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STUDIES OF AIRFRAME-PROPULSION-SYSTEM
INTEGRATION FOR MACH 6 CRUISE VEHICLES
by Frunk S. Kirkhum, Jumes M . Czcbbuge, Jr., WuZter A. VuhI,
and WiZZi~mj. SZVUZZ
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0330758 NASA TN'D-4128 STUDIES O F AIRFRAME -PROPULSION-SYSTEM INTEGRATION FOR MACH 6 CRUISE VEHICLES By Frank S. Kirkham, J a m e s M. Cubbage, Jr., Walter A. Vahl, a n d William J. Small Langley R e s e a r c h Center Langley Station, Hampton, Va.
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - CFSTl price $3.00 STUDIES OF AIRF€?AME-PROPULSION-SYSTEM INTEGRATION FOR MACH 6 CRUISE WICLES* By Frank S. Kirkham, James M. Cubbage, Jr., Walter A. Vahl, and William J. Small Langley Research Center SUMMARY A n exploratory, experimental, and analytic investigation of airframe- propulsion-system integration has been conducted at a Mach number of 6. A two- pod nacelle configuration, a four-pod nacelle configuration, and a two- dimensional nacelle configuration with and without boundary-layer diverters were tested at a Reynolds number sufficient to produce a turbulent boundary layer on the wing ahead of the nacelles.
These preliminary results indicate no particular advantage of pod-type nacelles over two-dimensional designs and that the best over-all performance is obtained when the nacelle expansion area is no larger than the minimum required to enclose the turboramjet engines. A potential for significantly improving the lift-drag ratio of a configuration by utilizing the exhaust from underex- panded nozzles is also shown.
INTRODUCTION One of the principal problems involved in the design of the hypersonic air-breathing aircreft is the efficient integration of the airframe and propul- sion system. The engine airflow requirements for cruise-type aircraft designed to operate in the Mach 6 to 8 speed range are such that the inlet can be placed between the wing surface and wing-leading-edge shocks to take advantage of the high pressure airflow beneath the wing (fig. 1). At this speed, the area between the wing and shock system is large enough to permit considerable lati- tude in the shaping and placement of the engine nacelles (ref. 1 ) . Some of the basic questions pertinent to this problem for Mach 6 cruise configurations are as follows: What is the most effective exit-to-inlet area ratio for an engine housing?
Is a two-dimensional engine housing more efficient than pod-type nacelles?
Can any jet effects present be used to advantage?
The present exploratory study attempts to provide first answers to these questions, using simplified analytic and experimental models.
~
*
Presented at the classified “Conference on Hypersonic Aircraft Tech- nology,” Ames Research Center, May 16-18, 1967, and published in NASA SP-148.
SYMBOLS A e nacelle exit area A i n a c e l l e i n l e t area t o t a l i n l e t area of a s p e c i f i c configuration *i,t incremental drag c o e f f i c i e n t , D r a g of wing with nacelles minus ACD Drag of wing without nacelles incremental l i f t coefficient, L i f t of wing with nacelles minus ACL L i f t of wing without nacelles l o c a l s k i n - f r i c t i o n c o e f f i c i e n t
c f
d i i n l e t diameter ( f i g . 10) height of boundary-layer d i v e r t e r f o r two-dimensional nacelle hd ( f i g . 13) h i height of i n l e t of two-dimensional nacelle ( f i g . 13) i n t e r n a l s p e c i f i c impulse, seconds ISP (L/D)m= maximum l i f t - d r a g r a t i o l i f t - d r a g r a t i o without j e t e f f e c t s (L/D)O r a t i o of l i f t - d r a g r a t i o with j e t e f f e c t s t o l i f t - d r a g r a t i o (L/D)/(L/D), without j e t e f f e c t s length of pod support s t r u t ( f i g . 10) Zstrut free-stream Mach number Ea, nozzle e x i t s t a t i c pressure Pe s t a t i c pressure under wing P1 l o c a l Reynolds number ' *Z
s , wing planform area
angle of a t t a c k , degrees U wing r e f l e x angle, degrees ( f i g . 18) E d i v e r t e r wedge angle, degrees ( f i g . 13) Rl3SULTS AND DISCUSSION Ehgine Nacelle Sizing Requirements AS a s t a r t i n g point, consider t h e s i z e engine and i n l e t required f o r A representative thrust-drag schedule f o r t h e cruising f l i g h t a t Mach 6.
acceleration t r a j e c t o r y of a Mach 6 cruise vehicle i s presented i n f i g u r e 2.
