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Declassified by a u t h o r i t y o f EASA 1 3 Change Notices N'o.-L---..
~ A R G E - S C A L E LOW-SPEED WIND-TUNNEL TESTS OF A DELTA WINGED SUPERSONIC TRANSPORT MODEL TO DETERMINE AERODYNAMIC EFFECTS OF FORWARD OR REVERSE THRUST By William H. Tolhurst and Kiyoshi Aoyagi Ames Research Center Moffett Field, Calif.
N O T I C E T h i s document should not be returned after i t has satisfied your requirements. It may be disposed of i n accordance with your local security regula- tions or the oppropriate provisions of the Industrial Security Monual for Safe-Guarding C l a s s i f i e d i s p r m i b i t e d by law. lnformat ion.
N A T I O N A L AERONAUTICS AND SPACE ADMINISTRATION m m m m m m m em m m m m m m m m m m m a m m m m m m m m e m m m m m m m m m m m m m m m m m m m m m m m m m m m m m m em LARGE-SCALE LOW-SPEED WIND-TUNNEL TESTS O F A DELTA WINGED SUPERSONIC TWSPORT MODEL TO .
DETERMINE AERODYNAMIC EFFECTS O F FORWARD OR FEVERSE THRUST* By W i l l i a m H. Tolhurst and Kiyoshi Aoyagi Ames Research Center Moffett F i e l d , C a l i f .
SUMMARY The purpose of t h e i n v e s t i g a t i o n was t o determine t h e aerodynamic e f f e c t s of t h e o p e r a t i o n of wing-pod-mounted j e t engines on t h e l o n g i t u d i n a l charac- t e r i s t i c s of a supersonic t r a n s p o r t model with a d e l t a wing of aspect r a t i o 2.17.
Data a r e p r e s e n t e d f o r various configuration combinations which i n c l u d e d wing t r a i l i n g - e d g e f l a p d e f l e c t i o n s from 0 ' t o 30°, h o r i z o n t a l - t a i l incidence angles from 0 ' t o -l5', and droop angles from 0 ' t o -25'. The a i r p l a n e angle-
o f - a t t a c k range extended from - 4 ' t o +lTo with a Reynolds number range from
17.2X106 t o 3 2 . 2 ~ 1 0 ~ . The d a t a include l o n g i t u d i n a l f o r c e and moment d a t a w i t h t h e engines i n b o t h forward and reverse t h r u s t and t h e m a x i m u m tempera- t u r e of t h e s u r f a c e of t h e h o r i z o n t a l tail.
The r e s u l t s i n d i c a t e t h a t w i t h t h e engines i n forward t h r u s t , t h e aerodynamic e f f e c t s of engine o p e r a t i o n on t h e l i f t and pitching-moment char- a c t e r i s t i c s of t h e a i r p l a n e were small compared t o t h e d i r e c t engine t h r u s t f o r c e s . With t h e engines i n r e v e r s e t h r u s t t h e e f f e c t s on t h e aerodynamic c h a r a c t e r i s t i c s were g r e a t e r t h a n t h e d i r e c t engine t h r u s t f o r c e s . F u l l r e v e r s e t h r u s t caused a reduction i n l o n g i t u d i n a l s t a b i l i t y when t h e wing t r a i l i n g - e d g e f l a p s were u n d e f l e c t e d and a l o s s i n s t a b i l i t y when t h e f l a p s were d e f l e c t e d t o 30'.
Under c o n d i t i o n s where t h e h o r i z o n t a l t a i l would be drooped i n t o t h e engine exhaust-gas flow f i e l d , i t s s k i n temperature may i n c r e a s e t o approxi- mately 35 p e r c e n t of t h e d i f f e r e n c e between t h e engine j e t - e x h a u s t temperature and f r e e - s t r e a m temperature.
INTRODUCTION S e v e r a l i n v e s t i g a t i o n s have been made a t t h e Ames 40- by 8 0 - ~ o o tWind Tunnel t o determine, a t l a r g e s c a l e , t h e low-speed aerodynamic c h a r a c t e r i s t i c s * T i t l e , U n c l a s s i f i e d 0 . 0 . . . 0.. . . . . 0.. 0 .
