AIAA-88-2985 The Challenges and Opportunities of Supersonic Transport Propulsion Technology William C. Strack, NASA Lewis Research Center, Cleveland, OH; and Shelby J. Morris Jr., NASA Langley Research Center, Hampton, VA

AIAA/ASME/SAE/ASEE 24th JOINT PROPULSION CONFERENCE

July 11-13, 1988/Boston, Massachusetts

For permission to copy or republish, contact the Amerian Institute of Aeronautics and Astronautics 370 L'Enfant Promemade. S.W., Washington, D.C. 20024

THE CHALLENGES AND OPPORTUNITIES OF SUPERSONIC TRANSPORT PROPULSION TECHNOLOGY

William C. Strack' NASA Lewis Research Center

Cleveland, Ohio 44135

and

Shelby J . Morris, Jr." NASA Langley Research Center Hampton, Virginia 23665-5225

Abstract

The major challenges confronting the propul- sion community for civil supersonic transport applications are identified: high propulsion sys- tem efficiency at both supersonic and subsonic cruise conditions, low-cost fuel with adequate thermal stability at high temperatures, low noise cycles and exhaust systems, low emission combustion systems, and low drag installations. Both past progress and future opportunities are discussed in relation to perceived technology shortfalls for an economically successful airplane that satisfies environmental constraints.

Introduction

Although the Concorde ushered in the super- sonic transport ( S S T ) era, it has not been a commercial success for a variety of reasons. Nevertheless, for several years there has been a growing interest in the subject of efficient sus- tained supersonic cruise technology applied to civil transport aircraft. A second-generation SST is envisioned that flies three times as many pas- sengers nearly twice as far and considerably faster than the Concorde while achieving economic and environmental near-parity with comparable technol- ogy subsonic transports (Fig. I ) . The central issue is the level of technology that must be attained in order to realize this vision.

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Two independent NASA sponsored studies, known as the High-speed Civil Transport (HSCT) studies, are currently in progress that address a broad spectrum of technical, economic, environmental, and related issues. While many issues are still being investigated, others have been partially resolved. For example, cruise speed analyses have narrowed the original Mach 0 to 25 range to the Mach 2 to 5 region - thus excluding hypersonic flight with its attendant hydrogen-fueled scramjet technology requirements. ogy improvements are required in all discipline areas to approach economic viability. lows, the propulsion technology gaps are discussed in broad perspective in hopes of highlighting the major issues and perhaps stimulating new ideas or novel solutions.

It is also clear that large technol-

In what fol-

An overview of the major HSCT technical challenges as viewed by a propulsion analyst is depicted in Fig. 2. The Concorde consumes about three times as much fuel per seat-mile as equiva- lent technology (circa 1976) subsonic long-range airplanes. This is largely responsible for its uncompeti tive economics - twice the total operating -

*Supervisory Aerospace Engineer. "Aero-Space Technologist; Member, AIAA

This D l p r iJ dcrhrrd 8 work a1 Ihr U.S. Gorrrnmrni tnd Is no8 rubl-I Io WVrChl prom~llon tn the Un1t.d S w n .

1

cost (TOC) as similar technology subsonic trans- ports and much worse than that relative to contem- porary technology aircraft. Very large airframe and propulsive efficiency improvements will be required to alter this situation. In our quest for greater productivity through increased speed, we are also confronted with ever increasing difficul- ties arising from high ram temperature levels. The challenge is to utilize advanced materials to cope with the hiah temoeratures without incurrina exces- ~ ~~ ~~

sive weight-and cost penalties. inability of traditional low-cost fuels to provide

In additioi, the

adequate thermal stabi I i ty impedes the pursuit of higher speeds. Expensive JP-type fuels reach ther- mal stability limits at speeds near Mach 4, but low cost Jet A is limited to speeds somewhat greater than Mach 2 .

