large-scale tind-tunnel investigation of an mi-l lifting ... · the mi-l lifting body concept...
TRANSCRIPT
LARGE-SCALE WIND-TUNNEL INVESTIGATION OF A N MI-L LIFTING BODY WITH A N INFLATABLE AND A RIGID AFTERBODY
bY Kenneth W. Mort
I A m e s Reseurch Center
und M i c h e l D. Falurski Army Aeronuzcticul Reseurch Laborutory
N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N W A S H I N G T O N , D. C. O C T O B E R 1 9 6 9
https://ntrs.nasa.gov/search.jsp?R=19690029390 2020-01-08T07:39:17+00:00Z
TECH LIBRARY KAFB, “I
19. Secur i ty Classif. (of t h i s report)
Unclassified
2. Governmen t A c c e s s i o n No. I 1. Report No. NASA TN D-5468
4. T i t l e a n d Sub t i t l e LARGE-SCALE WIND-TUNNEL INVESTIGATION O F AN M1-L LIFTING BODY WITH AN INFLATABLE AND A RlGID AFTERBODY
I
7. A u t h o d s ) Kenneth W. Mort and Michael D. Falarski
9. Performing Organ iza t ion Name and A d d r e s s
NASA Ames Research Center
20. Secur i ty C l a s s i f . (of t h i s p a g e ) of P a g e s 22. P r i c e *
Unclassified $3.00
and Army Aeronautical Research Laboratory Molfett Field, Calif. 94035
I 112. Sponsor ing Agency Name a n d A d d r e s s
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION Washington, D. C . , 20546
3. R e c i p i e n t ‘ s C a t a l o g N o .
5. Repor t D a t e October 1969
6. Pe r fo rming Organ iza t ion C o d e
8. Pe r fo rming Organ iza t ion Repor! A-2894
I O . Work Unit N o . 124-07-02-16-00- 21
11. C o n t r a c t or Gran t No.
13. T y p e of Repor t a n d P e r i o d Cov
TECHNICAL NOTE
14. Sponsor ing A g e n c y C o d e
115. Supp lemen ta ry N o t e s
The MI-L lifting body concept employs a high volumetric elficiency forebody and an inllatahle afterbody that i s
In deployed to provide lilt-to-drag rat ios sufficient fo r horizontal landings. The purpose of the investiLmtion descr ibed here was to determine the low-speed aerodynamic character is t ics of a large-scale model with an inllatable afterbody. addition, a model with a rigid afterbody was tested as a datum fo r aeroelast ic character is t ics .
than 2. of about 2-1/2 i s the niinlmum value required to accomplish a horizontal landing, this 20-percent reduction appeared to be the most s iq i f i can t effect of flexibility.
Deployment of the inflatable afterbody increased the maximum lift-to-drag ratio (L/D) l rom l e s s than 1 to slightly nior Because a maximum L, This value was about 15 to 20 percent less than that fo r the model with the rigid alterbody.
Studies indicated that depluyment of the inflatable afterbody i s mechanically feasible.
I I 17. Key Words Suggested by Author(s)
Lifting body Inflatable afterbody Subsonic aerodynamics, a i r c ra f t Subsonic aerodynamics, missi les and space vehicles
18. Dis t r ibu t ion S ta t emen t
Unclassified - Unlimited
1 1 1 I *For sale by the Clear inghousefor Federal Scientific and Technical Inlormation
Springfield, Virginia 22151
LARGE-SCALE WIND-TUNNEL INVESTIGATION OF AN M1-L LIFTING BODY
WITH AN INFLATABLE AND A R I G I D AFTERBODY
By Kenneth W . Mort Ames Research Center
and
Michael D . Fa l a r sk i Army Aeronaut ical Research Laboratory
SUMMARY
The M1-L l i f t i n g body concept employs a high volumetr ic e f f i c i e n c y forebody and an i n f l a t a b l e a f te rbody t h a t i s deployed t o provide l i f t - t o - d r a g r a t i o s s u f f i c i e n t f o r h o r i z o n t a l l andings . The purpose of t h e i n v e s t i g a t i o n descr ibed he re was t o determine t h e low-speed aerodynamic c h a r a c t e r i s t i c s of a l a rge - sca l e model wi th an i n f l a t a b l e a f te rbody. In a d d i t i o n , a model with a r i g i d a f te rbody was t e s t e d as a datum f o r a e r o e l a s t i c c h a r a c t e r i s t i c s .
Deployment of t h e i n f l a t a b l e a f te rbody increased t h e maximum l i f t - t o - d r a g r a t i o (L/D) from less than 1 t o s l i g h t l y more than 2 . This va lue was about 15 t o 20 percent less than t h a t f o r t h e model with t h e r i g i d a f te rbody. Because a maximum L / D of about 2 - 1 / 2 i s t h e minimum value r equ i r ed t o accom- p l i s h a ho r i zon ta l l anding , t h i s 20-percent reduct ion appeared t o be t h e most s i g n i f i c a n t e f f e c t of f l e x i b i l i t y .