In-line subsonic combustion turboramjets a r e used with t h e t r a n s i t i o n from tur- bojet t o ramjet operation occurring a t approximately Mach 3 . I n determining t h e s i z e of t h e propulsion system required, t h e t u r b o j e t thrust must be adequate t o provide t h e minimum acceleration desired i n t h e pinch regions occurring e i t h e r during t h e subsonic climb which may be imposed due t o sonic-boom con- sideration or during t h e transonic acceleration. The ramjet t h r u s t and t h e i n l e t area required are usually t a i l o r e d t o t h e cruise conditions, provided t h a t s a t i s f a c t o r y performance can be obtained over t h e complete acceleration t r a j e c t o r y . Once t h e t h r u s t requirements and hence t h e engine s i z e have been determined, t h e minimum propulsion pod dimensions a r e established - t h a t i s , t h e pod i n l e t area and engine exhaust nozzle e x i t area.
A parametric v a r i a t i o n of t h e r a t i o of nozzle-exit s t a t i c pressure pe t o underwing s t a t i c pressure p1 i s given a s a function of nacelle expansion r a t i o Ae/Ai i n f i g u r e 3 . These curves a r e generally applicable f o r subsonic combustion ramjets and a r e constant over a range of a l t i t u d e provided t h a t t h e maximum duct i n t e r n a l pressure l i m i t i s not exceeded. Also shown a r e t h e r e s u l t s from an in-house mission-analysis computer program which s i z e s t h e engine f o r a p a r t i c u l a r s e t of vehicle aerodynamics. The aerodynamics used herein were obtained from t h e study r e s u l t s of t h e d i s t i n c t d e l t a wing and blended wing-body cruise configurations described i n reference 2. The s i z e s of both i n - l i n e and wrap-around turboramjet engines required a r e indicated by bars i n t h e f i g u r e f o r an a l t i t u d e of approximately 100 000 f t . The lengths of these bars i n d i c a t e t h e v a r i a t i o n obtained a s t h e a i r c r a f t configuration w a s changed from t h e d i s t i n c t wing-body concept with a wing loading of 76 l b / f t 2 (a, 7 O ) t o a blended wing-body concept with a wing loading of 42 l b / f t 2 The minimum nacelle expansion r a t i o required f o r t h e i n - l i n e engine (a = 50).
i s about 1.3. The wrap-around engine i s somewhat l a r g e r and requires a nacelle expansion r a t i o of about 1.7. The nozzle exhausts a r e underexpanded p p1 > 1
) ( e/
f o r both engine types and nacelle expansion r a t i o s on t h e order of 2.5 would be required t o achieve f u l l expansion.
A s shown i n f i g u r e 4, about a ?-percent increase i n i n t e r n a l s p e c i f i c impulse can be obtained by f u l l y expanding t h e nozzle exhausts. If t h e nozzle flow can be expanded i n t o t h e wing surface, t h i s increased engine performance might be obtained with no penalty i n aerodynamic drag. I f , however, t h e f r o n t a l area of t h e nacelle must be enlarged t o achieve f u l l expansion, t h e increase i n engine performance with increasing nacelle e x i t - t o - i n l e t area r a t i o m u s t be traded off against t h e attendant drag penalty of an enlarged nacelle.