0 . e . 0 .
. . 0 . . 0 . .
.. 0 . . e .
0 . ..e . . .
of a delta-winged supersonic t r a n s p o r t a i r p l a i i e c o n f i g u r a t i o n without engines.
The r e s u l t s of t h e s e t e s t s a r e r e p o r t e d i n r e f e r e n c e s 1, 2, and 3. The h o r i - z o n t a l t a i l considered i n r e f e r e n c e 3 would be used f o r l o n g i t u d i n a l s t a b i l i t y and c o n t r o l during subsonic f l i g h t .
However, j u s t p r i o r t o supersonic f l i g h t t h e t a i l would be drooped t o provide a d d i t i o n a l d i r e c t i o n a l s t a b i l i t y at supersonic speed.
It would a l s o provide a means of reducing t h e e f f e c t of t h e aerodynamic c e n t e r v a r i a t i o n between subsonic and supersonic speeds and g i v e r e l a t i v e l y low t r i m drag throughout t h e speed range, as r e p o r t e d i n r e f e r e n c e 4 : “lie p r e s e n t t e s t s were made t o deterrine the effect of engine o p e r a t i o n on t h e s t a t i c l o n g i t u d i n a l s t a b i l i t y of t h e a i r p l a n e i n c o n f i g u r a t i o n s r e p r e - s e n t a t i v e of take-off and subsonic climb. T e s t s were a l s o made t o determine t h e e f f e c t of f u l l r e v e r s e t h r u s t on t h e s t a t i c l o n g i t u d i n a l c h a r a c t e r i s t i c s of t h e a i r p l a n e i n c o n f i g u r a t i o n s r e p r e s e n t a t i v e of letdown from a l t i t u d e a t subsonic speeds and during t h e l a n d i n g approach.
The l o n g i t u d i n a l f o r c e and moment d a t a of t h i s r e p o r t show t h e aerody- namic e f f e c t s of v a r i o u s v a l u e s of forward or r e v e r s e . t h r u s t on t h e model with s e v e r a l combinations of wing f l a p d e f l e c t i o n and h o r i z o n t a l - t a i l incidence and droop a n g l e s . A l s o p r e s e n t e d i s a b r i e f survey of t h e s k i n temperatures of t h e h o r i z o n t a l t a i l as it was drooped through t h e exhaust j e t .
NOTATION
engine i n l e t a r e a ( t o t a l of 4 engines - 5.58 f t 2 )
wing span, f t chord l e n g t h , f t
, q b I 2 c2 dy
mean aerodynamic chord
drag c o e f f i c i e n t , - b a g
qoos l i f t
l i f t c o e f f i c i e n t , -
qms p i t c h i n g moment pitching-moment c o e f f i c i e n t , QSC gross t h r u s t of f o u r engines, l b gross r e v e r s e t h r u s t of f o u r e n g i n e s , l b ‘ e ’ net t h r u s t of f o u r engines,
, l b
Fg - g
g r a v i t a t i o n a l a c c e l e r a t i o n , 32.2 f t / s e c *
-
h o r i z o n t a l - t a i l incidence, p o s i t i v e with t r a i l i n g edge down, deg t u n n e l f r e e - s t r e a m dynamic p r e s s u r e , l b / f t 2 wing a r e a , f t 2 temperature, OR t u n n e l f r e e - s t r e a m v e l o c i t y , f t / s e c engine i n l e t weight r a t e of flow, l b / s e c a n g l e of a t t a c k of wing chord plane, deg
hor i z o n t a1 - t a i l d i he dr a1 , de g
wing t r a i l i n g - e d g e f l a p d e f l e c t i o n , deg spanwise d i s t a n c e wing semispan s t a t i o n , b/2 S u b s c r i p t s j e t engine exhaust t a i l pipe h o r i z o n t a l t a i l wing t u n n e l f r e e stream MODEL Figure 1 shows t h e model i n s t a l l e d i n t h e wind t u n n e l and f i g u r e 2 ( a ) g i v e s t h e g e n e r a l arrangement and geometry of t h e model.