S tudies ind ica t ed t h a t deployment of t h e i n f l a t a b l e a f te rbody i s mechanically f e a s i b l e .
INTRODUCTION
Many s t u d i e s have been conducted i n developing l i f t i n g body r e e n t r y conf igura t ions capable of g l i d i n g t o a s p e c i f i e d recovery s i t e and making a conventional h o r i z o n t a l landing. One such concept i s r e f e r r e d t o as t h e M1-L l i f t i ng -body conf igu ra t ion . A high volumetr ic e f f i c i e n c y forebody (30" h a l f - cone) i s used as t h e prime space veh ic l e , and an i n f l a t a b l e af terbody i s deployed a t low speeds t o provide adequate l i f t - t o - d r a g r a t i o s f o r performing t h e landing maneuver. Several smal l - sca le i n v e s t i g a t i o n s examined t h e aero- dynamic c h a r a c t e r i s t i c s of t h e r e e n t r y conf igu ra t ion , o r forebody, and a r e repor ted i n r e fe rences 1 through 9 . To i n v e s t i g a t e t h e aerodynamic cha rac t e r - i s t i c s of t h e concept with t h e i n f l a t a b l e a f te rbody, two l a rge - sca l e models were cons t ruc ted : a model with an i n f l a t a b l e a f te rbody t h a t could be used f o r both s t e a d y - s t a t e aerodynamic measurements and deployment i n v e s t i g a t i o n s , and a model with a r i g i d a f te rbody used as a datum f o r a e r o e l a s t i c cha rac t e r - i s t ics . T e s t s of t h e s e models were performed i n t h e Ames 40- by 80-Foot Wind Tunnel, and t h e r e s u l t s are presented h e r e i n .
NOTAT I ON
b maximum span, 13.4 f t
CD
CL
cz
Cm
Cn
D
2
L
9
r
R
S
V
X
Y
0.
b f
2
D drag c o e f f i c i e n t , - qs
L l i f t c o e f f i c i e n t , - qs
r o l l i n g moment qSb
rolling-moment c o e f f i c i e n t ,
s i d e fo rce qs
s ide - fo rce c o e f f i c i e n t ,
p i t c h i n g moment qs 2
pitching-moment c o e f f i c i e n t ,
yawing moment yawing-moment c o e f f i c i e n t , qSb
drag f o r c e , l b
o v e r a l l l ength of model with a f te rbody, 1 4 . 1 f t
l i f t f o r c e , l b
f ree-s t ream dynamic p res su re , p s f
r a d i a l d i s t a n c e from cone ax i s (see f i g . 3 ) , i n .
VZ Reynolds number , kinematic v i s c o s i t y
re ference area (body planform area with a f t e rbody) , 142 f t 2
f ree-s t ream v e l o c i t y , f p s
d i s t a n c e from leading edge (see f i g . 3 ) , i n .
d i s t ance above cone a x i s , along a x i s (see f i g . 3 ) , i n .
angle of a t t a c k (see f i g . 3 ) , deg
f l a p d e f l e c t i o n (see f i g . 3 ) , deg
6, rudder d e f l e c t i o n , both s i d e s are moved toge the r ( see f i g . 3 ) , deg
Apc i n t e r n a l p re s su re of i n f l a t a b l e a f te rbody c a v i t y re ferenced t o atmospheric p re s su re , p s i
Api
The fo rces and moments developed by t h e model were reso lved along t h e wind a x i s .
i n t e r n a l p re s su re of primary s t r u c t u r e re ferenced t o atmospheric p re s su re , p s i
MODEL DESCRIPTION
Photographs of t h e models i n s t a l l e d i n t h e t e s t s e c t i o n of t h e Ames 40- by 80-Foot Wind Tunnel are shown i n f i g u r e 1. Photographs showing t h e model with t h e i n f l a t a b l e a f te rbody i n var ious s t a g e s of assembly a r e shown i n f i g u r e 2 . Basic model dimensions and geometry are presented i n f i g u r e 3 and t a b l e I .
The model with t h e i n f l a t a b l e a f te rbody was cons t ruc ted by Goodyear Aerospace Corporat ion. made of Goodyear A i r m a t . l r e a r su r f ace , t h e f i n s and con t ro l su r f aces , and t h e i n t e r n a l v e r t i c a l members. (see f i g s . 2(b) , 2(c) , and 3 ) . Fur ther d e t a i l s of t h e design and construc- t i o n a r e given i n r e fe rence 1 0 . a t 5 t o 25 p s i , and t h e i n t e r i o r of t h e a f te rbody o r c a v i t y was i n f l a t e d wi th a i r a t 0 .1 t o 0.2 p s i .