There are also several challenging environmen- tal issues. While the sonic boom problem is air- frame driven, the excessive airport noise levels are due to the very high takeoff exhaust velocities associated with supersonic engines. Engine exhaust oas emissions is another environmental issue

~~~~ ~~~~

requiring attention. report, each of these issues will be discussed in more detail, including a summary o f previous prog-

In the remainder of this

ress, current status,~and future research require- ments. Further information, especially regarding previous studies, is available in Ref. I .

Fuel Economy

Figure 3 summarizes prior progress made in reducing supersonic transport engine thrust spe- cific fuel consumption (TSFC). Results are normal- ized by the cruise TSFC of the 1971 U . S . engine that was first proposed for a U.S. supersonic transport. This afterburning turbojet (GE4) performed relatively well at supersonic cruise conditions. But its subsonic efficiency was very inferior to comparable high bypass ratio subsonic engines. To mitigate this mismatch between a fixed-cycle engine and varying mlssion require- ments, the nation embarked on a IO-year NASA spon- sored variable cycle engine (VCE) research program that achieved considerable progress during the 1970 's and earlv 1980's. Comoared to the 1971 GE4 afterburning tu;bojet, the hybothetical VCE engines defined in 1981 (which assumed technology levels beyond 1981) consumed IO percent less fuel at supersonic and transonic conditions, and 25 percent less at subsonic soeeds - reflectina the cvcle

~I ~~ ~~~ ~ ~ ~~~~~

changing feature of VCE's. A simulianeous'25 per- cent reduction in engine weight occurred malniy as a result of improved materials. Nevertheless these gains are insufficient by themselves to obtain good enough fuel economy to enable competi- tion with subsonic aircraft. The subsonic effi- ciency of the 1981 VCE engines, for example, is

s t i l l o n l y about one-hal f t h a t of t o d a y ' s h i g h bypass r a t i o turbofans.

The p r imary cause o f the Concorde's h i g h f u e l consumption i s t he dramat ic C a l l i n a i r p l a n e LID a t supersonic speeds - on t h e o r d e r of one-hal f t h a t of subsonic t r a n s p o r t s . Th is i s o n l y p a r t i a l l y o f f s e t by the t r e n d towards i n c r e a s i n g o v e r a l l engine e f f i c i e n c y w i t h f l i g h t speed as shown i n F i a . 4 . " I n s t a l l e d c r u i s e e f f i c i e n c v " shown here i n i l u d e s i n l e t and nozz le l osses , b u i n o t n a c e l l e d rag , and rep resen ts des ign p o i n t va lues . The lowes t curve rep resen ts c u r r e n t l y o p e r a t i o n a l power- p l a n t s . The m idd le curve i n d i c a t e s t h a t s i g n i f i c a n t imorovement i s o o s s i b l e w i t h t o d a v ' s a v a i l a b l e tech- ~ ~~

n o iogy f o r b o t h ' s u b s o n i c and 8upe;sonic regimes. The t o p band p r o j e c t s f u t u r e o p p o r t u n i t i e s based p r i n c i p a l l y on NASA c y c l e analyses. Several a l t e r - n a t i v e c v c i e conceo ts -a re reoresented. i n c l u d i n a v e r y adv inced VCE and t u r b i n e bypass engines ( l ower boundary). and r a d i c a l l y d i f f e r e n t concepts such as supersonic th rough f low tu rbo fans (upper boundary). These advanced technology concepts p o t e n t i a l l y extend the peak p r o p u l s i o n e f f i c i e n c y l e v e l s from Mach 2+ t o a t l e a s t Mach 4. Gains o f 40 pe rcen t or more over Concorde's Olympus a re p o s s i b l e .

Us ing a s imp le c r i t e r i o n such as des ign p o i n t e f f i c i e n c v i s i n s u f f i c i e n t t o DroDer lv convev the o v e r a l l i i p a c t of advanced technology: p l e , t h i s p l o t shows a r e l a t i v e l y modest 8 pe rcen t g a i n between 1987 technology V C E ' s and advanced

For exam-

VCE's ( lower l i n e of t o p band). Not shown, b u t v i t a l l y i m p o r t a n t , a r e even l a r g e r improvements i n c l i m b e f f i c i e n c v (due t o n o n a f t e r b u r n i n a o o e r a t i o n ) dnd engine .eight 'of aavanced VCE s i o '!ll.s- [ - a r e tne ccn'plere improrenent p o t e n t i a l . F i g 5 d i s p l a y s an example of a "goa l " VCE t h a t rep resen ts what oavoffs would accrue if r e v o l u t i o n a r v advances i n make;ials and s t r u c t u r e s technoloav a re

~ ~ 1 1 ~ ~ ~ ~~ ~ ~ ~ ~ ~~ ~~~~ ~~ ~~~

a:hie/ed T h i s p a r t ' c J l a r aes lgn w a s generatea oy General E l e < t r i c i n t n e i r recen t hASA-sponsoreo ROMS s tudy.2 I t assumes e s s e n t i a l l v uncooled near s t o i c h i o m e t r i c engine m a t e r i a l s coubled t o advanced aerodynamics and s t r u c t u r a 1 , d e s i g n techno log ies . Th is i m p l i e s e x t e n s i v e use of nonmeta l l i c8 and i n t e r m e t a l l i c m a t e r i a l s .