I The f u s e l a g e was c y l i n d r i c a l w i t h a n ogive nose and t a i l . The diameter The of t h e f u s e l a g e where it subtended t h e wing c e n t e r l i n e was 4.06 f t .
f i n e n e s s r a t i o w a s 16.6.
Wing The wing had an a s p e c t r a t i o of 2.17 with t h e l e a d i n g edge swept back 5 9 ' swept forward 1 0 ' . The a i r f o i l s e c t i o n w a s hexagonal with and t r a i l i n g edge 3 - p e r c e n t - c h o r d m a x i m u m t h i c k n e s s between the 30- and 70-percent-chord p o i n t s and a s t r a i g h t l i n e t a p e r frorri these p o i n t s t o t h e l e a d i n g and t r a L l i n g edges.
The leading-edge r a d i u s was 0.008 inch and t h e t r a i l i n g edge, 0.074 inch.
The t r a i l i n g - e d g e f l a p s were 15-percent chord extending from 1 3 t o 90 percent of t h e wing semispan. The f l a p s could be d e f l e c t e d t o 30°,0except for.
t h e s e c t i o n s above t h e n a c e l l e s , which could be d e f l e c t e d t o only 5 .
Horizontal T a i l The a x i s of r o t a t i o n of t h e all-movable h o r i z o n t a l t a i l was t h e q u a r t e r - * chord point of t h e mean aerodynamic chord; t h e incidence-angle range was 0 ' The t a i l droop was obtained by r o t a t i o n about l o n g i t u d i n a l axes t o -15'.
l o c a t e d i n t h e chord plane of t h e wing 8.0 inches outboard of' t h e f u s e l a g e Both t h e horizon- c e n t e r l i n e . The tail droop angles ranged from 0 ' t o -25'.
t a l and v e r t i c a l t a i l s had t h e same a i r f o i l s e c t i o n as t h e wing. Thermocou- p l e s were imbedded i n t h e upper s u r f a c e s k i n c l o s e t o t h e l e a d i n g edge t o i n d i c a t e t h e temperature rise as t h e t a i l was drooped through t h e exhaust j e t The thermocouples were l o c a t e d o n a constant chord of t h e inboard motors.
l i n e 0.040 Ft back from t h e l e a d i n g edge a t spanwise s t a t i o n s 0.29, 0.44, 0.60, and 0.76.
Engines and Nacelles J e t t h r u s t was provided by f o u r YJ-85 GE-5 engines mounted i n d i v i d u a l l y i n n a c e l l e s a t t a c h e d d i r e c t l y t o t h e lower s u r f a c e of t h e wing. The i n l e t s were s t r a i g h t ducts of double w a l l c o n s t r u c t i o n , t h e i n s i d e diameter being t h e same as t h a t of t h e compressor f a c e . The i n l e t l i p was b e v e l l e d and had a leading-edge r a d i u s of 0.06 inch.
The t h r u s t r e v e r s e r s were of t h e cascade type as shown i n d e t a i l i n f i g u r e 2 ( b ) . Several of t h e bays were blocked o f f , both t o reduce t h e open a r e a of t h e r e v e r s e r s and t o d i r e c t t h e h o t gas flow away from t h e engine support s t r u c t u r e . The vane angles were f i x e d a t 55' measured from a l i n e normal t o t h e engine t h r u s t axis. For t h e forward t h r u s t c o n f i g u r a t i o n s t h e t h r u s t . r e v e r s e r s were removed and t h e normal t a i l pipes s u b s t i t u t e d .