The primary s t r u c t u r e of t h e i n f l a t a b l e a f te rbody was This s t r u c t u r e included t h e upper su r face , t h e
The ca tenary and remaining af te rbody e x t e r i o r were rubberized f a b r i c
The primary s t r u c t u r e was i n f l a t e d with a i r
TEST PROCEDURE
The t e s t s were performed by varying angle of a t t ack f o r var ious i n t e r n a l p re s su res , f ree-s t ream dynamic p res su res , and con t ro l s e t t i n g s .
Deployment t e s t i n g was performed a t dynamic pressures of 24 and 62 p s f . The procedure involved r e l e a s i n g t h e harness t h a t r e s t r a i n e d t h e d e f l a t e d a f te rbody, i n f l a t i n g t h e primary s t r u c t u r e with high p res su re a i r contained i n compressed a i r b o t t l e s i n t h e forebody, and i n f l a t i n g t h e a f te rbody c a v i t y with ram a i r from t h e scoop on t h e lower s u r f a c e of t h e forebody (see f i g . 3 ) . Signals from p res su re senso r s i n t h e cav i ty ac tua t ed a d r i v e system t o open and c l o s e t h e scoop au tomat ica l ly t o r e g u l a t e t h e c a v i t y p re s su re . During the s t e a d y - s t a t e t e s t i n g , t h i s scoop was c losed t o provide more uniform cav i ty p re s su re (suppl ied from an e x t e r n a l source) and t o e l imina te t h e random flow e f f e c t s t h a t would appear as sca t te r i n t h e d a t a from t h e ape r iod ic opening and c los ing of t h e scoop.
~
'Trademark, Goodyear Aerospace Corporat ion, Akron, Ohio.
3
REDUCTION OF DATA
Correc t ions
No tunnel -wal l co r rec t ions were app l i ed t o t h e d a t a presented because es t imates ind ica t ed t h a t such e f f e c t s on t h e d a t a were wel l w i th in t h e ind ica t ed accuracy.
The d a t a were co r rec t ed for t a r e s obta ined f o r t h e unshielded s t r u t s . These t a r e s were obtained without t h e models and hence a r e s u b j e c t t o e r r o r s from d i f f e rences due t o i n t e r a c t i o n with t h e models. However, experience with o the r models us ing these same s t r u t s i n d i c a t e s t h a t t h e i n t e r f e r e n c e e f f e c t s a r e small .
Accuracy of Measurements
The var ious q u a n t i t i e s measured i n t h e wind tunnel were accu ra t e wi th in t h e fol lowing l i m i t s . The values given inc lude e r r o r l i m i t s due t o c a l i b r a - t i o n s , co r rec t ions , and record ing methods. The f o r c e p re s su re , and moment measurements f o r each d a t a po in t were obta ined by averaging 10 samples. Hence, t he accuracy l i m i t s l i s t e d f o r t h e s e i tems are f o r t h e average va lues .
Angle o f a t t a c k t o . 3'
L i f t 210 l b
Drag t 3 l b
Side fo rce 23 l b
P i tch ing moment k300 f t - l b
Rol l ing moment t 4 0 0 f t - l b
Yawing moment k100 f t - l b
Free-stream dynamic p res su re k 0 . 5 percent above 20 p s f , 20.1 p s f below 20 psf
Pr imary-s t ruc ture p re s su re t 0 . 5 p s i
Cavity p re s su re t 0 . 0 2 ps i
Control s u r f a c e s e t t i n g s + 2 O
4
I
RESULTS
Table I1 is a complete index t o t h e f i g u r e s which show t h e t e s t r e s u l t s . Basic long i tud ina l aerodynamic c h a r a c t e r i s t i c s are presented i n f i g u r e s 4 through 10. Resul t s are shown i n f i g u r e 4 with t h e af terbody o f f and i n f i g u r e s 5 through 9 wi th t h e i n f l a t a b l e a f te rbody on. Data f o r s e v e r a l i n f l a t i o n p res su res and dynamic p res su res a r e p re sen ted . Figure 10 shows r e s u l t s f o r t h e r i g i d model. Longi tudinal aerodynamic c h a r a c t e r i s t i c s w i th and without t he a f t e r b o d i e s are summarized i n f i g u r e 11. A comparison of t h e r e s u l t s f o r t h e i n f l a t a b l e and r i g i d model i s presented i n f i g u r e 1 2 . I n f i g u r e 13 photographs of t h e model wi th t h e i n f l a t a b l e a f te rbody show t h e d e f l e c t i o n of t h e a f te rbody dur ing t e s t i n g . Longi tudinal and d i r e c t i o n a l con t ro l e f f ec t iveness d a t a are presented i n f i g u r e s 14 and 15, r e s p e c t i v e l y , f o r both models. Photographs of var ious s t a g e s of t h e deployment of t h e i n f l a t a b l e a f te rbody are shown i n f i g u r e 16. The r e s u l t s shown i n f i g u r e s 11 through 16 w i l l be d iscussed i n t h e next s e c t i o n .