Two l e v e l s of technology a r e quoted here : ( 1 ) t h e uncooled s t o i c h i o m e t r i c goal l e v e l i s denoted by the r i g h t - h a n d va lues ( G E ROMS). and ( 2 ) a 600 'F c o o l e r l e v e l i s denoted by t h e l e f t - hand va lues (NASA e s t i m a t e ) . One- th i rd o f t h e 32 pe rcen t u l t i m a t e fue l r e d u c t i o n p o t e n t i a l i s due t o a 45 pe rcen t engine weight r e d u c t i o n r e l a - t i v e t o a h y p o t h e t i c a l 1984 technology readiness b a s e l i n e enaine. Whi le a c h i e v i n a uncooled s t o i - c h i o m e t r i c i echno loav i s c e r t a i n i v a v e r v lona- term goa l , t h e magnitude-of t h e payo f f - i s so i a r g e - t h a t p u r s u i t o f h i g h temperature, m i n i m a l l y coo led cores and advanced VCE components i s key t o s u b s t a n t i a l improvement i n supersonic f l i g h t e f f i c i e n c y

To o b t a i n even b e t t e r powerp lant performance than af forded by a p p l y i n g advanced technology t o t h e VCE, nove l h i g h r i s k concepts w i l l be r e q u i r e d . One p o t e n t i a l HSCT breakthrough i s t h e supersonic f a n concept ( F i g . 6). I n s t e a d o f u s i n g a l o n g and heavy i n l e t system t o d e c e l e r a t e t h e i n t a k e a i r f l o w to subsonic speeds r e q u i r e d by conven t iona l turbo- machinery, t he supersonic Can e f f i c i e n t l y processes a i r a t supersonic throughf low v e l o c i t i e s . The advantages i n c l u d e much lower i n l e t system weight ,

2

l i g h t e r fan ( l e s s stages r e q u i r e d for a g i ven pres- sure r a t i o ) , l e s s i n l e t cowl and boundary l a y e r b leed drag. b e t t e r i n l e t pressure recove ry , and b e t t e r match ing of bypass r a t i o v a r i a t i o n s t o f l i g h t speed. Of course, t h e r e are many unknowns

o p e r a t i n g c h a r a c t e r i s t i c s ? How can the core i n l e t l osses assoc ia ted w i t h unsteady, s w i r l i n g , super- son ic i n f l o w be c o n t r o l l e d - or i s an a c t fan con- f i g u r a t i o n a b e t t e r s o l u t i o n ? L i t t l e e f f o r t has been expended on t h i s concept t o da te , a l t hough NASA has i n i t i a t e d two concept f e a s i b i l i t y research e f f o r t s , i n c l u d i n g an i n i t i a l fan model t e s t a t NASA Lewis Research Center .

and cha l l enges . What a re such a f a n ' s low-speed W

Tie p o t e n t i a ' pa)of f of s.pe*sonlc tncc .qvc l%. f a n ( S S T F ) recnnology for 4 t y p i c a l SST a p p l l c i - t m has been analyzed by NASA.3 t r i b u t o r s i s t h e i n l e t s i z e and we iah t r e d u c t i o n t o

One o f t h e major con-

about one -ha l f t h a t o f a conven t iona l supersonic i n l e t . Th i s a l s o reduces t h e i n l e t b leed drag pen- a l t y about 70 pe rcen t . The i n s t a l l e d e f f i c i e n c y i s improved n e a r l y IO percen t r e l a t i v e t o a comparable technology conven t iona l t u rbo fan engine, t he pro- p u l s i o n system weight i s reduced about 25 pe rcen t , and toge the r these improvements would y i e l d approx- i m a t e l y a 22 pe rcen t r e d u c t i o n i n m i s s i o n f u e l for a Mach 3. 5500 nm HSCT ( F i g . 7 ) .

F igu re 8 d i s p l a y s t h e impact o f p o t e n t i a l f u t u r e technology advances on a i r p l a n e f u e l con- sumption - r e c o g n i z i n g t h a t t h e key t o v i a b l e HSCT economics i s m iss ion fue l c o s t l e v e l s approaching those f o r f u t u r e subsonic a i r p l a n e s . Ach iev ing 100 Dercent f u e l usaae o a r i t v w i t h t h e subsonic ~. ,~ ~ ~ ~~~~~~~ ~

c o m p e t i t i o n i s n o t n e c e s s a r j because o f t h e increased p r o d u c t i v i t y assoc ia ted w i t h the much h i g h e r speed HSCT's. However, i t i s impor tan t to a t l e a s t be i n the same neiahborhood. which the v Concorde and p rev ious SST s h d y a i r p l a n e s cannot achieve d e s p i t e t h e i r r e l a t i v e l y s h o r t range capa- b i l i t i e s . Even a c u r r e n t technology 4500 nm HSCT f a l l s f a r s h o r t o f a c h i e v i n g f u e l - p a r i t y . p o r a t i n g advanced p r o p u l s i o n technology w i t h a 3500 "F l e v e l co re coupled t o a supersonic through- flow fan y i e l d s a major g a i n towards f u e l - p a r i t y i n the Mach 2 t o 3 r e g i o n even i f the des ign range i s increased from 4500 t o 5500 nm. To achieve t r u e f u e l - p a r i t y and 6500-nm range would a l s o r e q u i r e s u b s t a n t i a l a i r f r a m e improvements - on the o r d e r o f 20 pe rcen t L I D improvement and 25 pe rcen t s t r u c - t u r a l we igh t r e d u c t i o n . f i r s t o r d e r r e s u l t s s u b j e c t to m o d i f i c a t i o n s as t h e on-going s t u d i e s evo lve (e .g . , t h e r e s u l t s i n F i g . 8 do n o t p r o p e r l y recogn ize an a i r p o r t no i se c o n s t r a i n t ) . Another u n c e r t a i n t y i s t h e p o s s i b l e i n t r o d u c t i o n of a v e r y advanced a l l - new subsonic a i r p l a n e . An es t ima te of t h a t p o s s i b i l i t y i s i nc luded he re t h a t has an 11 pe rcen t L I D improve- ment, a 15 pe rcen t s t r u c t u r a l we igh t improvement, and a 33 pe rcen t p r o p u l s i o n e f f i c i e n c y improvement. The conc lus ion t o be drawn from t h i s a n a l v s i s i s

