Powered-Lift Aerodynamics and Acoustics

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POWERED-LIFT

AERODYNAMICS AND

ACOUSTICS

Q A conference held a t LANGLEY RESEARCH CENTER Hamptsn, Virginia May 24-26, 1976 (NASA-SP-406) POWEZED-LIFT A E R O D Y N A M I C S A Y D U78--24046 P C O ' J S T I C S Q N b S A ) 502 p HC 922/AF 101 T H R U C S C L O l k N78-2407f l l n c l a s H1/02 17138 NATIONAL AERONAUTICS AND SPACE ADMINISTR~TION

NASA SP-406

POWERED-LIFT

AERODYNAMICS AND

ACOUSTICS

A conference he'd at NASA Langley Research Center, Hamptor,, Virginia, on May 24-26, 1976 P r r ~ a r e d by L ~ n g i e y Research Center 19-0 PA< E AI)MINISTRATION IY'arhh,~[on. D.C.

PREFACE This compilation consists of papers presented at a conference on Powered- Lift Aerodynamics and Acoustics held at the NASA Langley Research Center on h y 24-26, 1976. L.e presentations were made in sessions subdivided according t o subject matter as follows:

I - High-Lift Aerodynamics, I1 - High-Speed and

The purpose of the conference was to provide an in-depth review of powered- lift technology generated by in-house and NASA sponsored research over the last several years.

Papers were presented by members of NASA Centers, Universities, and Industry.

iii

CONTENTS

PREFACE. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . iii

. . . . . . . . . . . . . . . . . . 1. OVERVIEW OF POWERED-LIFT TECHNOLOGY

John P. Campbell

SESSLON I - HIGH-LIFT AERODYNAMICS

2 . UPPER-SURFACE-BLOWING FLGW-TURNING PERFORMANCE . . . . . . . . . . . . 29

William C. Sleeman, Jr., and Arthur E . Phelps I11 3. RESULTS OF STATIC TESTS OF A 1.14-SCALE MODEL OF THE BOEING YC-14

. . . . . . . . . . . . . . . . . . . . . . . . . . POWERED-LIFT SYSTEM 45

James L. Hassell, Jr.

4. SUMMARY OF LOW-SPEED AERODYNAMIC CHARACTERISTICS OF UPPER-SURFACE-

BLOWN JET-FLAP CONFIGUXATIONS . . . . . . . . . . . . . . . . . . . . . 63

Arthur E. Phelps 111, Joseph L. Johnson, Jr., and Richard J. Margason 5. APPLICATION OF POWERED-LIFT CONCEPTS FOR IMPROVED CRUISE EFFICIENCY

OF LONG-RANGE AIRCRAFT . . . . . . . . . . . . . . . . . . . . . . . . 89

Paul Z. Coe, Jr., and Paul G. Fournier 6. COMPARISON OF AERODYNAYIC THEORY AND EXPERIMENT FOR JET-FLAP WINGS . . 103 Thomas G. Gainer, Long P. Yip, and Raymond D. Vogler , i

7 . EXTERVALLY BLOWN FLAP IMPINGEMENT PARAMETER . . . . . . . . . . . . . . 119

Danny R. Hoad

I

! 1 : 8 . SOHE MEASUREMENTS OF AN EBF POWERED-LIFT WAKE 135 . . . . . . . . . . . . .

William G. Johnson, Jr. ; I ; .

i 1 4 i ! I . :

9. AERODYNAMIC CHARACTERISTICS IN GROUND PROXIMITY . . . . . . . . . . . . 145 i . L

James L. Thomas, Janes L. Hassell, Jr., and Luat T. Nguyen ! ! : , .

8 , . I .

' 1 10. DISTRIBUTED UPPER-SURFACE BLOWING CONCEPT . . . . . . . . . . . . . . . 159 . , Paul G. Fournier and Paul L. Coe, Jr. , ) . ~ , . , , < 1 , ?

SESSION II - HIGH-SPEED AND CRUISE AERODYNAMICS j , : 11. CRUISE AERODYNAMICS OF USB NACELLE/WING GEOMETRIC VARIATIONS . . . . . 165 1 1 . : I John A . Braden, John P. Hancock, and Kenneth P. Burdges

i i i

PAC r" 1NTENTlONALLY BLANK

12. EFFECTS OF NOZZLE DESIGN AND POWER ON CRUISE DRAG FOR . . . . . . . . . . . . . . . . . . . .

UPPER-SURFACE-BLOWING AIRCRAFT 183 Edward T. Meleason 13. THEORETICAL PREDICTIONS OF JET INTERACTION EFFECTS FOR USB AND . . . . . . . . . . . . . . . . . . . . . . . . . .