I Aerodynamic Characteristics A wind-tunnel program t o examine t h e aerodynamic c h a r a c t e r i s t i c s of various nacelle configurations w a s i n i t i a t e d . The types of nacelles considered are shown i n figure 5 . T e s t s w e r e conducted i n t h e Langley 20-inch Mach 6 tunnel with a sharp leading-edge TO0 sweep d e l t a wing. ( A description of t h e wind tun- Flow-through nacelles with constant i n t e r n a l duct n e l i s given i n ref. 3 . ) areas w e r e used t o simulate a two-pod nacelle configuration, a four-pod nacelle a two-dimensional nacelle configuration. The t o t a l i n l e t configuration, and area w a s 1.8 percent of t h e wing planform area f o r a l l configurations. Neither t h e i n l e t compression surfaces nor t h e correct nozzle exhaust flow were simu-
l a t e d i n t h i s investigation. The pod external contour w a s parabolic with a 140
i n i t i a l angle a t t h e l i p . The distance between adjacent pod center l i n e s w a s diameters long f o r a l l pod configurations 2 i n l e t diameters and t h e pods were 5 t e s t e d . The pods were placed longitudinally so t h a t t h e outboard pods would be behind t h e wing shocks throughout t h e angle-of-attack range (Oo 5 - - a 5 8O).
It w a s determined experimentally ( r e f . 4) t h a t t h e boundary l a y e r on t h e wing i s f u l l y turbulent i n t h e hatched region shown i n t h e sketches. The nacelles were t e s t e d i n this region of f u l l y turbulent wing boundary layer.
The boundary l a y e r on t h e two-dimensional and four-pod engine nacelles i s believed t o be t r a n s i t i o n a l while a region of f u l l y turbulent flow probably existed on t h e rearmost portion of t h e two-pod engine nacelles.
The importance of t e s t i n g with turbulent boundary layers i s i l l u s t r a t e d i n f i g u r e 6. The r e l a t i v e drag p e n a l t i e s of two-dimensional nacelles and pod nacelles a r e shown f o r both lamjnar and turbulent wing boundary layers. T e s t s conducted i n t h e Langley 11-inch hypersonic tunnel a t Mach 6.8 where t h e wing boundary l a y e r was laminar i n d i c a t e t h a t t h e pod-nacelle i n s t a l l a t i o n has 50 percent more drag than t h e two-dimensional nacelle. E a r l i e r r e s u l t s of this kind were i n t e r p r e t e d as v i r t u a l l y ruling out t h e use of pods on hypersonic However, t e s t s a t Mach 6 with turbulent wing boundary layers vehicles ( r e f . 1).
a 20-percent drag penalty r e l a t i v e t o t h e two- show t h e pods t o have only dimensional i n s t a l l a t i o n and t h i s can probably be f u r t h e r reduced with refine- ments i n design. Thus, by t e s t i n g with t h e correct type of boundary layer, t h e pod nacelle i s restored t o a more competitive position r e l a t i v e t o t h e two- dimensional design. The Reynolds number f o r t h e Mach 6 t e s t s i s s t i l l only 3 percent of t h e f u l l - s c a l e f l i g h t Reynolds number and, thus, additional scale e f f e c t s may be expected.
The r e s u l t s obtained i n t h e wind-tunnel program a r e presented i n f i g u r e s 7 t o 16. For two-pod nacelles, t h e e f f e c t of increasing t h e nacelle expansion r a t i o i s shown i n figure 7. The oil-flow photograph shows t h a t a strong i n t e r - action occurs between t h e nacelle shocks and t h e wing surface which produces s i g n i f i c a n t interference forces on t h e wing and nacelles.
The increments i n l i f t and drag obtained when t h e nacelles a r e added t o t h e basic wing are shown as a function of a i n f i g u r e 7. These data were obtained with a six-component strain-gage balance. All data a r e corrected f o r nacelle i n t e r n a l drag and f o r nacelle base drag. The i n t e r n a l drag correction w a s obtained by calculating t h e i n t e r n a l skin f r i c t i o n with t h e assumption of a laminar boundary layer. The base pressure was corrected to free-stream static pressure by assuming that the base pressure coefficient in the wind-tunnel tests was equd to -1,/~~*.
The theoretical predictions, shown as dashed curves, consisted of an esti- mate of the external pressure and friction drag on the nacelles and their sup- port struts plus an estimate of the forces on the wing in the interference region. The pressure forces on the pod nacelles were calculated by the method of characteristics for axisymmetric rotational flow. Shock-expansion theory was used on the support struts. The drag of all leading edges was obtained from Newtonian theory with a maximum pressure coefficient of 1.2. The inter- ference pressures on the wing were obtained from the pressure distribution in the axisymmetric flow field in the plane of the wing. Skin friction on the nacelles was assumed laminar for calculation purposes and to follow the equa-
tion Cfpl = 0.625. The theory predicts the trends but not the magnitude of
the experimental data.