I n c o r -

These a re p r e l i m i n a r y

t h a t t h e l a r g e HSCT fue l consumption impehiment can be overcome, b u t i t w i l l r e q u i r e v e r y l a r g e tech- no logy ga ins i n a l l d i s c i p l i n e s - p r o p u l s i o n , a i r - p lane aerodynamics, and a i r f rame s t r u c t u r e s

Mixed Compression Supersonic I n l e t s

Ach iev lng supersonic f l i g h t a t moderate Mach 2 Concorde speeds can be viewed as r e l a t i v e l y s t r a i g h t f o r w a r d t e c h n o l o g i c a l l y , a l t hough l a r g e d technology improvements a re needed even a t Mach 2

t o y i e l d an economica l l y successfu l a i r p l a n e . Pushing the c r u i s e speed s u b s t a n t i a l l y h i g h e r i s c e r t a i n l y d e s i r a b l e , b u t i n t roduces a s e r i e s of eve r - i nc reas ing t e c h n o l o g i c a l chal lenges - beyond the fue l economy o f j u s t t he engine by i t s e l f . One of these new cha l l enges i s i l l u s t r a t e d i n F i g . 9 . - Convent ional e x t e r n a l compression i n l e t s accompl ish a l l o f t h e i r d i f f u s i o n o u t s i d e o f the i n t a k e d u c t through severa l o b l i q u e shocks and a t e r m i n a l nor- mal shock l o c a t e d a t t he cowl l i p . Th is type o f i n l e t d e l i v e r s adequate performance and i s w e l l - hehavpd ( ? t a b l e ) under a l l t r a n s o o r t f l i a h t condi - .. .~ .. .~~~ ~ ~ ~~ ~

t i o n s up to about Mach 2 . t he performance of e x t e r n a l compression i n l e t s r a o i d l v d e t e r i o r a t e s because o f t h e excess ive cowl

Beyond Mach 2: though,

drag a i s o c i a t e d w i t h the i n c r e a s i n g cowl l i p angle and the i n a b i l i t y t o inc rease the number of o b l i q u e shocks due t o excess i ve i n l e t l e n g t h and weight p e n a l t i e s . F l i g h t beyond Mach 2, t h e r e f o r e , r e q u i r e s a mixed compression type i n l e t t h a t pe r - forms some o f t he d i f f u s i o n i n s i d e the i n t a k e d u c t through more o b l i q u e shocks and a normal shock near the t h r o a t . Th i s i n t roduces o t h e r problems - nota- b l y more boundary l a y e r b leed t o a v o i d adverse shock-boundary l a y e r i n t e r a c t i o n s ( separa t i on ) and i n l e t shock system i n s t a b i l i t y . much more complex i n l e t and c o n t r o l system. N e i t h e r t r a n s p o r t s nor f i g h t e r s have been f l o w n w i t h such i n l e t s , y e t t h e need for utmost p r o p u l s i o n e f f i - c i ency w i l l r e q u i r e i t fo r high-speed t r a n s p o r t s .

t o a phenomena known as i n l e t i n s t a b i l i t y or " u n s t a r t " a s i l l u s t r a t e d i n F i a . 10. Whenever a

The r e s u l t i s a

Mixed compression i n l e t s a re q u i t e s u s c e p t i b l e

d i s tu rbance t h a t reduces the i n l e t mass flow occu rs , t he i n t e r n a l shock system moves a b r u p t l y upstream and r e p o s i t i o n s i t s e l f comp le te l y o u t s i d e t h e i n t a k e d u c t . Th is causes an ab rup t and severe drop i n t h r u s t due t o lower recove ry and mass flow, and a l a r g e i nc rease i n d rag . The p r e c i p i t a t i n g d i s t u r b - ance cou ld be r e l a t i v e l y smal l such as encoun te r ing a s t r o n g gus t or r a p i d l y changing the angle-of- a t t a c k . If the i n i t i a l d i s tu rbance i s l a r g e ( e . g . , compressor s t a l l ) , t he t r a n s i e n t response can be v e r y severe - p o s s i b l y u n s t a r t i n g n e i g h b o r i n g i n l e t - engine systems, which would l i k e l y throw the a i r - p lane i n t o a v i o l e n t yaw and r o l l maneuver. To p reven t such unacceptable behav io r , some form of s t a b i l i t y c o n t r o l system i s needed.

...,