OWB CONFIGURATIONS 197 C. Edward Lan and James F. Campbell SESSION 111 - ACOUSTICS . . . . . . . . .

14. USB FLOW CHARACTERISTICS RELATED TO NOISE GENERATlON 213 W. H. Brown and N. N. Reddy . . . . . . . . . . . . . . . . . . . . .

15. CHARACTERISTICS OF USB NOISE 227 J. S. Gibson and N. Searle 16. ANALYTICXL DEVELOPHENTS FOR DEFINITION AND PREDICTION OF . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

USB NOISE 241 N. N. Reddy and C. K. W. Tam 17. MODELING OF UNDER-THE-WING EXTERNALLY BLOWN FLAP . . . . . . . . . . . . . . . . . . . . . . . . . .

POWERED-LIFT NOISE 263 Daniel J. McKinzie, Jr.

. . . . . . . . .

18. USB NOISE REDUCTION BY NOZZLE AND F W P MODIFICATIONS 283 Richard E. Hayden . . . . . . . . . .

19. EBF NOISE REDUCTION THROUGH NOZZLE/FLAP POS1TLc)NISG 307 T. Kadman and K. L. Chandiramani

SESSION IV - PROPULSION AERODYNAMICS AND ACOUSTICS

. . . . . . . . . . . . . . . . . . . . .

20. OVERVIEW OF THE QCSEE PROGRAM 325 Carl C. Ciepluch

. . . . . . . . . . . . 21. ACOUSTIC DESIGN OF THE QCSEE PROPULSIDI. SYSTEMS 335

Irvin J. Loeffler, Edward B. Smith, and Harry D. Sowers 22. INLET/NAcELLE/ExHAusT SYSTEM INTEGRATION FOR THE QCSEE . . . . . . . . . . . . . . . . . . . . . . . . . .

PROPULSION SYSTEMS 357 John T. Kutaey

. . . . . . . . . . . . . . 2 3 . INLET TECHNOLOGY FOR POWERED-LIFT AIRCRAFT 369

Roger W. Luidens I I

24. REVERSE-THRUST TECHNOLOGY FOR VARTABLE-PITCH FAN PROPULSION SYSTEMS . . 387

David A . Sagerser, John W. Schaefer, and Donald A . Dietrich

:

\:.%I I ! 25. ACOUSTICS AND AERODYNAMICS OF OVER-THE-WING THRVST REVERSERS . . . . . 403

i

D a l e L. S t i m p e r t a n d R o b e r t C. Ammer !

1 !

I .

!

SESSION V - AERODYNAMIC AND ACOUSTIC LOADS

26. MEASURED AND CALCULATED STEADY AERODYNAMIC LOADS O N A LARGE-SCALE

UPPER-SURFACF BLOWN MODEL . . . . . . . . . . . . . . . . . . . . . . . 415

Boyd F e r r y 111 a n d M i c h a e l R. > l c n d e n h a l l

. . . . . . . . 2'1. ACOUSTIC-LOADS RESEARCH FOR POWERED-LIFT CONFIGLlRATTONS 4 2 9

-. . , James A. S c h o e n s t e r , C o n r a d M. Willis, J a m e s C. S c h r o e d e r , a n d J o h n S . M i x s o n

. . . . . . . . . . 28. INVESTIGATIONS OF SCALING LAWS FOR JET IMPINGETlENT 4 4 5

J. B. M o r t o n , J. K . H a v i l a n d , C . D. C a t a l a n o , a n d W. W. H e r l i n g ....

SESSION ' ! I - RILL-SCALE AND FLIGHT RESEARCK

.... 29. NASA P:iRTICIPATION IN T;;Z AMST PROGRAM . . . . . . . . . . . . . . . . 4 6 5

E a r l J . Montoya a n d A l a n E. Fnve, J r .

30. USB ENVI41)NMENT ELEIISLIREMENTS BASED O N FULL-SCALE STATIC ENGINE

GROUNDTESTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 7 4

M . B. Sussrnrin, D. L. I h r k . o n e n , anti J . B. Reed

. -. 4

OVERVIEW OF POWERED-LIFT TECHNOLOGY John P. Campbell The George Washington U n i v e r s i t y , J o i n t I n s t f t u t e f o r A c o u s t i c s and F l i g h t S c i e n c e s

- 1 S U M M A R Y l i

i l

T h i s introduc:ory paper is intended t o s e t t h e s t a g e f o r t h e c o n f e r e n c e by !

reviewing p r o g r e s s t o d a t e i n t h e p o w e r e d - l i f t ' f i e l d . The concept and a p p l i c a - t i o n of powered l i f t and t h e e f f e c t s of some fundamental d e s i g n v a r i a b l e s a r e d i s c u s s e d . A b r i e f chronology of s i g n i f i c a n t developments i n t h e f i e l d is a l s o

, i

p r e s e n t e d and t h e d i r e c t i o n of r e s e a r c h e f f o r t s i n r e c e n t y e a r s is i n d i c a t e d .