The increments in lift and drag were used with the drag polar obtained from wind-tunnel tests of the delta-wing hypersonic cruise vehicle configura- tion described in reference 5 to obtain the effects of engine nacelle modifica- tions on the maximum untrimmed lift-drag ratio for a practical cruise vehicle design. The results are shown in figure 8. Addition of two pods with an expansion ratio of 1.5 decreased the maximum from about 4 to 3.8 while L/D increasing the nacelle expansion ratio to 2 decreased maximum to L/D about 3.6.
The effect of changing the length of the pod support strut is shown in figure 9. At low angles of attack, the incremental lift ACL is increased as pods are pulled closer to the wing surface but the incremental drag is unchanged. The drag is affected at the higher angles because the contribution of normal force to drag becomes significant. When these increments are applied 10) only a slight benefit in to the delta-wing cruise configuration (fig.
( L / D ) , = is obtained by decreasing the strut length because ( L / D ) " occurs at about a = 8'. A higher performance configuration with (L/D)max occurring
at lower angles of attack (a 5 bo) would benefit more from short pod support
struts because additional lift could be obtained with no increase in drag.
Thus, pod nacelles should be placed as close to the wing surface as is practical and perhaps merged into the wing surface.
A four-pod nacelle configuration is compared with the two-pod nacelle con- figuration in figure 11. The theory, contrary to the data, predicts a larger This ACL for the four-pod configuration than for the two-pod configuration.
discrepancy is possibly a result of the nacelles being placed farther forward from the wing trailing edge (relative to the length of the nacelle) for the four-pod configuration than the nacelles of the two-pod configuration. Both force and pressure tests to determine the effects of varying longitudinal placement of pod-type nacelles are needed to see whether significant changes in the aerodynamic forces occur. The theory does, however, predict an increase in drag as the number of pods is increased as would be expected since the wetted area and leading-edge a r e a increase as the number of pods increases. The a r e shown i n f i g u r e 12.
e f f e c t s of these increments on (L/D)mm The r e s u l t s of t h e investigation of a two-dimensional nacelle with an e x i t - t o - i n l e t area r a t i o of 1.5 a r e shown i n f i g u r e s 13 t o 15. The nacelle w a s t e s t e d with and without boundary-layer diverters. Without d i v e r t e r s , t h e nacelle w a s mounted f l u s h with t h e wing surface and t h e wing boundary l a y e r w a s allowed t o flow through t h e nacelle. The lower surface of t h e nacelle was a circular-arc p r o f i l e with an i n i t i a l angle of 8 . 4 O . The s i d e p l a t e s were swept 7 5 O with sharp leading edges and a 5 ' wedge angle.
BoundaSy-layer d i v e r t e r s may be used t o avoid degrading engine performance due t o ingesting t h e wing boundary layer. To simulate boundary-layer d i v e r t e r s , w a s supported on two s t r u t s ( f i g . 1 3 ) . The upper t h e two-dimensional n a c e l l e surface of t h e n a c e l l e had a 5 O wedge angle which diverted t h e wing boundary l a y e r toward t h e wing surface. The center portion of t h e boundary l a y e r flowed a through constant a r e a duct between t h e wing and nacelle, whereas t h e outer portion w a s diverted toward t h e s i d e s of t h e nacelles by t h e wedges on t h e d i v e r t e r s . The d i v e r t e r height w a s sized such t h a t t h e e n t i r e wing boundary l a y e r i n t h e wind-tunnel model would be diverted away from t h e i n l e t .
O n a f l i g h t vehicle t h e wing boundary l a y e r would be r e l a t i v e l y 60 percent as t h i c k as i n t h e wind tunnel and t h e d i v e r t e r height could be correspondingly reduced.
The curved shocks produced by t h e d i v e r t e r s (shown i n t h e photograph) indicate t h a t there i s a complicated i n t e r a c t i o n between t h e wing, t h e d i v e r t e r , and t h e nacelle which may be i n p a r t caused by choking i n t h e d i v e r t e r duct and boundary-layer separation.
The l i f t and drag increments f o r t h i s configuration a r e shown i n f i g u r e 14.