DISCUSSION
The E f f e c t s of t h e I n f l a t a b l e Afterbody
As s t a t e d i n t h e in t roduc t ion t h e purpose of t h e i n f l a t a b l e a f te rbody was t o improve t h e long i tud ina l aerodynamic c h a r a c t e r i s t i c s of t h e forebody s u f f i c i e n t l y t o allow a h o r i z o n t a l landing. To examine t h i s improvement, f i g u r e 11 is presented showing t h e aerodynamic c h a r a c t e r i s t i c s of t h e fo re - body with and without t h e i n f l a t a b l e a f t e rbody .2 I t i s ev ident from t h i s f i g u r e t h a t t he i n f l a t a b l e a f te rbody s i g n i f i c a n t l y improved t h e aerodynamic c h a r a c t e r i s t i c s of t h e model. The l i f t curve s lope was increased from about 0.012 t o about 0 .027 p e r degree , drag c o e f f i c i e n t a t zero l i f t was reduced from 0.32 t o 0 . 1 2 , and t h e s t a t i c margin went from 0.1045 t o -0.110 (at Cm = 0 ) . Maximum L/D, which is probably the most important parameter , was increased from less than 1 t o somewhat over 2 . As i nd ica t ed by s e v e r a l o t h e r s t u d i e s ( e . g . , r e f . l l ) , a maximum L / D o f 2 - 1 / 2 i s about t h e minimum value requi red f o r h o r i z o n t a l landing c a p a b i l i t y .
A f t erbody F l e x i b i li t y
The aerodynamic c h a r a c t e r i s t i c s are shown i n f i g u r e 11 f o r t h e model with t h e i n f l a t a b l e a f te rbody a t an i n f l a t i o n p res su re of 10 p s i i n t h e primary s t r u c t u r e , a long i tud ina l con t ro l s e t t i n g of - 3 0 ° , and a f ree-s t ream dynamic p res su re of 50 p s f . t o affect t h e aerodynamic c h a r a c t e r i s t i c s because of a f te rbody f l e x i b i l i t y . To examine t h i s , r e s u l t s are shown i n f i g u r e 1 2 f o r several i n f l a t i o n p res - su res and con t ro l s e t t i n g s ; t h e dynamic p res su re is 25 p s f . Results f o r t h e
Var i a t ions i n t h e s e parameters would be expected
2These d a t a were obta ined from f i g u r e s 4 and 9 .
5
model wi th t h e r i g i d a f te rbody are a l s o shown.3 s lope and s t a b i l i t y vary with i n t e r n a l p re s su re and f l a p s e t t i n g . between t h e two models are a l s o ev iden t . The L/D does not appear t o vary apprec iab ly with p re s su re ; however, it i s always lower f o r t h e model with t h e i n f l a t a b l e a f te rbody - t h e maximum value i s about 20 pe rcen t lower. The reasons f o r t h i s can be seen i n photographs i n f i g u r e 13. A s can be seen, t h e r e were wrinkles a long t h e s i d e and bulging of t h e lower s u r f a c e near t h e junc tu re of t h e forebody and af terbody t h a t caused t h e drag t o be h ighe r .4 (The minimum drag c o e f f i c i e n t was about 50 percent h ighe r . )
I t i s ev ident t h a t l i f t - c u r v e Differences
The effect of dynamic p res su re on a f te rbody f l e x i b l i t y was compared f o r t h e two models wi th t h e con t ro l s on a t var ious dynamic p r e s s u r e s . Increas ing t h e dynamic p res su re from 25 t o about 50 p s f (maximum f o r which comparisons could be made) d i d no t s i g n i f i c a n t l y change t h e incremental d i f f e r e n c e s between the i n f l a t a b l e and r i g i d models a t primary s t r u c t u r e p re s su res equal t o o r g r e a t e r than 10 p s i . (For t h i s reason , t h e i n t e r n a l p re s su re of t h e primary s t r u c t u r e was not macle dimensionless with dynamic p res su re . ) ing the dynamic p res su re below 25 p s f a l s o d id not s i g n i f i c a n t l y change t h e incremental d i f f e rences between t h e two models b u t showed l a r g e v a r i a t i o n s i n aerodynamic c h a r a c t e r i s t i c s due t o flow c h a r a c t e r i s t i c s s e n s i t i v e t o Reynolds number ( see , e . g . , f i g s . 5 (a) and 1 0 ) .