!

A l l powered-lift concepts a r e i n c l u d e d , b u t emphasis is on t h e two e x t e r n a l l y blown schemes which i n v o l v e blowing e i t h e r above o r below t h e wing and which

/

a r e now b e i n g u t i l i z e d i n t h e YC-14 and YC-15 a i r p l a n e s . 'This review d e a l s i p r i m a r i l y w i t h aerodynamics and v e h i c l e d e s i g n , and only touches b r i e f l y on t h e I a r e a s of a c o u s t i c s , p r o p u l s i o n , and l o a d s .

I

INTRODUCTION It i s perhaps a p p r o p r i a t e t o s t a r t t h i s review with a b i t of h i s t o r i c a l backg-ound which i l l u s t r a t e s one of t h e f a c t o r s t h a t s p u r r e d i n t e r e s t i n powered l i f t back i n t h e 1950's. Richard E. Kuhn brought o u t t h i s p o i n t very w e l l by t h e u s e of f i g u r e 1 which is a h i s t o r y of maximum l i f t development from t h e Wright B r o t h e r s t o t h e p r e s e n t day. The upper s o l i d l i n e shows t h a t w i t h t h e i n t r o d u c t i o n of t r a i l i n g - e d g e f l a p s and w i t h t h e c o n t i n u i n g refinement and s o p h i s t i c a t i o n of t h e s e f l a p s , t h e maximum l i f t c o e f f i c i e n t C L , ~ ~ ~ o b t a i n e d i n wind-tunnel t e s t s i n c r e a s e d a t a r a p i d r a t e up u n t i l t h e 1940's b u t a t a much more modest r a t e a f t e r w a r d . Of c o u r s e , t h e v a l u e s of CL,,,, a t t a i n e d w i t h o p e r a t i o n a l a i r c r a f t lagged w e l l behind t h e wind-tunnel p r o g r e s s a t f i r s t , b u t i t l a t e r became apparent t h a t a i r p l a n e s would soon be u s i n g up most of t h e mechanical-flap h i z h - l i f t technology developed i n le winds t u n n e l . T h i s t r e n d was f o r e s e e n by r e s e a r c h e r s i n t h e e a r l y 1950's who recognized t h a t t h e c e i l i n g o b t a i n a b l e w i t h mechanical f l a p s could be bypassed by making f u l l on CL,rnax use of t h e energy of t h e t u r b o j e t p r o p u l s i o n e n g i n e s t o augment wing l i f t , a s i n d i c a t e d by t h e dashed l i n e . E x p l o r a t o r y r e s e a r c h on t h e j e t - f l a p p r i n c i p l e was t h e r e f o r e s t a r t e d i n an e f f o r t t o r e a l i z e t h i s p o t e n t i a l .

I n t h i s j e t - f l a p concept, a h i g h - v e l o c i t y j e t s h e e t is t u r n e d downward by a t r a i l i n g - e d g e f l a p and e f f e c t i v e l y i n c r e a s e s t h e chord of t h e f l a p t o produce h i g h e r l i f t . The t o t a l l i f t produced is made up of t h e t h r e e compone,lts shown t h e power-off l i f t produced by t h e wing and f l a p , t h e l i f t due t o i n f i g u r e 2 : t h r u s t d e f l e c t i o n ( t h a t i s , t h e v e r t i c a l component of t h e t h r u s t ) , and powered

c i r 1 : u l a t l o n l i f t which is t h e a d d i t i o n a l c i r c u l a t i o n l i f t induced on t h e wing ;;it\l hv t h e p r e s e n c e of t h e j e t s h e e t . The p r o p o r t i o n s of t h e t l ~ r e e corn- ! : k x , 1 L I ? + + :, ~ r t v a r y q u i t e a b i t , d e p e n d i n g on t h e t y p e o f f l a p and t h e p a r t i c u l a r pof.t, l . c . , \ - l l f t c o n c e p t used.