One i n l e t s t a b i l i t y improvement concept con- s i s t s o f a s e t of s e l f - a c t u a t i n g b leed va l ves l o c a t e d i n the i n l e t n a c e l l e ( F i g . 1 1 ) . These r a o i d resoonse r a t e oneumatic va i ves w i l l ODen i n response 'to t h e i nc rease i n d u c t p ressu re pkoduced by a t r a n s i e n t excu rs ion o f the i n l e t t e r m i n a l shock fo rward f r o m i t s s teady -s ta te p o s i t i o n . A S t he shock moves f o r w a r d i t exposes the s t a b i l i t y bvoass olenum to inc reased oressure and automat i - c i i l y a k t i v a t e s t h e b leed va l ves which then s p i l l i n l e t b leed a i r overboard. Th is increases t h e i n l e t mass flow and f o r c e s t h e shock rearward, and thereby r e e s t a b l i s h e s s t a b i l i t y . The va l ves c l o s e when the t r a n s i e n t d i s tu rbance subsides and the shock has r e t r e a t e d t o i t s s teady s t a t e p ~ s i t i o n . ~ - ~

A n exaer imenta l wind tunne l t e s t aroaram a t NASA L e w i s ' v e r i f i e d the f e a s i b i l i t y o f ' t h f s concept d u r i n g t h e m id 1 9 7 0 ' s . A f i v e - f o l d i nc rease i n s t a b i l i t y marg in was demonstrated u s i n q a YF-12 system s j m u l a t i o n . Whi le encouraging, - these i n i - t i a l t e s t s r e p r e s e n t j u s t a beg inn ing - n o t an e s t a b l i s h e d d a t a base. Considerable research l i e s ahead t o adequate ly address t h i s i m p o r t a n t i s s u e .

v

Exhaust Nozzle Performance

The exhaust nozz le for an HSCT must pe r fo rm w e l l a t t h r e e c r i t i c a l f l i a h t c o n d i t i o n s - t a k e o f f . subsonic c r u i s e , and supergonic c r u i s e . imen ta l model t e s t r e s u l t s shown i n F i g . 12 (NASA

The exper-

Lewis 0- by 6-Ft Supersonic Wind Tunnel) o f an e j e c t o r nozz le show t h a t , w h i l e good takeof f and supersonic c r u i s e performance was achieved, t he subsonic c r u i s e performance was d i s a p p o i n t i n g l y low due t o flow separa t i on over the i n l e t doors of t he e j e c t o r . Th i s s h o r t f a l l i s impor tan t because i t s i g n i f i c a n t l y increases the rese rve f u e l a l lowance r e q u i r e d t o reach an a l t e r n a t e a i r p o r t . For long- range HSCT's, t he amount o f rese rve fue l i s q u i t e l a r g e - equal i n magnitude t o the pay load we igh t . I n a d d i t i o n . i t i s c r i t i c a l t o o b t a i n h i a h nozz le per formance.at t he t r a n s o n i c t h r u s t minu; a i r p l a n e drag "p inch p o i n t ' ' t o a l l o w adequate a i r c r a f t a c c e l e r a t i o n

Transonic P ropu ls ion System Draq

J u s t as exhaust nozz le performance i s c r i t i c a l d u r i n g subsonic f l i g h t , so a l s o i s t he m i n i m i z a t i o n of t r a n s o n i c i n s t a l l a t i o n losses assoc ia ted w i t h i n l e t s and nozz les . The i n l e t problem a r i s e s from a major mismatch i n i n l e t a i r f l o w supply c a p a c i t y ( too much) compared t o the engine demand as d i s - p layed i n F i g . 1 3 . Th i s mismatch causes some o f the i n le t -p rocessed a i r t o be s p i l l e d overboard d u r i n g of f -des ign c o n d i t i o n s . Th is rep resen ts a l o s s c a l l e d s p i l l a g e d rag t h a t i s p r o p o r t i o n a l t o the amount o f s p i l l e d a i r f l o w . The l e f t - h a n d p l o t shows t h a t s o i l l a a e increases r a o i d l v as an a i r - p lane i s slowed down from i t s d e s i g n - c r u i s e speed c o n d i t i o n to t r a n s o n i c speeds, and t h a t t h i s pen- a l t y becomes worse as des ign c r u i s e speed inc reases . S i m i l a r l y , a supersonic exhaust nozz le i s c l osed down d u r i n a t r a n s o n i c f l i a h t t o o b t a i n maximum t h r u s t b u t - t h i s a l s o prod;ces an e x t e r n a l f low loss known as b o a t t a i l d rag ( F i g . 13). The nozz le boat- t a i l t r a n s o n i c d raq Dena l t v a l s o r i s e s r a p i d l y w i t h des ign c r u i s e speed. the sum of the t r a n s o n i c i n l e t s o i l l a a e draa. boat-

A t h j g h des ign Mach numbers, ~~~ ~ ~ , . ~~~~

t a i l d rag , and n a c e l l e d rag c o u l d equal or exceed the drag of the a i r f rame. F i n d i n g s o l u t i o n s t o these i n s t a l l a t i o n problems i s e s s e n t i a l t o achieve an acceptable a i r p l a n e des ign

The Hiqh-Speed Fuel I ssue

Convent ional low-cost J e t A f ue l cannot w i t h - s tand the h i a h temoeratures assoc ia ted w i t h f l i a h t speeds much i n excess o f about Mach 2 . j e c t e d to temperatures above approx ima te l y 250 "C ( t i m e dependent a l s o ) they t h e r m a l l y decompose and

I f sub--