The t h e o r e t i c a l prediction of t h e forces on t h e lower surface of t h e nacelle w a s obtained by using two-dimensional shock-expansion theory with a correction f o r edge e f f e c t s by t h e method of reference 6. The t h e o r e t i c a l predictions of t h e e f f e c t s of adding boundary-layer d i v e r t e r s were done by first finding t h e equivalent Mach number i n t h e turbulent wing boundary l a y e r at t h e forward f a c e of t h e two-dimensional i n l e t ( r e f . 7). The flow w a s assumed i n v i s c i d a f t of t h i s point and t h e pressures were calculated by shock-expansion theory using t h e equivalent boundary-layer Mach number as a s t a r t i n g point. The skin f r i c - t i o n i n t h e d i v e r t e r ducts w a s assumed turbulent and calculated by t h e T ' method described i n reference 8. The seemingly accurate predictions of l i f t and drag given i n f i g u r e 14 a r e f o r t u i t o u s since t h e a x i a l force w a s underpre- d i c t e d and t h e normal f o r c e overpredicted which tended t o compensate each other when l i f t and drag were calculated.
(L/D)max i s shown i n f i g u r e 15. The The e f f e c t of t h e s e increments on two-dimensional nacelle without boundary-layer d i v e r t e r s caused only a s m a l l l o s s i n i n s p i t e of t h e f a c t t h a t t h e lower surface of t h e nacelle (L/D)mx w a s contoured t o give an e x i t - t o - i n l e t area r a t i o of 1.5. When d i v e r t e r s a r e added, however, a s i g n i f i c a n t penalty i n (L/D)max i s incurred. T h i s penalty Increasing i s unchanged by reducing t h e d i v e r t e r height by about 30 percent.
t h e d i v e r t e r wedge angle from t o 100 reduced (L/D)max as expected.
A comparison of the various nacelle concepts i s shown i n figure 16.
The best aerodynamic performance w a s obtained with t h e two-dimensional nacelle without boundary-layer diverters.
The addition of boundary-layer diverters, however, decreased t h e performance of t h e two-dimensional nacelle t o below t h a t of t h e two-pod nacelle configuration. More carefully designed diverters would undoubtedly increase t h i s performance l e v e l but t h e drag penalty f o r pod nacelle i n s t a l l a t i o n s can a l s o probably be reduced by proper integration of t h e pods and a i r c r a f t . Some of t h e pertinent variables f o r integrating pod nacelles with the a i r c r a f t a t lower speeds (M = 3 ) a r e described i n references 9 t o 12.
Since additional work t o optimize both nacelle types i s needed, no clear-cut choice between two-dimensional nacelles and pod nacelles can be made a t this time.
The trade-off between t h e increase i n engine performance against t h e decrease i n aerodynamic performance as t h e nacelle expansion r a t i o i s increased can now be examined. The pertinent parameter (L/D)-(Isp) i s shown on t h e r i g h t of f i g u r e 16. Assuming expansion of t h e nozzle flow i n t o t h e wing surface and no increase i n external drag, t h e two-dimensional nacelle without boundary- i s layer d i v e r t e r s can obtain a 2- percent gain i n performance as Ae/Ai increased from 1.3 t o 2. This increase i s due solely t o increasing Isp and t h e degradation i n engine performance due t o boundary-layer ingestion has not This performance l e v e l would be reduced been included f o r t h i s configuration.
i f boundary-layer ingestion e f f e c t s were included. For t h e two-pod configura- t i o n , t h e increase i n drag with increasing expansion r a t i o more than counter- balances t h e improvement i n engine performance, and t h e best over-all perfor- Considering t h e s m a l l per- mance i s obtained with t h e low area r a t i o nacelle.
formance gains obtainable by a d d i t i o n a l nozzle expansion, it appears t h a t t h e nozzle e x i t s should not be enlarged beyond t h e minimum s i z e required by t h e engine. If t h e nozzle exit areas are thus r e s t r i c t e d , t h e nozzle exhaust flow w i l l be underexpanded a s w a s discussed i n conjunction with f i g u r e 3 . The flow from underexpanded nozzles may impinge on adjacent a i r c r a f t surfaces and pro- duce j e t interference forces on t h e a i r c r a f t . A preliminary estimate of the jet interference e f f e c t s on t h e blended wing-body configuration has been made and i s considered next.