POWERED-LIFT CHRONOLOGY 'i k ~ u n b ~ r of d i f f e r e n t c o n c e p t s have been s t u d i e d a s i n d i c a t e d by t h e ~ ? i ~ ; ~ t ~ r ~ i l - l i f t c h r o ~ l o l o g y p r e s e n t e d i n f i g u r e 3. D a t e s a r e shown f o r t h e f i r s t : i A : . t :.irc!l conducted on a g i v e n c o n c e p t a n 3 For t h e f i r s t f l i g h t of an a i r p l a n e i n c o r p o r a t i n g t h e c o n c e p t .

The blowing b o u n d a r y - l a y e r - c o n t r o l (BLC) scheme i l l u s t r n t e d a t t h e t o p is n o t u s u a l l y c o n s i d e r e d a t r u e p o w e r e d - l i f t c o n c e p t s i n c e i t o n l y x s e s e n g i n e b l e e d a i r and hence d o e s n o t make f u l l u s e o f t h e a v a i l a b l e e n g i n e t h r u s t . I t is i n c l u d e d h e r e , however, b e c a u s e ' o f its b a s i c s i m i l a r i t y t o t h e j e t f l a p and b e c a u s e some o f t h e work o n blowing BLC p r o v i d e d u s e f u l i n f o r m a t i o n i n t h e devclopmtmt o f t h e j e t f l a p . E x p l o r a t o r y s t u d i e s o f blowing BLC were c a r r i e d o u t as c a r l y iis t h e 1 9 2 0 ' s b u t i t was n o t u n t i l t h e 1 9 4 0 ' s and 1 9 5 0 ' s t h a t s y s - t e m a t i c r e s e a r c h was c o n d r ~ c t e d t h a t l e a d t o a p p l i c a t i o n o f t h e c o n c e p t . Some o f t h e most i m p r e s s i v e work was dons on t h e Navy's F9F-5 a i r p l a n e i n t h e e a r l y 1 9 5 0 ' s under t h e d i r e c t i o c o f John A t t i n e l l o ( r e f . 1 ) . A number o f o t h e r a i r - c r a f t w i t h blowing BLC have been flown, i n c l u d i n g t h e Boeing 367-30 a i r p l a n e which was used by NASA f o r low-speed f l i g h t r e s e a r c h i n t h e e a r l y 1 9 6 0 ' s . (See The p r i n c i p l e o f t h e j e t f l a p was proposed and v e r i f i e d by Schubauer i n 1932, b u t v e r y l i t t l e a t t e n t i o n was g i v e n t o t h e c.otlct3pt u n t i l 20 y e a r s l a t e r when A t t i n e l l o ' s s t u d i e s i n t h e U n i t e d S t a t e s ( r e f . 1 ) and Davidson's s t u d i e s i n England ( r e f . 2 ) showed g r e a t promise f o r t h e j e t f l a p . T h i s work l e d t o e x t e n s i v e r e s e a r c h programs on t h e c o n c e p t i n England, F r a n c e , and t h e United S t a t e s . (For example, s e e r e f s . 1 t o 4 . ) The Hunting j e t f l a p r e s e a r c h a i r - p l a n e ( f i g . 5 ) was b u i l t i n t h e e a r l y 1 9 6 0 ' s t o s t u d y t h e f l i g h t c h a r a c t e r i s t i c s a s s o c i a t e d w i t h t h e j e t f l a p . (See r e f . 5.) U n f o r t u u n t o l y , t h e a i r p l a n e had a number o f d e f i c i e n c i e s which l i m i t e d i t s u s e f u l n e s s as a r e s e a r c h a i r c r a f t .

. , I n t h e l a t e 1 9 5 0 ' s D e H a v i l l a n d of Canada i n i t i a t e d r e s e a r c h on a v a r i a t i o n o f t h e j e t f l a p c a l l e d t h e augmentor wing. T h i s c o n c e p t i n c o r p o r a t e s a s h r o u d assembly o v e r t h e f l a p t o c r e a t e a n e j e c t o r s y s t e m which augments t h e t h r u s t o f The augmentor wing was t h e s u b j e c t of t h e n o z z l e by e n t r a i n i n g a d d i t i o n a l a i r .

a comprrhensive r e s e ~ r c h program c a r r i e d o u t j o i n t l y by NASA and t h e Canadian governmrnt s t a r t i n g i n 1965. T h i s program c u l m i n a t e d i n t h e d e s i g n and con- s t r u c t i o n o f t h e C - 8 augmentor wing r e s e a r c h a i r p l a n e by Boeing and De H a v i l l a n d .