form coke depos i t s t h a t c l o g f u e l supply components and fue l i n j e c t o r s . Consequent ly, a cha l l enge e x i s t s to extend the thermal s t a b i l i t y o f conven- t i o n a l j e t f ue l t o h i g h e r temperatures w i t h o u t i n c u r r i n g a s i g n i f i c a n t f ue l p r i c e i nc rease - e i t h e r i n t h e fue l manufacture or assoc ia ted w i t h s o e c i a l f u e l t r a n s o o r t a t i o n and h a n d l i n a r e a u i r e - ments such as wi th 'JP-7 and c ryogen ics ?Fig.' 14) Whi le the p r a c t i c a l use of hydrogen l i e s f a r i n t o the f u t u r e , l i q u i d methane remains an i n t r i g u i n g p o s s i b i l i t y due to i t s c u r r e n t low p r i c e and h i g h thermal s t a b i l i t y . Endothermic f u e l s o f f e r more heat s i n k c a p a c i t y , b u t a r e f r a u g h t w i t h o f f s e t t i n g p r a c t i c a l and economic p e n a l t i e s . U n c e r t a i n f u t u r e fue l p r i c e s and i n f r a s t r u c t u r e cos ts c l o u d t h e i s s u e of fue l s e l e c t i o n and, consequent ly , a i r p l a n e des ign speed as w e l l . I n t he i n t e r i m , the c u r r e n t HSCT

3

s t u d i e s a r e proceeding under the assumptions def ined i n F i g . 1 5 . The dark squares rep resen t the b a s e l i n e f u e l p r i c e s w h i l e t h e shaded bands rep resen t the range of p r i c e s t o be considered i n s e n s i t i v i t y analyses. g a l l o n s of J e t A (EGJA) t o account f o r v a r i a t i o n s i n energy c o n t e n t .

P r i c e s a r e quoted i n t e r m s o f e q u i v a l e n t

Takeof f Noise Reduct ion

The f i r s t genera t i on of h y p o t h e t i c a l S S T ' s o f the e a r l y 1970's used a f t e r b u r n i n g t u r b o j e t s and would have provoked the i r r i t a t i o n o f many people l i v i n g around major a i r p o r t s . j e t exhaust v e l o c i t i e s ( o v e r 4000 f t / s ) by ove r - s i z i n g the engines and t h r o t t l i n g back d u r i n g take- o f f reduces no ise somewhat, b u t i t a l s o increases a i r p l a n e s i z e t o o r a p i d l y t o be an e f f e c t i v e method f o r more than a f e w d e c i b e l s as shown i n F i g . 16. Each curve i n F i g . 1 6 rep resen ts va r ious amounts of engine o v e r s i z i n g f o r a f i x e d miss ion. b l e n o i s e r e d u c t i o n progress evolved d u r i n g the 1970 's and e a r l y 1980's through a combinat ion of v a r i a b l e c y c l e fea tu res and many noise suppress lon concepts e x p e r i m e n t a l l y t e s t e d . U n f o r t u n a t e l y , even t h i s progress i s i n s u f f i c i e n t t o m e e t c u r r e n t FAR 36 Stage I 1 1 requ i remen ts . s h o r t f a l l shown i n F i g . 16 i s f o r a Mach 2.4. 4000 nm range a i r p l a n e . If we s e l e c t even g r e a t e r speeds and ranges w i t h o u t s imultaneous a i r p l a n e weight reduc ing technology improvements, then the no ise s h o r t f a l l can become s i g n i f i c a n t l y worse as dep ic ted by the s e n s i t i v i t y curves o f F i g . 1 7 .

a q u i e t HSCT w i t h o u t excess ive noise r e d u c t l o n pen- a l t i e s . Some of the no ise r e d u c t i o n concepts i l l u s t r a t e d i n F ig . 18 have been exp lo red i n a x i - symmetric c o n f i g u r a t i o n s s u i t a b l e f o r f l i g h t speeds up t o Mach 2 . 5 . These concepts need database ex tens ions i n bo th ax isymmetr ic and two-dimensional nozz le c o n f i g u r a t i o n s . Other concepts have p r a c t i - c a l l y no database a t a l l and a r e q u i t e Specu la t i ve . For example, the concept o f enhancing exhaust j e t m l x i n g w i t h pneumatic o s c i l l a t o r s rep resen ts a v e r y s p e c u l a t i v e and t e c h n i c a l l y c h a l l e n g l n g s t r a t e g y . The remote augmented t h r u s t system concept guaran- tees low n o i s e w i t h i t s h l g h mass f l o w , low pres- su re ra t i o fan. But i t in t roduces d i f f e r e n t problems - n o t a b l y , how t o i n t e g r a t e the dep loyab le remote t a k e o f f fans i n t o the a i r f rame.

Reducing t h e i r h i g h

Considera-

The 6-dB n o i s e

Much research l i e s ahead I f we are t o achieve

Emissions Reduct ion

Prev ious a i r p o r t p o l l u t i o n concerns p r e c i p i - t a t e d a NASA emisslons r e d u c t i o n research oroaram .-... t h a t l e d t o severa l emlss lon c o n t r o l m e c h a n i s k i n c l u d i n g the development o f two-zone combustors. The conven t iona l s lnqle-zone combustors have t h e i r h i g h power e f f l c i e n c j compromised to o b t a i n good i g n i t i o n and s t a b i l i t y a t low engine power l e v e l s . The improved two-zone combustors u t i l i z e d a p i l o t stage o p t i m i z e d f o r i d l e c o n d i t i o n s and a main s tage o p t i m i z e d f o r c r u i s e power. This r e s u l t e d i n l eaner , wel l -mixed high-power combustion w i t h app rox ima te l y one-hal f as much NOX emissions assum- i n g the engine c y c l e remains unchanged ( F i g . 19). However, ou r con t inued quest f o r h ighe r o v e r a l l engine e f f i c i e n c y produces even h ighe r c y c l e tem- p e r a t u r e s which increases NOX produc t ion . Hence, the f i n a l engine designs of the VCE program, if b u i l t , would have produced about as much NOX as the decade-ea r l i e r technology Concorde engines. Today, we face the same dilemma - performance d r i v e n