Decreas-
Control Ef fec t iveness
The long i tud ina l con t ro l e f f ec t iveness of t h e two models i s presented i n f i g u r e 14. The con t ro l e f f ec t iveness of t he i n f l a t a b l e model was l e s s than t h a t of t he r i g i d model (15 t o 30 percent depending on the angle of a t t a c k and dynamic p r e s s u r e ) . i n f l a t a b l e con t ro l su r f aces , although it should be poin ted out t h a t t h e con- t r o l s were d i f f e r e n t : t h e hinge l i n e was s e a l e d f o r t h e i n f l a t a b l e c o n t r o l s and not s ea l ed f o r t h e r i g i d c o n t r o l s . (The ev ident n o n l i n e a r i t y from 0" t o -10" d e f l e c t i o n occurred f o r t h e r i g i d c o n t r o l s as w e l l as f o r t h e i n f l a t a b l e con t ro l s and hence appears t o be due mainly t o t h e con t ro l system conf igu ra t ion . )
This d i f f e r e n c e could be due t o d e f l e c t i o n of t h e
Di rec t iona l con t ro l e f f ec t iveness f o r t h e two models was s i g n i f i c a n t l y d i f f e r e n t . A s shown i n f i g u r e 15 t h e con t ro l e f f e c t i v e n e s s of t h e i n f l a t a b l e model appears t o be nonl inear , whereas t h a t o f t h e r i g i d model i s e s s e n t i a l l y l i n e a r . A t rudder d e f l e c t i o n s g r e a t e r than 10" t h e i n f l a t a b l e model appears t o have h igher 'control e f f ec t iveness than t h e r i g i d model. Again, t h e hinge l i n e w a s sea led f o r t h e i n f l a t a b l e con t ro l su r f aces and not f o r t h e r i g i d con t ro l su r f aces .
3These r e s u l t s were determined from t h e d a t a of f i g u r e s 5 , 7 , 9 , 10, and
4The wrinkles and bulging appear t h e same as were evident i n f i g u r e 1, 14 ; some minor ex t r apo la t ion was r equ i r ed .
which were due t o drooping of t h e a f te rbody about t h e upper-surface i n t e r s e c - t i o n of t h e forebody and af te rbody because of i t s own weight.
6
Deployment
Deployment of t h e i n f l a t a b l e af terbody was s tud ied t o e s t a b l i s h t h e mechanical f e a s i b i l i t y . Two deployment sequences a r e shown i n f i g u r e 16. The f irst deployment was made a t a dynamic p res su re of 24 p s f ( f i g , 1 6 ( a ) ) . The primary s t r u c t u r e was f u l l y i n f l a t e d i n about 10 seconds; however , , the c a v i t y d id not i n f l a t e as r a p i d l y as was intended because t h e scoop on t h e bottom of t h e forebody (see f i g . 3 ) , which was t h e ram a i r i n l e t f o r t h e c a v i t y , d i d not func t ion proper ly . The second deployment was made a t a dynamic p res su re of 62 ps f ( f i g . 1 6 ( b ) ) . During t h i s deployment t h e c a v i t y i n f l a t i o n system operated proper ly , and t h e c a v i t y was completely i n f l a t e d i n about 3 seconds; bu t t h e high p res su re system t h a t i n f l a t e s t h e primary s t r u c t u r e f a i l e d t o opera te , hence, it was not i n f l a t e d .
Even though t h e primary s t r u c t u r e and c a v i t y were no t i n f l a t e d s imultaneously during e i t h e r deployment, it i s apparent from t h e photographs of f i g u r e 16 t h a t t h e shape of t h e f i n s was good i n one deployment and t h a t t he shape of t he a f te rbody w a s good i n t he o t h e r . This sugges ts t h a t deploy- ment of t h e i n f l a t a b l e a f te rbody is mechanically f e a s i b l e and would have been success fu l i f both i n f l a t i o n systems had operated s imul taneous ly .
CONCLUDING REMARKS
Deployment of t h e i n f l a t a b l e a f te rbody increased t h e maximum L / D from less than 1 t o s l i g h t l y more than 2 . A s i nd ica t ed by s e v e r a l s t u d i e s ( see , e . g . , r e f . 11) a maximum L / D of 2 - 1 / 2 i s about t h e minimum value r equ i r ed f o r a ho r i zon ta l landing. Hence, t h e maximum L / D of t h e model wi th t h e i n f l a t a b l e a f te rbody i s too low. The maximum L / D f o r t h e model with t h e r i g i d af terbody i s only about 2-3/4, which i s near t h e minimum requ i r ed f o r a ho r i zon ta l landing.
The maximum L/D f o r t h e model with t h e i n f l a t a b l e a f te rbody is about 15 t o 20 percent lower than t h a t of t h e r i g i d model because of increased drag caused by bulg ing , wr inkles , e t c . The bulging and wr inkles depend on t h e design and m a t e r i a l s ; however, any i n f l a t a b l e s t r u c t u r e w i l l probably have h ighe r drag than a r i g i d s t r u c t u r e . with s u f f i c i e n t l y h igh maximum L / D t o allow f o r l o s ses due t o f l e x i b i l i t y . I t i s apparent t h a t t h e M1-L con f igu ra t ion does not have s u f f i c i e n t l y h igh L/D.
Hence, it i s necessary t o use body shapes
Deployment of t h e i n f l a t a b l e a f te rbody was considered mechanically f e a s i b l e . bu t t hese problems were not r e l a t e d t o t h e conf igura t ion o r t o t h e method o f deployment.