TESTS AND PROCEDUKE Force and moment d a t a were o b t a i n e d through a n angle-of'-attack range i'rom - 4 ' t o + 1 8 ' . Tlie t e s t Reynolds numbcr ranged from l 7 . % 1 O 6 t o 32.a<106 wliicli corresponded t o wind- Lunnel lYee-stream dynamic p r e s s u r e s from 25 t o 100 pounds per square r o o t . The t o t a l g r o s s t h r u s t from a l l f o u r engines was v a r i e d I'rom 6,230 pounds i n Yorward t h r u s t t o 1,770 pounds i n r e v e r s e t h r u s t .
Tlie reverse t h r u s t ei'l'iciency v a r i e d betwccn 42 and 52 p e r c e n t of t h e forward g r o s s t l i r u s t .
ma a m m a a a a a a a a a am m a e a a m a mo a am a a a m a m a m a a a am The forward t h r u s t of each engine was c a l i b r a t e d s t a t i c a l l y a g a i n s t t h e i n t a k e weight r a t e of flow t o o b t a i n n e t t h r u s t with forward v e l o c i t y . The drag due t o t h e r e v e r s e t h r u s t of each engine was c a l i b r a t e d a g a i n s t inJet a t weight rate of flow with t h e wind-tunnel free-stream dynamic p r e s s u r e 1 0 pounds p e r square f o o t . These data were then c o r r e c t e d f o r a i r p l a n e drag and r a m drag t o o b t a i n r e v e r s e gross t h r u s t .
The a i r p l a n e aerodynamic data were obtained a t various angles of a t t a c k b u t with constant engine i n l e t weight r a t e of flow and wind-tunnel dynamic * p r e s s u r e . The d a t a f o r no t h r u s t were obtained with t h e engine i n l e t s plugged.
CORRECTIONS The following wind-tunnel w a l l c o r r e c t i o n s were a p p l i e d t o t h e f o r c e and moment d a t a :
. -
ACm = -0.00939 CL The i n t a k e weight r a t e of flow and engine t h r u s t were c o r r e c t e d t o s t a n d a r d atmospheric conditions. The f o r c e d a t a are r e f e r r e d t o t h e wind axes system with moments taken about t h e quarter-chord p o i n t of t h e mean aero- dynamic chord.
RFSULTS Table I i s an index of t h e configurations tested during t h i s i n v e s t i g a - t i o n and t h e f i g u r e numbers t o which t h e y apply.
Fg/QS, as p r e - The data are p r e s e n t e d a t constant values of t h e r a t i o and i n forward t h r u s t a l s o i n terms of t h e parameter s e n t e d i n r e f e r e n c e 5, Fi/qmAi which i s d i s c u s s e d l a t e r .
F i g u r e 3 shows t h e e f f e c t of t a i l droop, without engine operation, on t h e l o n g i t u d i n a l c h a r a c t e r i s t i c s of t h e model.
F i g u r e 4 p r e s e n t s t h e e f f e c t of engine operation, a t s e v e r a l values of
r t = 0 '
t h r u s t , on the l o n g i t u d i n a l c h a r a c t e r i s t i c s of t h e b a s i c model where and it = - 5 ' .
F i g u r e 5 p r e s e n t s t h e e f f e c t of engine t h r u s t with t h e h o r i z o n t a l t a i l a t s e v e r a l a n g l e s of droop. Figure 6 shows t h e e f f e c t of i n c r e a s e d free- stream dynamic p r e s s u r e with engine t h r u s t increased t o o b t a i n values of . . ... ..
0 . 0 . 0 .
0 . ... 0 . . .
0 . 0 . ...
0 . 0.. . . . 0
Fn/%Ai similar t o those i n f i g u r e 5 ( b ) w i t h t h e same model c o n f i g u r a t i o n .
Figure 7 shows t h e l o n g i t u d i n a l c h a r a c t e r i s t i c s of t h e model a t n e a r l y t h e same v$Lues of F n / q A i f o r t h e same model configuration as t h a t of f i g u r e 5(b) except w i t h t h e h o r i z o n t a l t a i l a t 0 ' incidence angle.
...
The l o n g i t u d i n a l c h a r a c t e r i s t i c s of t h e model with t h e wing t r a i l i n g - edge f l a p s d e f l e c t e d t o loo, 20°, and 3 0 ' are presented i n f i g u r e s 8(a), ( b ) , and ( e ) , r e s p e c t i v e l y .