designs w l l l i nc rease NOX w h i l e emissions d r i v e n designs w i l l reduce performance.

There a r e severa l nonexc lus ive approaches t o NOX r e d u c t i o n . Reducing a i r p l a n e weight o b v i o u s l y reduces emissions through reduced f u e l consumption and t h i s may be achieved v i a a i r f rame aerodynamic and s t r u c t u r a l e f f i c i e n c y improvements as w e l l as p r o p u l s i o n system improvement. A more d i r e c t and e f f e c t i v e s t r a t e g y i s t o a l t e r the engine cyc le parameters, e i t h e r by compromising the c y c l e p r e s - sure r a t i o and lo r t u r b i n e i n l e t temperature o r by the a d d i t i o n o f a heat-p ipe combustor p recoo le r ( o r o t h e r heat exchange mechanism) t o lower t h e compressor e x i t temperature. These approaches are based on an e m p i r i c a l database t h a t r e l a t e s NOX p r o d u c t i o n r a t e s t o an exponent ia l f u n c t i o n of com- bus to r i n l e t temperature and a l i n e a r f u n c t i o n of combustor o u t l e t temperature. The combustor p re- c o o l i n g approach ( F i g . 20) i s a new idea and ana ly - s e s a r e c u r r e n t l y underway t o assess i t s m e r i t .

More d r a s t i c approaches i n v o l v e changing the fundamental combustion process. This i n c l u d e s con- cepts t o avo ld n e a r - s t o i c h i o m e t r i c flame tempera- t u r e s ( F i g . 2 1 ) and reduced res idence t i m e s . The lean premixed p revapor i zed combustor approach avoids l a r g e r e c i r c u l a t i o n reg ions w i t h i n the p r i - mary combustion zone thereby reduc ing the res idence t ime c o n s i d e r a b l y , b u t compromising combustor Sta- b i l i t y and a l t i t u d e r e l i g h t r e l i a b i l l t y w h i l e i n t r o d u c l n g f l ashback tendencies. q u i c k quenchl lean bu rn concept appears q u i t e a t t r a c t i v e i f a homogeneous r i c h m i x t u r e can be achieved i n reasonably s h o r t leng ths . exper lments performed for ground-based appl i c a - t l o n s . w h i l e encouraaina. need t o be extended t o

v

The r i c h burn1

L i m i t e d

<.

f l i g h t a p p l i c a t i o n s . be f e a s i b l e for near- term s e r v i c e w h i l e o t h e r s w i l l

Some of these concepts might , -.

r e q u i r e cons ide rab le technology e f f o r t j u s t t o determine t h e i r p r a c t i c a l i t y . aoo rox ima te l v 113 o f c u r r e n t l e v e l s seem f e a s i b l e

NOX r e d u c t i o n s t o

for the n e a r l t e r m approaches and to 1 /10 f o r the f a r - t e r m approaches ( F i g . 22) .

AS the 2 l s t c e n t u r y approaches, i t i s becom- i n q i n c r e a s i n g l y c l e a r t h a t e f f l c i e n t supersonic c r u i s e f l i g h t - i ; w i t h i n our t e c h n o l o g i c a l reach - a l b e l t a v e r y l a r g e reach. Many c h a l l e n g l n g pro- p u l s i o n problems need to be r e s o l v e d be fo re a s t a t e o f t echno logy readiness i s achieved ( F i g . 2 3 ) . A ve ry l a r g e improvement i n p r o p u l s i o n system e f f i - c i ency i s needed bo th a t supersonic c r u i s e and sub- son ic c r u i s e c o n d i t i o n s . Toward t h a t end, severa l advanced englne concepts are be lng considered t h a t , t oge the r w i t h advanced d i s c i p l l n e and component techno log ies , promise a t l e a s t 40 pe rcen t b e t t e r e f f i c i e n c y than the Concorde engine. The quest f o r h ighe r p r o d u c t i v i t y through h ighe r speed i s a lso t hwar ted by the l a c k o f a conven t iona l , l ow-p r i ced f u e l t h a t i s t h e r m a l l y s t a b l e a t the h i g h e r temper- a t u r e s assoc ia ted w i t h f a s t e r f l i g h t . Extending J e t A type f u e l t o h i g h e r temperatures and the use o f methane are two p o s s l b l l i t i e s r e q u i r i n g f u r t h e r i n v e s t i g a t i o n . A i r p o r t no ise remains a tough c h a l - lenge because p r e v i o u s l y researched concepts fa1 I s h o r t of a c h i e v i n g FAR 36-111 no ise l e v e l s . Inno- v a t i v e s o l u t i o n s may be necessary t o reach accepta- b l y low n o i s e . emiss lons w i l l r e q u i r e c y c l e compromises and uncon- - v e n t i o n a l combustor approaches. Whi le the t e c h n i - c a l cha l l enges are indeed formidable, i t i s

S i m i l a r l y , ach iev ing low exhaust