Operat ional problems were encountered dur ing deployment s t u d i e s ,
Ames Research Center Nat ional Aeronautics and Space Adminis t ra t ion
Moffet t F i e ld , C a l i f . , 94035, June 1 2 , 1969
7
REFERENCES
1.
2.
3 .
4 .
5.
6.
7 .
8.
9 .
10.
11.
Eggers, Al f red J . , J r . ; and Wong, Thomas J . : Re-entry and Recovery of Near-Earth S a t e l l i t e s , With P a r t i c u l a r A t t en t ion t o a Manned Vehicle . NASA MEMO 10-2-581\, 1958.
Savage, Howard F . ; and T in l ing , Bruce E . : Subsonic Aerodynamic Charac- ter is t ics of Severa l Blunt, L i f t i n g Atmospheric-Entry Shapes. NASA MEMO 12-24-58AY 1959.
Hassell, James L . , J r . : I n v e s t i g a t i o n of t h e Low-Subsonic S t a b i l i t y and Control C h a r a c t e r i s t i c s of a 1/3-Scale Free-Flying Model of a L i f t i ng - Body Reentry Configurat ion. NASA TM X-297, 1960.
Dennis, David H . ; and Edwards, George G . : The Aerodynamic Character- i s t ics of Some L i f t i n g Bodies. NASA TM X-376, 1960.
Sarabia , Michael F . : Aerodynamic C h a r a c t e r i s t i c s of Blunt Half-Cone Entry Configurat ion a t Mach Numbers From 3 t o 6 . NASA TM X-393, 1960.
Tunnel l , P h i l l i p s J . : The S t a t i c and Dynamic S t a b i l i t y Der iva t ives of a Blunt Half-Cone Entry Configurat ion a t Mach Numbers From 0.70 t o 3.50. NASA TM X-577, 1961.
DeRose, Charles E . : The Aerodynamic C h a r a c t e r i s t i c s of a Blunt Half- Cone Entry Configurat ion Obtained From Bal l i s t ic Range Tes t s a t Mach Numbers Near 3 . NASA TM X-578, 1961.
Holtzclaw, Ralph W . : S t a t i c S t a b i l i t y and Control C h a r a c t e r i s t i c s of a Half-Cone Entry Configurat ion a t Mach Numbers From 2 . 2 t o 0 .7 . NASA TM X-649, 1962.
Holdaway, George H . ; Polek, Thomas E . ; and Kemp, Joseph H . , J r . : Aero- dynamic C h a r a c t e r i s t i c s of a Blunt Half-Cone Entry Configurat ion a t Mach Numbers of 5 .2 , 7.4, and 10 .4 . NASA TM X-782, 1963.
Anon.: M-1-L Re-entry Vehicle (Wind Tunnel Model, Fu l l - sca l e ) Design, Fabr i ca t ion , and Tes t ing . F ina l Report , 10 Aug. 1962 - 15 Nov. 1963, Goodyear Aerospace Corporation Report Number GER-11141, (NASA Contract NAS 2-1037) , NASA CR-53716, 1964.
Hoey, Robert G . : Horizontal Landing Techniques f o r Hypersonic Vehicles . AGARD Rep. 428, J a n . 1963.
8
I
TABLE I.- BODY COORDINATES (See f i g . 3)
x, i n .
0
.44
1.66
2.20
3.13
4.08
7.83
11.40
12.52
18.80
25.06
28 .20
77.00
77.55
81.08
84.60
88.12
91.65
93.90
98.70
105.75
112.80
119.85
130.42
141.00
151.58
162.15
169.50
~
y , in.
i k cd
I+
2 k .d U
I 16.40
c, t
7 A Ma) .d d cd .d k - 1 c,
23.83
23.97
24.25
24.32
24.18
23.90
23.62
22.84
20.94
18.54
15.37
9.52
2.04
-6.56
-15.65
-22.28
r , i n .
0
4.95
9.55
10.91
12.75
14.26
18.70
22.62
26.88
30.70
32.54
t - 1 6 cdo 0 .d .rl c, c o oa)
70.50 i 72.76
75.58
77.48
78.75
80.02
80.44
80.23
79.66
79.10
9
TABLE 11. -
Rigid
I N D E X TO FIGURES
Longi t udina l .
j o f A f t e r b o dy onf i gurat ion d a t a
n f l a t a b l e with cont r o 1s
O f f 1 Longitudinal
n f l a t a b l e Longi tudinal ~
Photos of d e f l e c t i o n
n f l a t a b l e with I Longitudinal cont r o 1s
n f l a t a b l e and r i g i d
n f l a t ab l e Longitudinal con t ro l I e f fec t iveness
n f 1 a t ab 1 e Di rec t iona l con t ro l e f f ec t iveness
Rigid
if l a t a b l e I Deployment . ~~~
99 PSf
50
5,25 10 - 33 10-33
15
30 - 70
30 ,SO 15
30,50 30 ,SO 30 ,SO
5-50
50
25
29
30,50
! 5 , 5 0
!8,48
48
!4,62
'Pi 9
p s i
5 10 15 15
15
10,15 5,10,1:
10,15 15
10,15 . . .