Figure 9 p r e s e n t s t h e v a r i a t i o n i n temperature r a t i o of t h e m a x i m u m recorded EiorizonLal- Lail s k i n temperature t o j e t engine exhaust t a i i - p i p e temperature with change i n t h e t h r u s t parameter f o r t h e t a i l droop Fn/%Ai angles t e s t e d .
Figures 10 through 1.3 show t h e e f f e c t of r e v e r s e t h r u s t on t h e l o n g i t u - Figures 10 and 1 1 p r e s e n t d a t a with t h e d i n a l c h a r a c t e r i s t i c s of t h e model.
h o r i z o n t a l t a i l undrooped a t a f r e e - s t r e a m dynamic p r e s s u r e of 50 and 25 pounds per square f o o t , r e s p e c t i v e l y .
Figure 1 2 p r e s e n t s d a t a with t h e t a i l drooped -15'. Figure 1.3 shows r e s u l t s obtained with t h e wing f l a p s d e f l e c t e d 30' and t h e t a i l undrooped a t two t a i l incidences.
Figure 14 shows t h e v a r i a t i o n of p i t c h i n g moment with change i n
F n / L A i a t s e v e r a l values of Figure 1 5 shows t h e v a r i a t i o n of p i t c h i n g - s, and F t .
moment increment a t a = 0 ' due t o change i n Fn/GAi along with c a l c u l a t e d engine thrusL c o n t r i b u t i o n t o t h e pitching-moment increment a t s e v e r a l f l a p d e f l e c t i o n angles.
Figure 1 6 shows t h e pitching-moment increment due t o v a r i a t i o n of r e v e r s e t h r u s t Fg/%S f o r f l a p s u n d e f l e c t e d and d e f l e c t e d 30'. Calculated pitching-moment increments due t o r e v e r s e t h r u s t are a l s o p r e s e n t e d i n t h e f i g u r e .
DISCUSSION C o r r e l a t i o n Parameter Fn/qmAi The parameter F g / ~ S was i n t r o d u c e d i n r e f e r e n c e 5 t o enable c o r r e l a - t i o n of t h e e f f e c t s of engine t h r u s t r e v e r s a l on t h e aerodynamic c h a r a c t e r i s - t i c s of an a i r p l a n e over a range of t h r u s t and v e l o c i t y . I n o r d e r t o c o r r e l a t e t h e e f f e c t s of forward t h r u s t on t h e a i r p l a n e c h a r a c t e r i s t i c s i n t h e p r e s e n t i n v e s t i g a t i o n t h e parameter Fn/qmAi i s used which i s t h e r a t i o Of t h e change of t h e momentum of t h e n e t forward t h r u s t t o t h e momentum of t h e free-stream a i r flow i n t o t h e engines.
I n f i g u r e lb t h e results from f i g u r e s 4, 5, and 6 are compared i n terms Fn/%,Ai; reasonable of t h e v a r i a t i o n of p i t c h i n g moment w i t h momentum r a t i o c o r r e l a t i o n exists over t h e range of t h r u s t and dynamic p r e s s u r e s i n v e s t i - gated. It t h e r e f o r e appears t h a t t h e parameter should be usable i n determin- i n g t h e e f f e c t s of t h r u s t on t h e c h a r a c t e r i s t i c s of a f u l l - s c a l e a i r p l a n e from those of a l a r g e - s c a l e model presented h e r e i n by r e l a t i n g t h e r e s u l t s with as follows, t h i s c o r r e l a t i o n f a c t o r
(L) =(3)
q-Ai Model %*i Airplane Forward Thrust
I n general, f i g u r e s 4 through 8 show l i t t l e aerodynamic e f f e c t on t h e
l o n g i t u d i n a l c h a r a c t e r i s t i c s of t h e model o t h e r than t h a t produced by t h e d i r e c t t h r u s t f o r c e s of t h e engines. However, w i t h t h e f l a p s d e f l e c t e d 1 0 ' ( f i g . l5), t h e c a l c u l a t e d pitching-moment increment, ACm, due t o d i r e c t t h r u s t was g r e a t e r t h a n t h e t o t a l measured A C , , i n d i c a t i n g t h a t t h e flow f i e l d induced by t h e engine exhaust increased t h e pitch-dgwn moment due t o f l a p d e f l e c t i o n . 6f = 2 0 ° and 30 at low values of This w a s t r u e a l s o f o r Fn/LAi, b u t a t t h e higher values, t h e induced flow f i e l d changed so t h a t t h e measured ACm i s g r e a t e r t h a n t h e c a l c u l a t e d d i r e c t t h r u s t c o n t r i b u t i o n .