reasonable to assume that the current shortfalls 3. Franciscus, L .C . , "The Supersonic Through-Flow in fuel economy and environmental acceptability Turbofan for High Mach Propulsion," AIAA Paper can be overcome through an aggressive propulsion 87-2050, July 1987. (NASA TM 100114.) research program. Achievement of the propulsion goals outlined herein would indeed revolutionize 4 . Sanders, B.W.. and Mitchell, G . A . . "Increasing the future world air transportation system. the Stable Operating Range of a Mach 2.5 Inlet,

'-' AIAA Paper 70-686, June 1970. (NASA TM References X-52799.)

I . Morris, J.M. Jr., and Strack, W.C. , Propulsfon 5 . Sanders, B.W. , and Mitchell, G.A., "Throat Bypass Bleed Systems for Increasing the Stable Air Flow Range o f a Mach 2.50 Axisymmetric Inlet with 40-Percent Internal Contraction."

System Issues for the High-speed Civil Trans- port Study, AIAA Paper 87-2938, Sept. 1987.

Materials and Structures Study," NASA 2. Feig. P.D., "Revolutionary Opportunities for NASA TM X-2779. 1973.

CR-179642. 1987.

MACH 2 SEED -ILICnss J?m N. MI. W E - E.W N. MI. 1W PASSENGERS SIZE * 3m + PASSENGERS

<%F&AHl WITH FAR 8 TICKET COS - ZERO OR SMALL SURCHARGE

M R U N D C R U S SONK WJM - E E R FAR b - AIRPan NOISE

4r SUBSONIC ->ma

Fig. 1 A vision of future high-speed transport

ECONOMICS

Fig. 2 Challenges to high-speed transports.

5

1 .O

0.9

RELATIVE WEIGHT

1 .oo v

RELATIVE

CONSUMPTION FUEL 0.8

0.7

0.6 0.75

Fig. 3 SST propuls ion p r o g r e s s

INSTALLED CRUISE EFFICIENCY

r l - i i

0 1 2 3 4 5 6 FLIGHT MACH NUMBER

01

Fig. 4 F u t u r e high-speed propuls ion p e r f o r m a n c e potent ial .

POTENTIAL MAC11 3 CRUISE CONDITIONS

DRUM-' U I V ~ 2 - 2 4 COMP OPR 9 - 16

UNCCOLED & A D D

nENEPlTS iMATERlA1.S & ACKO) 290 PAX 5000 NM TKANSPORT

llrlvlivr io Cuneni Technology @ 16D/gsl.

.-SINGLE STAGE SUPERSONIC FAN

SUPERSONlC ;3 STAGE CONVENTIONAL FAN :, :-DUCT NOZZLE DlFFVS

SUBSONIC SUPERSONIC OIFFUSER ---,

CONVENTIONAL TURBOFAN SUPERSONIC FAN ENGINE

SUPERSONIC FAN ENGINE FEATURES IMPLICATIONS .SHORT. ALL SUPERSONIC "ET . SlNGLE STAGE SUPERSONIC FAN . BPR OECREASES WlTH Mo

* LOWER WEIGHT. LOWER INLET DRAG .LOWER WElGHT AND COST. RUGGED SWDlNG . HlGHER CRUlSE THRUST

Fig. 6 Supersonic throughf low f a n enone

Fig. 5 V a r i a b l e cycle e n g i n e goal.

6

40 -

30

PERCENT IMPROVE.

20

1.0

0.0 TOTAL

PRESSURE RECOVERY

0.0

0.4

0.2

SUPERSONIC FAN

0

~

~ Shock

~ . Low p",orma"es abors UQCh 2 - Pro,,. '0Cw"y

- Cowl drag . Hlqhor Weigh1 Higher B.L. Blsad Droq

I I I I I I

Fig. 7 Benefit of supersonic throughflow fan. Mach 3 mrnrnercial transport, 300 passengers, 5500 n.m. range.

CRUISE MACH NUMBER

Fig. 8 Impact of technology on fuel economy. 300 passengers, unmnstrained noise.

Fig. 9 Supersonic inlet performance.

7

0.8 0.~1 PRESSURE I RECOVERY

0.6

O ' l 4 I 0.6 0.7 0.8 0.9 1 .o

MASS FLOW RATIO

Fg. IO Mixed compression supersonic inlet instability.

1.0,

0:5

Fig. 11

0.6 0.7 0.8 0.9 1 .o MASS FLOW RATIO

Mixed compression inlet stabilty improvement.

1.00 I

.98

GROSS .96 THRUST COEFF. .g1

.92

.30

.88

.86

c f

SUBSONIC CRUSE

MACH NO.

SCR GOALS

MODEL TFSTS

LEWIS SWT

Fig. 12 Supersonic mise nozzle performance. VCE research ejector.

8

INLET NOZZLE

FUEL PRICE $lEGJA

Fllohl Uwh Number C r u h Uaoh Numbar

Fig. 13 Transonic propulsion system drag.

3

mNVWTlONAL RELATIVE rnKAREm5

FUEL 2 PRICE

1

01,,,,,, 5 6 0 1 2 3 4

MACH NUMBER

Fig. 14 The high-speed transport fuels issue.

THERMAL STABILITY LIMIT,'F

Fig. 15 HSCT studies fuel assumptions for thermally s t a ~ e jet fuel (~Kanci methane (cH 4).

9

L E;& 4 0 0 7

100 105 110 115 120 125 SIDELINE NOISE, EPNdB @ 1500 It.

Fg 16 SST takeoff n o w reaxion progress Mach 24. 4000 n rn range, 300 passengers

SIDELINE NOISE

EXCEEDING FAR 36-111.

EPNdB

4.5 5.0 5 2.0 2.5 3.0 3.5 4.0

MACH NUMBER

Fig. 17 High-speed civil transport take-off noise challenge. Suppressed VCE turbofanlrarnjet engines, constant liftoff velocity, 1988 technology.

I ACOUSIIC LlNlNC

Fig. 18 Jet noise reduction concepts.

10

EMISSION INDEX

(g/kg FUEL) 20

lo I

INVARIANT EMISSION

r" FlRSl GENERATION SSl prl.nmr*

SINGLE TONE COMBUSTOR 7 '* EARLY 1 9 8 ~ s VCE STUDY ENGINE

g'\ @ 2-ZONE INVARIANT COMBUSTOR CYCLE

0 1 I I I 1970 1980 1990

YEAR

Fig. 19 History of SST cruise NO, emission reduction.

Haol Exshongar Cools A b Enlsrlng Cornbudei

0 REDUCES NOX BY 4X FOR 3% TSFC PENALTY

WEIGHT AND COMPLEXITY PENALTIES

Fig. 20 Combustor precooling approach to reducing NO,.

P I

EQUIVALENCE RATIO. 0

LEAN PREMIXED PREVAPORIZED RICH BURN/OulcK QUENCH/LEAN BURN I I

Fig. 21 Variation of NO, with equivalence ratio

11

ease

NOX EM IS S IO N S

BaEB 3 -

Fig. 22 Emissions technology impact

4 0 % I'UOI. SAVINGS . ENVIRONMIINTAL A C C E l T h i I l l . l T Y ECONOMlC,Al.lSThlll .B FUEL

THTRMAI.1.Y STABLE w r I ~ L

LOW

Fig. 23 High-speed propulsion technologies

12