t a l l f l u t t e r can occur at much lower speeds than would be predicted f o r c l a s s i c a l f l u t t e r . Figure 9 c l e a r l y i l l u s t r a t e s the aerodynamic mechanism which can s u s t a i n s t d l f l u t t e r of a pee o s c i l i d t a r y nature ( r e f . 4 ) . Three h y s t e r e s i s loops f o r t h e moment c o e f f i c i e n t a r e shown f o r a 0012 a i r f o i l o s c i l l a t i n g about mean a n g l e s of a t t a c k of 0, 1 2 and 24 degrees. For each loop t h e reduced frequency equals .I12 and t h e amplitude of ir equals 6 degrees. For a = O0 ?: 6O, t h e a i r f o i l is u n s t a l l e d s o t h a t t h e closed i n t e g r a l of CM over a r e p r e s e n t s work which must be done by t h e system t o s u s t a i n t h e o s c i l l a t i o n .

For a = 24O +- 6O, t h e a i r f o i l is completely s t a l l e d s o t h a t , again, t h e area within t h e counter- clockwise closed h y s t e r e s i s loop r e p r e s e n t s work which must be done by t h e system. For a = 12' + 6 O t h e s i t u a t i o n is d i f f e r e n t . Here t h e a r e a enclosed by the clockwise loop minus t h a t enclosed by t h e counter-clockwise loop is p o s i t i v e as a r e s u l t of t h e a i r f o i l operating i n and out of t h e s t a l l e d region. This net a r e a represents work being done on t h e a i r f o i l which represents negative damping; and hence, t h e p o s s i b i l i t y of a s e l f - s u s t a i n i n g o s c i l l a t i o n .

Some recent unpublished test d a t a obtained at NASA LeRC with a model p r o p e l l e r a r e shown i n f i g u r e 10. For a given blade angle, the s e c t i o n a n g l e s of a t t a c k increase a s t h e advance r a t i o decreases. Thus, decreasing J, a value is reached below which a portion of t h e blade is s t a l l e d r e s u l t i n g i n t h e inception of stall f l u t t e r as noted on t h e figure.

FUTURE STUDIES A lumped parameter model of an e l a s t i c p r o p e l l e r blade is being developed a s an a l t e r r - a t e t o t h e blade dynamics program described e a r l i e r . The lumped parameter model w i l l allow one t o c a l c u l a t e t h e dynamic response of t h e p r o p e l l e r t o e x c i t i n g forces and torques without determining t h e normal modes f i r s t . A simplified sketch t o i l l u s t r a t e t h e model f o r t h e clamped hub case is shown i n f i g u r e 11. Here t h e blade is divided i n t o f i n i t e elements. The inboard end of each element is attached by t h r e e orthogonal equivalent springs t o t h e adjacent element. The axes of two of t h e s e t o r s i o n a l s p r i n g s l i e along t h e p r i n c i p a l axes of t h e blade s e c t i o n s . Their s p r i n g c o n s t a n t s can be e a s i l y c a l c u l a t e d knowing t h e s e c t i o n modulus of e l a s t i c i t y , moment of i n e r t i a and element length. Each element is a l s o allowed t o r o t a t e along a locus through t h e shear c e n t e r s of t h e sections. The t h i r d equivalent t o r s i o n a l spring i n t h e r a d i a l d i r e c t i o n a t the shear c e n t e r allows f o r t h i s t o r s i o n a l motion.

I f t h e mass c e n t e r and e l a s t i c a x i s a r e not coincident, an a c c e l e r a t i o n transverse t o t h e blade produces an i n e r t i a l moment which tends t o t w i s t t h e blade. Also, t h e f a c t t h a t t h e blade is twisted and tends t o bend about the s e c t i o n major a x i s w i l l produce a coupling between blade bending and torsion.

A program has been developed, based on modified vortex theory, which p r e d i c t s t h e time-dependent blade loading given t h e v e l o c i t y vector f i e l d i n t h e p r o p e l l e r plane ( r e f . 5). I n order t o v a l i d a t e t h i s program and t o l e a r n more about t h e d e t a i l s of t h e flow through a p r o p e l l e r , an experimental f l i g h t t e s t program w i l l be conducted t o measure t h e unsteady v e l o c i t i e s immediately behind a p r o p e l l e r f o r d i f f e r e n t f l i g h t conditions. Figure 12 i l l u s t r a t e s a three-component, hot-film anemometer mounted on a t r a n s v e r s i n g mechanism which is supported on a t r u s s attached t o t h e f i r e w a l l . The probe is i s o l a t e d from fuselage v i b r a t i o n by a s o f t mounting. The system is designed so t h a t measurements can be taken around t h e azimuth a t varying r a d i a l locations and distances downstream of t h e propeller. The a i r p l a n e t o be used f o r t h i s experiment is a Piper Cherokee 180 having a fixed-pitch propeller. For these tests t h e standard r e a r s e a t is replaced by a s i n g l e s e a t on t h e l e f t s i d e with an instrument rack on t h e r i g h t side. The r i g h t front s e a t is removed and a b a t t e r y pack put i n its place.

I n addition t o the flow f i e l d measurements, it is planned t o measure the unsteady bending moment d i s t r i b u t i o n along the blade. Miniature s t r a i n gages w i l l be bonded along the blade and t h e output transmitted across t h e hub by means of an FM multiplexing system.

CONCLUDING REPIARKS The subject of propeller blade dynamics promises t o become more important i n t h e future. Advanced turboprop designs incorporating t h i n , suept e l a s t i c blades, and operating i n a non-uniform inflow environment, w i l l require sophisticated analyses of t h e i r a e r o e l a s t i c behavior. I n addition t o avoiding resonances a t multiples of t h e propeller r o t a t i o n a l speed, the designer may have t o be concerned with t h e p o s s i b i l i t y of s t a l l f l u t t e r .

REFERENCES

I. Jumper, S . J. : :;,\?cputer Predict ion of Three-Dimensional P o t e n t i a l Flow-

f i e l d s i n which Aircraft Propellers Operate. M.S. Thesis, The Pennsylvania S t a t e University, February 1980.

2. Martinovic, 2. N.: A Study of t h e Dynamic Behavior of Propeller Blades.

The Pennsylvania S t a t e University, March 1979.

M.S. Thesis, 3. Bollay, William and Brown, C. D.: Some Experimental Results on Wing Flutter. Jour. Aero. Sci., vol. 8, no. 8, June 1941, pp. 313-318.

4. Carta, F. 0. and Niebanck, C. F.: Prediction of Rotor I n s t a b i l i t y a t High Forward Speeds, Volume 111, S t a l l F l u t t e r . USAACZABS Tech.

Report 68-18C, February 1969.

5 . Aljabri, A. S.: Prediction of Propeller Performance and Loading i n Uniform and Nonuniform Flowfields. M.S. Thesis, The Pennsylvania S t a t e University, November 1978.

PROPELLER TURBOFAN CMPARISON TURBOFAN CHOKED E X I T A I R F R M ISOLATION I-UNI FORM FLOW VELOCITY CONST PROPELLER NACELLEIAI RFRME INTERFEEKE F I G U R E 1 8 OXRATING CDNDI JIONS GEOrETRY 8 MATERIAL FIGURE 2 42 7 T R A I L I N G VORTEX

SWLIFIED WING-FUSELAGE-NACEUE WRICAL mrn

FIGURE 3 PREDICTED A X I A L V E L O C i M RATIO AT PROPELLER PROP PADIUS FRACTION R/R = 0,75 (TOP) AZIRUTH ANGLE