J e t Interference IXfects The configuration shown i n f i g u r e 17 i s t h e blended wing-body concept.
Further description of t h i s configuration i s given i n reference 5 . The nacelle housing t h e engines i s 30 f t wide and t h e nozzle e x i t s a r e 40 f t upstream of t h e wing t r a i l i n g edge. If t h e nozzle e x i t pressure pe i s greater than t h e underwing s t a t i c pressure p1 t h e nozzle flow continues t o expand along t h e surface of t h e wing creating an interference pressure f i e l d and a r e s u l t i n g force on t h e wing. I n t h e flow model used t o obtain a preliminary estimate of these forces, t h e flow w a s assumed t o be two dimensional, t h e wing t o be f l a t , and t h e pressure pe t o be constant along t h e wing surface t o t h e point where t h e t r a i l i n g expansion wave s t r i k e s t h e wing.
An example j e t e f f e c t calculation u t i l i z i n g these assumptions i s i l l u s - t r a t e d i n f i g u r e 18 where t h e r a t i o L/D with j e t e f f e c t s t o L/D without j e t e f f e c t s i s p l o t t e d as a function of t h e wing r e f l e x angle E. Significant improvements i n L/D can be achieved throughout t h e range of static-pressure r a t i o s considered without reflexing t h e wing. Wing reflex has a small benefi- c i a l e f f e c t a t t h e higher pressures but i s detrimental a t lower pressures.
A discussion of t h e u t i l i z a t i o n of underexpanded exhausts from asymmetric nozzles i s given i n reference 13.
The e f f e c t of j e t interference on t h a t might be obtained with r e a l - L/D i s t i c engines w a s determined through a range of c r u i s e Mach numbers from 5 t o 8, a s shown i n f i g u r e s 19 and 20. I n f i g u r e 19, t h e static-pressure r a t i o s f o r both i n - l i n e and wraparound turboramjet engines, a s obtained by t h e methods discussed i n conjunction with f i g u r e 3 , a r e shown a s a function of cruise Mach number. Both engine types have underexpanded nozzle exhausts throughout this Mach number range. The L/D improvements obtained by u t i l i z i n g these exhaust overpressures range from 5 t o 13 percent f o r t h e i n - l i n e engines and from 3 t o 10 percent f o r t h e wraparound engines ( f i g . 20). Some of t h e implications of these interference forces on a i r c r a f t s t a b i l i t y a r e discussed i n reference 5 .
Although these L/D improvements were obtained from an idealized analyti- c a l flow model, it i s apparent that s i g n i f i c a n t improvement i n L/D can prob- ably be r e a l i z e d by proper u t i l i z a t i o n of t h e exhausts of underexpanded nozzles.
Because of t h e extreme complexity of t h e exhaust flow f i e l d , a more r e a l i s t i c assessment of these e f f e c t s must be obtained through experimental t e s t s .
CONCLUDING RESIARKS These exploratory studies of airframe-propulsion-system integration a t Mach 6 have revealed t h a t underexpanded nozzles, whose e x i t a r e a s - a r e no l a r g e r than t h e maximum area required by t h e engine, appear t o produce b e t t e r over-all performance than f u l l y expanded nozzles. Preliminary calculations indicate t h a t underexpanded nozzles exhausting well ahead of t h e wing t r a i l i n g edge have a s i g n i f i c a n t p o t e n t i a l f o r increasing t h e l i f t - d r a g r a t i o . These j e t exhaust e f f e c t s , however, require d e t a i l e d experimental v e r i f i c a t i o n .
I n regard t o nacelle type, these preliminary r e s u l t s indicated no particu- l a r advantage of two-dimensional designs over individual pods.
The important interference e f f e c t s of t h e nacelles on both l i f t and drag a t hypersonic speeds were i n general not predicted accurately by t h e simple analytic techniques currently i n use.
Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., May 16, 1967, 126-13-03-31-23.
REFERENCES 1 . Fetterman, David E.; McLellan, Charles H.; Jackson, L. Robert; Henry, Beverly Z., Jr.; and Henry, John R.: A Review of Hy-personic Cruise Vehicles. NASA T M x-1276, 1966.
