Conceptual study of hypersonic airbreathing missiles

Abstract

The purpose of this paper i s t o r e p o r t r e c e n t r e s u l t s of an in-house conceptual study t o e v a l u a t e t h e performance p o t e n t i a l and research needs of a i rbreathing hypersonic m i s s i l e s . An alkylated-borane (noncryogenic) f u e l e d , dual-mode, ramjet/scranijet propulsion system s t r u c t u r e d w i t h a Rene'41 i n l e t and a carbon-carbon combustor was assumed along w i t t l a Lockalloy heat sink f u s e l a g e s t r u c t u r e a n d b e r y llium wings and control s u r f a c e s . Performance f o r an air-launched b a s e l i n e m i s s i l e w i t h a 961 pound s t a gi n g weight containing a 100 pound payload i n d i c a t e d e x c e l l e n t long range c r u i s e , moderate a c c e l e r a t i o n and high maneuverability p o t e n t i a l . For example: a t Mach 6 c r u i s e conditions (103,000 f t . ) , c r u i s e range exceeded 700 n.m.; horizontal a c c e l e r a t i o n s of 9 g ' s a r e o b t a i n a b l e f o r Mach numbers froni 4 t o 7 a t 30,000 f e e t ; approximately 40 g thrust-equal-drag turns a r e achievable w i t h t h e b a s e l i n e concept a t f l i g h t dynamic pressures of 10,000 psf ( M = 6 , 40,000 f t . ) . ranges of t h e o r d e r of 2500 nautical miles f o r payloads of 300 pounds can be achieved with moderate s i z e n.,issile c a r r y weights (9000 l b s , ) . Aerodynami c heatina analvses i n d i c a t e t h a t unprotected heat-A s i z i n g study i n d i c a t e s t h a t Mach 6 c r u i s e t o a s s e s s t h e p o t e n t i a l of t h e technology being developed i n Langley's hypersonic R&T programs f o r a p p l i c a t i o n s t o a i r b r e a t h i n g m i s s i l e s and t o ident i f y p o t e n t i a l a r e a s f o r f u r t h e r research. The i n t e n t was t o d e f i n e a concept f o r an a i r -b r e a t hing hypersonic l i f t i n g m i s s i l e and t o determine p o t e n t i a l performance and s c a l i n g r e l a t i o n s h i p s . To s p e c i f y weight and deterinine t h e magnitude of t h e technological o b s t a c l e s , i t was necessary t o d e f i n e a t l e a s t one s t r u c t u r a l approach f o r eng i n e s , a i r f r a m e , and radomes, and t o a s s e s s ihermos t r u c t u r a l l i m i t a t i o n s . Ooeratino Mach numbers from 4 t o 9 were i n v e s t i g a t e d . Concepts w i t h ranqes from 20 t o 3500 nautical miles and payloads from 70 t o 500 pounds were s t u d i e d . Symbois A a r e a A' area a t shock p o s i t i o n C s p e c i f i c heat modulus of e l a s t i c i t y ultiiiiate t e n s i l e s t r e n g t h P ET F t u sink s t r u c t u r e s -w i t h i n t e r n a l i n s u l a t i o n a r e f e a s ib l e f o r ranges of several hundred miles. For ranges of several thousands of miles a multiwall r a d i a t i o n s h i e l d ( I n c o n e l f t i t a n i u m ) was s e l e c t e d f o r protect i o n of t h e i n t e r n a l l y i n s u l a t e d heat sink s t r u c t u r e . Research areas in propulsion, propulsion i n t eg r a t i o n , c o n f i...

Performance Potential and Research Needs of a Hypersonic, Airbreathing, Lifting Missile Concept

“…15. These results indicate, for example, that for Mach 6 cruise conditions (q-24 kPa), essentially sharp, swept (74 deg) leading edges of beryllium or advanced titanium could be used without protective coatings for almost indefinite flight time.…”

“…Additional acceleration/range information for a modified baseline missile which achieved maximum speeds approaching Mach 9 is presented in Ref. 15.…”

“…22 of Ref. 15) with a weight per unit surface area of 7.8 kg/m 2 . (TPS is not required for beryllium wings and control surfaces since beryllium has extremely high heat-sink capability and could also operate at radiation equilibrium temperatures at Mach 6 and 32 km altitude.)…”

“…20 of Ref. 15). A three-dimensional woven, carbon-carbon material 24 could be used for the remainder of the rectangular inlet and the combustor.…”

“…Such an engine could provide efficient operation with subsonic combustion at a high equivalence ratio for acceleration from staging at Mach 3 or 4 and convert to supersonic combustion between Mach 5 and 6 for further acceleration and/or cruise. To investigate this type of dualmode operation for the Langley scramjet, subsonic combustion cycle analyses were performed 15 at Mach 4 flight conditions assuming that the fuel was injected from the wall downstream of the normal shock. (The Langley engine has an experimentally measured inlet contraction ratio of 5.7 at Mach 4, and a combustor-exit to combustor-entrance effective area ratio of approximately 3.7.)…”