With t h e h o r i z o n t a l t a i l drooped Oo, - l 5 ' , o r - 2 5 ' , t h e j e t exhaust had e s s e n t i a l l y no e f f e c t on t h e s t a t i c l o n g i t u d i n a l s t a b i l i t y throughout t h e a n g l e - o f - a t t a c k range i n v e s t i g a t e d ( f i g . 5 ) .
Horizontal - T a i l Skin Temperatures Figure 9 shows t h e v a r i a t i o n of t h e m a x i m u m h o r i z o n t a l - t a i l s k i n tempera- f o r t a i l droop angles of Oo, - 1 5 ' , t u r e with change i n values of F n / ~ A i and - 2 5 ' . T t - T,/TE - T , .
The t a i l temperatures are represented by t h e r a t i o Undrooped, t h e h o r i z o n t a l t a i l w a s above the exhaust j e t and t h e s k i n tempera- t u r e w a s only s l i g h t l y higher t h a n free-stream temperature even a t Drooped -15O, Fn/q&i = 14, t h e approximate value f o r take-off a t 160 knots.
F /q A i = 1 4 t h e t h e t a i l w a s c l o s e t o t h e c e n t e r of t h e exhaust j e t , and a t
v -
t a i l temperature i n c r e a s e was approximately 60 percent of t h e increment of 'temperature between t h e t u n n e l free stream and t h e exhaust j e t .
A t h i g h subsonic Mach number where t h e h o r i z o n t a l t a i l would be drooped t h e value of would through t h e exhaust p r i o r t o supersonic f l i g h t , Fn/GAi To t h e e x t e n t be approximately 2.0 a t a Mach number of 0.9 a t 40,000 f e e t .
t h a t Fn/\Ai can be considered a c o r r e l a t i n g parameter f o r temperature when based on t h e d a t a of f i g u r e 9, t h e r e s u l t i n g temperature rise would be about 35 p e r c e n t of t h e d i f f e r e n c e between f r e e - s t r e a m and j e t - e x h a u s t temperatures.
With t h e h o r i z o n t a l t a i l drooped t o - 2 5 ' , t h e temperature rise was consider- a b l y less t h a n a t - 1 5 ' droop s i n c e t h e t a i l w a s again o u t s i d e of t h e j e t - exhaust stream.
Reverse Thrust The e n t i r e r e v e r s e - t h r u s t p o r t i o n of t h e i n v e s t i g a t i o n was conducted w i t h the engines o p e r a t i n g a t f u l l r e v e r s e t h r u s t . The r e s u l t i n g aerodynamic.
e f f e c t s on t h e p i t c h i n g moment were l a r g e .
The d a t a presented i n f i g u r e s 10 and 1 1 i n d i c a t e t h a t with t h e wing .
s h i f t i n t h e negative d i r e c - t r a i l i n g - e d g e f l a p s undeflected t h e r e was a Cm t i o n which :EX opposite t o and of much g r e a t e r magnitude t h a n the s h i f t I n a d d i t i o n t o t h e caused by forward t h r u s t a t comparable values of F /%Sa CL = 0.3 s h i f t i n Cm, t h e r e was a l s o a r e d u c t i o n i n t h e s t a g i l i t y above which w a s probably due t o e i t h e r a r e d u c t i o n i n dynamic p r e s s u r e o r a change ' i n downwash angle a t t h e t a i l as a r e s u l t of t h r u s t r e v e r s a l . Figure 1 2 Figure 13 shows e s s e n t i a l l y t h e same r e s u l t s with t h e t a i l drooped t o -15'.
shows t h a t with t h e f l a p s d e f l e c t e d 30°, t h e r e d u c t i o n i n s t a t i c l o n g i t u d i n a l s t a b i l i t y was considerably g r e a t e r t h a n with f l a p s at 0 ' .