2. Drake, Hubert M.; Gregory, Thomas J.; and Petersen, Richard H . : Hy-personic Technology Problems Identified in Mission Studies.
Conference on Hyper- sonic Aircraft Technology, NASA SP-148, 1967, pp. 1-19.
3. Sterrett, James R.; and Emery, James C.: Extension of Boundary-Layer- Separation Criteria to a Mach Number of 6.5 by Utilizing Flat Plates With Forward-Facing Steps. NASA TN D-618, 1960.
4. Sterrett, James R.; Morrisette, E . Leon; Whitehead, Allen H., Jr.; and
Hicks, Raymond M.: Transition Fixing for Hy-personic Flow. Conference on Hypersonic Aircraft Technology, NASA SP-148, 1967, pp. 203-222.
(Also available as NASA TN D-4129. ) 5. Penland, Jim A.; Edwards, Clyde L. W.; Witcofski, Robert D.; and Marcum, Don C., Jr.: Comparative Aerodynamic Study of Two Hypersonic Cruise Aircraft Configurations Derived From Trade-off Studies. Conference on Hy-personic Aircraft Technology, NASA SP-148, 1967, pp. 45-62. (Also available as NASA TM X-1436. ) 6 . Liepmann, H. W.; and Roshko, A.: Elements of Gasdynamics. John Wiley & Sons, Inc., c.1957.
7 . Pinckney, S. Z.: Semiempirical Method for Predicting Effects of Incident-
Reflecting Shocks on the Turbulent Boundary Layer. NASA TN D-3029, 1965.
8. Peterson, John B., Jr.: A Comparison of Experimental and Theoretical Results for the Compressible Turbulent-Boundary-Layer Skin Friction With Zero Pressure Gradient. NASA TN D-1795, 1963.
9. Nichols, Mark R.: Aerodynamics of Airframe-Engine Integration of Super- sonic Aircraft. NASA TN D-3390, 1966.
10. Swan, Walter C.: A Discussion of Selected Aerodynamic Problems on Integra- tion of Propulsion Systems With Airframe on Transport Aircraft. Aerody- namics of Power Plant Installation, Part 1, AGARDograph 103, Oct. 1965, pp. 23-68.
11. Robins, A. Warner; Morris, Ode11 A.; and Harris, Roy V., Jr.: Recent Research Results in the Aerodynamics of Supersonic Vehicles. J. Aircraft, vol. 3, no. 6, Nov.-Dee. 1966, pp. 573-577.
E. Landrum, Ehma Jean: Effect of Nacelle Orientation on the Aerodynamic Char- acteristics of an Arrow Wing-Body Configuration at Mach Number 2.03.
NASA TN D-3284, 1966.
13. Lewis, W. G. E.; Herd, R. J.; and Herbert, M. V.: Lift Characteristics of Asymmetric Exhaust Nozzles at High Flight Speeds. J. Roy. Aeron. SOC.
(Tech. Notes), vol. 70, no. 671, Nov. 1966, pp. 1036-1040.
I MACH 6 CRUISE CONFIGURATION L-28-71-19 Figure 1 TYPICAL THRUST -DRAG SCHEDULE IN-LINE TURBORAMJET
I
I I I I
I k 3 4 5 6 0
Mco Figure 2
I -
NOZZLE-EXIT STATIC-PRESSURE RATIOS AT M,=6 STATIC - PRESSURE WRAP-AROUND 'RATIO, pe/pI TURBORAMJET I - EXIT -TO- INLET AREA RATIO, Ae/Ai Figure 3 RAMJET SPECIFIC IMPULSE Mm=6 ANGLE OF ATTACK,a , dag INTERNAL SPECIFIC IN-LINE TURBORAMJET 0 - TURBORAMJET I ~. L- . _ _ I EXIT-TO-INLET AREA RATIO, Ae/Ai Figure 4 NACELLE CONFIGURATIONS I NVESTlGATED Mm.6; -- Ai*t - 0.01 8 S W 2- POD 4-POD TWO-DIMENSIONAL ~ CONFIGURATION CONFIGURATION NACELLE ~ 0 0 O O a O - - -
A A
Figure 5 COMPARISON OF NACELLE DRAG IN LAMINAR AND TURBULENT FLOW RELATIVE OF POD AND OD NACELLES 2-0 NACELLES 1.2 ACELLE I .o Figure 6 EFFECT OF A,/Ai 2-POD NACELLES ; M , = 6 ; Zstrut/di = 0.250 .012 THEORY.