10
; , l o , 1 5
15
15
. .
15
~~
Vary
AP? PSI
0 . 2 . 2 . 2 .1 . 2
. 2
. 2
. 2
.1
. 2
. 2
. 2 -
. 2
. 2
. 2
. 2
V a r y -
O f f
O f f O f f O f f O f f
0
- 10 - 10 - 10 - 20 - 30
O f f ...
- 30
O f f , -10,-30
- 30
Vary
0
..
O f f
Figure
4
6
11
1 2
13
1 4
. . .
15
_ _ 16 . . . . . . -
l o
.......
.......
-=-- -=::::::::::::
A-35755 . 1 A- 35756.1
(a) Model with inflatable afterbody .
Figure 1.- Ml-L mounted in the Ames 40- by 80-Foot Wind Tunnel .
>-"0 0 ,0 H <1.) "0 ~ <1.)
lH "0 ell ;:l
M "0 (J . r-i ~ ~ 0
.r-i U H
..c: .n ~ M .,. . r-i 0>
~ <1.) C>
"'" H I «: M ~ <1.)
"0 . r-i 0 ~
::E r---. ,0 '--'
12
~.--- - --- ._---
(a) Forebody.
(b) Forebody and inflatable afterbody.
Figure 2.- Model with inflatable afterbody in various stages of assembly.
13
(c) Afterbody attached to forebody.
Cd) Model assembled with controls.
Figure 2.- Concluded.
14
- -- - -- -
153. I I
169.5-
Figure 3.- Basic model dimensions.
Note : All dimensions in inches Hinge lines sealed for inflatable afterbody and open for rigid afterbody (1/2 in. gap) Unless otherwise noted all radii=3in. Primary structure shown shaded in side view
.8
1 .- . - . . . . . _. . . . . - . . ..
. . .- .
--I(----- _- _____
.04 0 -.04 -.08 -.I2 Cm
- 2 0 2 4 L/D
Figure 4 . - Aerodynamic c h a r a c t e r i s t i c s without t he af terbody a t q = 50 p s f .
.8
.7
.6
.5
4
.3
.2
.I
0
CL
- I 0
CL
.I .2 .3 5l CD
.I
0
- I
4 8 12 16 20 24 28 32 .04 0 -.04 -.08 -.I2 -2 0 2 4 a,deg Cm L I D
(a) Api = 5 psi and Apc = 0.2 psi.
0 _I .2 .3 4 0 4 8 12 16 20 24 28 32 .04 CD a,deg
0 -.04 -.08 42-2 0 2 4 Cm L/D
(b) Ap. = 10 psi and Ap = 0.2 psi. 1 C
Figure 5.- Aerodynamic characteristics with the inflatable afterbody and without the control surfaces.
17
.8
.7
.6
.5
4
.3
.2
.I
0
CL
_ I I
0 .I
.8
.7
.6
.5
4
.3
.2
.I
0
CL
- I
.2 .3 4 0 4 8 12 16 20 24 28 32 04 0 -.04 708 a,deg Cm
(c) Api = 15 psi and Apc = 0.2 psi.
12 -2 0 2 4 L ID
0 .I .2 .3 4 0 4 8 12 16 20 24 28 32 .04 0 -.04 -.08 42 -2 0 2 4 a,deQ Cm L I D CD
(d) Api = 15 psi and q = 15 psf.
Figure 5.- Concluded.
18
.8
.7
.6
.5
4
.3
.2
.I
0
CL
-. I 0 .I .2 . 3 .4
CD
i
I
i I I 0 4 8 12 16 20 24 28 32 .04 0 -.04 -.08 42 -2 0 2 4
a,deg Cm L/D
Figure 6.- Aerodynamic characteristics with the inflatable afterbody and with inflatable control surfaces at Ap, = 0.2 psi.
6f = 0' and 6, = Oo; Ap. = 15 psi, 1
.8
.7
.6
.5
.4
.3
.2
.I
0
CL
-. I 0 .I .2 .3 4 0
CD
4 8 12 16 20 24 28 32 .04 0 -.04 -.08 -.I2 -2 a,deg Cm
(a) Apc = 0.2 psi.
0 2 4 L I D
Figure 7.- Aerodynamic characteristics with the inflatable afterbody and with inflatable control surfaces at 6f = -10" and 6, = 0 ' .
19
.8
.7
.6
.5
4
.3
.2
.I
0
CL
- I
CL
' / / j 1
' I
" 0 .I .2 .3 4 0 4 8 12 16 20 24 a d e g
.a
.7
.6
.5
4
.3
.2
.I
0
-I
28 32 .04 0 -.04 -.08 Cm
(b) Apc = 0.2 p s i and q = 15 psf.