ACm t h a t occurs with Figure 15 i n d i c a t e s t h a t a t 0 ' angle of a t t a c k t h e t h e engine o p e r a t i n g i n forward t h r u s t i s due l a r g e l y t o t h e d i r e c t t h r u s t I n c o n t r a s t , f i g u r e 16 f o r c e s and i s n o t g r e a t l y a f f e c t e d b y f l a p d e f l e c t i o n .
shows t h a t with t h r u s t r e v e r s a l , t h e ACm due t o engine o p e r a t i o n i s g r e a t e r and of opposite d i r e c t i o n t o t h a t of t h e forward t h r u s t . The aerodynamic e f f e c t s arc iiiucli g r e a t e r than t h e e f f e c l s of Lhe d i r e c t r e v e r s e t h r u s t f o r c e s .
D e f l e c t i n g t h e f l a p s t o 30' produced o n l y a small change i n t h e p i t c h i n g moment r e l a t i v e t o t h e values obtained with 0 ' f l a p d e f l e c t i o n with t h r u s t r e v e r s a l .
The model as t e s t e d was capable of only f u l l forward o r f u l l r e v e r s e t h r u s t , t h e t h r u s t being c o n t r o l l e d by t h e engine speed. Therefore, it w a s not determined whether modulated t h r u s t r e v e r s a l , as s t u d i e d i n r e f e r e n c e 6, would prevent t h e l a r g e changes i n l o n g i t u d i n a l s t a b i l i t y caused by f u l l reverse t h r u s t .
Ames Research Center National Aeronautics and Space Administration
Moffett F i e l d , C a l i f . , J u l y 7, 1964
REFEFENCES 1. Brady, James A., Page, V. Robert, and Koenig, David G.: Large-Scale Low- Speed Wind-Tunnel Tests of a Delta Winged Supersonic Transport Model With a Delta Canard Control Surface. NASA TM X-643,1962.
?. Koenig, David G., Brady, James A., and Page, V. Robert: Large-Scale Wind- Tunnel Tests at Low Speed of a Delta Winged Supersonic Transport Model in the Presence of the Ground. NASA TM X-644, 1962.
3. Koenig, David G., and Corsiglia, Victor R.: Large-Scale Low-Speed Wind-
Tunnel Tests of a Delta Winged Supersonic Transport Model With Various Canard, Horizontal Tail, and Wing Modifications. NASA TM X-857, 1964.
4. Fletcher, LeRoy S.: Static Stability Characteristics of a Delta-Winged
Airplane Configuration With Nacelles, a Trapezoidal Canard and a NASA D 4 x-780, 1963.
Drooped Tail at Mach Numbers From 0.70 to 3.52.
5. Tolhurst, William H., Jr., Kelly, Mark W., and Greif, Richard K. : Full-
Scale Wind-Tunnel Investigation of the Effects of a Target-Type Thrust Reverser on the Low-Speed Aerodynamic Characteristics of a Single Engine Jet Airplane. NASA TN D-72, 1959.
6. Hickey, David H., Tolhurst, William H., Jr., and Aoyagi, Kiyoshi: Investigation of the Longitudinal Characteristics of a Large-Scale Jet NASA Transport Model Equipped With Controllable Thrust Reversers.
TN D-786, 1961.
..
TABLF: I. - MODEL C O ~ I C ~ A T I O N S FOR WHICH T~HEELCOMPONENT FORCE DATA ARE PRFSENTED Figure 0, -15, -25 -5
- 1
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