.008 .008 r
Figure 7 L-2871-4
EFFECT OF Ae/Ai ON UNTRIMMED (b)
max b r u t
2- POD NACELLES; M , = 6 ; - di = 0.250
A A 3 = 2.0
e = 1 . 5 WITHOUT
Ai Ai NACELLES 4 . 0 3.5 3.0 Figure 8 0 4 8 0 4 8
EFFECT OF STRUT LENGTH ON UNTRIMMED (b)
max 2-POD NACELLES; M,=6; Ae/Ai =1.5 WITHOUT NACELLES
Lstrut 4.250
4.0 di 3 . 0 OT Figure 10 EFFECT OF NUMBER OF PODS M,= 6 ; A,/Ai=1.5 ; Zs+ru+/di=0.250 Figure 11 L-2871-8 EFFECT OF NUMBER OF PODS ON UNTRIMMED(b) max
M,=6; A,/Ai=1.5; 'strut -- - 0.250
WITHOUT 2 PODS 4 PODS NACELLES Figure 12 TWO-DIMENSIONAL NACELLE WITH FLOW-TH ROUGH DIVERTER Ae/Ai 2 1.5 Figure 13 L-2871 -10 EFFECT OF 2-DIMENSIONAL NACELLE ON L I F T AND DRAG M,=6 Ae/Ai =1.5 DATA THEORY TYPE DIVERTER 0 --- WITHOUT DIVERTER 8 ~ 5 " hd/hi ~ 0 . 2 9
--- &5"
hd/hi ~0.17 A - - - - -- - 8=IO" hd/hi ~ 0 . 2 9 .om - .016 r . 0 0 6 ' ACD
.ow -
.002.
I I I I 0 2 4 6 8 0, deg Figure 14
EFFECT OF 2-DIMENSIONAL NACELLES ON UNTRIMMED (k)
'" "ax M,=6; Ae/Ai=l.5 EXPERIMENT 0 THEORY NACELLE WITH DIVERTERS A r NACELLE WITHOUT WITHOUT 5" 5 O I O " NACELLE DIVERTER h 4.0 = 0 . 2 9 0.17 0 . 2 9 3.5 3 . 0 Figure 15
I
SUMMARY OF 2-DIMENSIONAL AND POD NACELLE PERFORMANCE Figure 16 I.
-- I
JET- I NT E R F ER E N CE FLOW M 0 D E L CRUISE CONFlGURAT!OL FLOW MODEL NORMAL FORCE7
-v
SHOCK 40ft SHOCK JET BOUNDARY
EXPANSION WAVE^
\
Figure 1-7
L
- - PI pe - +
1.20 3.0 PI
NOZZLE - EX IT STAT1C -PRESSURE RAT1 OS AT CRUISE
4- IN-LINE 3- TURBORAMJET STATIC - PRESSURE 2 - RATIO, pe/pl WRAP-AROUND TURBORAMJET
,LL- - . I I I
5 6 7 8 CRUISE MACH NUMBER Figure 1 9
! = IMPROVEMENT DUE TO ENGINE-EXHAUST- WING
D INTERFERENCE E
-
OPTIMUM --- 0" 6 5.09 7 4.92 1.15 8 4.65
/
TURBORAMJET
b/(k)o 1 . 1 0
1.05 TURBORAMJET
I I d L O O , L
6 7 8 CRUISE MACH NUMBER Figure 20 NASA-Langley, 1961 - 1 L-5579 “The aeronautical and space activities of the United States shall be
conducted so a.~ to contribute . . . to the expansion of human knowl-
edge of phenomena i n the atmosphere and space. The Administration sball provide for the widest pracjicable and appropriate dissemination of information concerning its activities and the resalts tbereof .” -NATIONAL AERONAUTICS AND SPACLI Am OF 1958
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