- t E
712 -2 0 2 4
0 .I .2 .3 .4 0 4 8 12 16 20 24 28 32 .04 0 -.04 -.08 712-2 0 2 CD a,deg Clll L/D
( c ) Ap, = 0 . 1 psi.
Figure 7.- Concluded.
4
20 I
I I
.8
.7
.6
.5
4
.3
.2
.I
0
CL
_ I
"0 .I .2 .3 4 0 4 8 12 16 20 CD a,deg
Figure 8.- Aerodynamic c h a r a c t e r i s t i c s with t h e i - i n f l a t a b l e c o n t r o l s u r f a c e s a t 6f = - 2 O O
Apc = 0 . 2
.7
-6
.5
.4
.3
.2
.I
0
-.I
72
CL
t ~
i
.I .2 CD
p s i .
.3 4 0 4 8 12 16 20 24 a,deg
I
,-- I
.04 0 -.04 -.08 -.I2 -2 0 2 4 Cm L/D
m f l a t a b l e a f te rbody and with and 6 , = 0'; Api = 15 p s i ,
c + I - j 1
I
28 32 .04 0 -.04 -.08 42 -2 0 2 4 CtJl L I D
Figure 9.- Aerodynamic c h a r a c t e r i s t i c s wi th t h e i n f l a t a b l e a f te rbody and with t h e i n f l a t a b l e c o n t r o l s u r f a c e s a t 6f = -30' and 6 , = Oo; Apc = 0 . 2 p s i .
2 1
I
. - -. . . . I . . .,I 9 1 1 . ..I.. 11.1.11111111111
.8
.7
.6
.5
4
.3
.2
.I
0
CL
- I "0 . I .2 .3 4 0 4 8 12 16 20 24 28 32 .W 0 -.04 -.08 42 -2 0 2 4
a,deg Cm L/D CD
Figure 10.- Aerodynamic characteristics with the rigid afterbody and without control surfaces.
.8
.7
.6
.5
4
.3
.2
.I
0
CL
- I
.' 0 .I .2 CD
1
I i ' !
11 i .3 4 0 4 8 12 16 20 24
a,deg 28 32 .04 0 -.04 -.08 -.I2 -2 0 2
Cm L I D
Figure 11.- Comparison of aerodynamic characteristics with and without inflatable afterbody at a dynamic pressure of 50 psf.
22
~
Ii Controls off
0 4 8 12 16 20 24 .04 0 Cm
704 -.08 0 2 4 L/D
Figure 12.- Comparison of the inflatable configuration with the rigid configuration at a dynamic pressure of 25 psf and Apc = 0.2 psi.
23
24
a,deg
4 .3
12
22
Figure 13.- Model with inflatable afterbody for three angles of attack, a dynamic pressure of 29 psf and an upper flap setting of _30°; ~Pi = 15 psi, ~Pc = 0.2 psi.
Inflatable afterbody (Ap,=15psi. Apc=.2 psi) I
.5
.4
.3
.2
CL .I
0
-.I
-. 2
.2
c, .I
0
-30 -20 -10 0 -30 -20 -10 0
8t.deg
Rigid afterbody A- <
q=25psf
Figure 14.- Effect of flap deflection on the longitudinal aerodynamic characteristics.
25
Inflatable afterbody (Api = 15 psi, Apc= .2psi) Rigid afterbody
L- YL, q=48psf -
q= 28 psf q=48psf .I
CY O
-.I
.06
.04
C" .02
0
-.02
.02
0
CZ
-.02
-10 0 10 20 -10 0 10 20 -10 0 10 20 art deg &, deg Br, deg
:04
Figure 15.- Effect of rudder deflection on the lateral-directional aerodynamic characteristics.
26
T=Osec, t:.P j =Opsj, t:.Pc=Opsi T=O.7sec, t:.pj=Opsi, t:.Pc =Opsi
T= 1.3sec, t:.Pj= Opsi, t:.Pc = Opsi T=2.3sec, t:.Pj =Opsi, t:.Pc =Opsi
T=4.0sec, t:.pj=4.8psi, t:.Pc=Opsi T=IO.Osec, t:.pj =14psi, t:.Pc=·02psi
T=60.0sec, t:.pj=20psi, t:.Pc= .05psi
(a) q = 24 psf, controls off.
Figure 16.- Deployment.
27
T = Osee. t.p; = Opsi, t.pr. = Opsi T = 1 sec, t.p; = Opsi, t.pr. = .01 psi
T = I. 5 sec, t.p; = Opsi, t.Pc = .01 psi T=3.lsee, t.Pi=Opsi, t.Pc= .l1 psi
T=21.2see, t.p;=Opsi, t.Pc= .12psi
(b) q = 62 psf, flaps on.
Figure 16.- Concluded.
28 NASA - Langley,1969 - 1 A-2894
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