The a i r c r a f t was f i r s t flown i n 1 9 7 2 and s i n c e th.it time h a s been (See f i g . 6 . ) (See r e f . 6 . ) u s e d i n n j o i n t NASA-Ames and Canadian f l i g h t r e s e a r c h frogram.

Both t r ~ ? augmentor wing and t h e jet f l a p proved t o bc v e r y e f f i c i e n t a e r o - d y n a m i c a l l y i n t h a t t h e y produced a large i n c r e a s e i n wing l i f t w i t h a givc?n But t b c y a r e i n t c r n o l l y blown s y s t e m s and hence s u f f e r amount of engL~le t h r u s t .

t h e disadvnntilge of r e q u i r i n g i u t e r n a l d u c t i n g which adds t o t h e w e i g h t , cotit, and c o m p l e x i t y of t h e wing s t r u c t u r c .

I n an e f f o r t t o e l i m i n a t e i n t e r n a l d u c t i n g and t o p r o v i d e much s i m p l e r p o w e r e d - l i f t s y s t e m s , NASA Langley R e s e a r c h C e n t e r s t a r t e d work i n t h e 1 9 5 0 ' s on t h e s o - c a l l e d " e x t e r n a l l y blown systems" - t h e e x t e r n a l l y blown f l a p rlsed w i t h c o n v e n t i o n a l pod-mounted e n g i n e s , and t h e uppci- s u r f a c e blown f l a p .

E x p l o r a t o r y r e s e a r c h w a s f i r s t c a r r i e d o u t on t h e e x t e r n a l l y blown f l a p i n 195b ( r e f . 7 ) ; r e s e a r c h on t h e upper s u r f a c e blown f l a p s t a r t e d a b o u t a y e a r l a t e r ( r e f . 8). I n i t i a l r e s u l t s a p p e a r e d t o be p r o m i s i n g f o r b o t h c o n c e p t s . A f a i r l y e x t e n s i v e r e s e a r c h program was c a r r i e d o u t t o d e v e l o p t h e t e c h n o l o g y f o r t h e e x t e r n a l l y blown f l a p ; b u t t h e r e were no i n d i c a t i o n s of s e r i o u s i n t ~ e r e s t by t h e i n d u s t r y i n a p p l y i n g t h e c o n c e p t u n t i l Boeing i n c o r p o r a t e d i t i n i t s p r o p o s a l Although B o e i n g ' s e n t r y d i d n o t win, t h i s show o f f o r t h e C-5 c o m p e t i t i o n .

i n t e r e s t a c c e l e r a t e d t h e r e s e a r c h 011 t h e e x t e r n a l l y blown f l a n and l e d t o an e a r l i e r b u i l d - u p o f t h e t e c h n o l o g y b a s e r e q u i r e d f o r a p p l i c a t i o n o f t h e c o n c e p t .

The c u l m i n a t i o n o f a l l t h i s r e s e a r c h is, o f c o u r s e , t h e McDL>nnell-Douglas YC-15 AMST ( f i g . 7 ) which h i s been f l y i n g s i n c e August 1975.

As p o i n t e d o u t e a r l i e r , t h e i n i t i a l r e s u l t s o b t a i n e d on t h e c o n c e p t f o r t h e upper s u r f a c e blown f l a p i n 1957 appeared t o be prom!\-ing. The aerodynamic p e r - formance was comparable w i t h t h a t o f t h e e x t e r n a l l y blown f l a p , and preli1:linary n o i s e s t u d i e s showed i t t o be a p o t e n t i a l l y q u i e t e r c o n c e p t beca:lse o f t h e s h i e l d i n g e f f e c t o f t h e wing. (See r e f . 9 . ) However, s i n c e t h e u p p e r s u r f a c e blowing arrangement i n v o l v e d a change i n e n g i n e l o c a t i o n away from t h e g e n e r a l l y a c c e p t e d u n d e r s l u n g pods and s i n c e t h e r e was a t t h a t time no s p e c i a l c o n c e r n w i t h t h e n o i s e problem, r e s e a r c h on t h e upper s u r f a c e blown f l a p was dropped a f t e r t h e i n i t i s i s t u d i e s . Research was resumed i n t h e e a r l y 70's when i t was becoming apparenL t h a t t h e e x t e r n a l l y blown f l a p might have d i f f i c u l t y meeti,lg i n c r e a s i n g l y st rirl2er r n o i s e r e q u i r e m e n t s . S i n c e t h a t t i m e , of c o u r s e , r e s e a r c h on t h e upper s u r f a c e blown f l a p h a s been c a r r i e d o u t a t a n a c c e l e r a t e d pace; t h i s r e s e a r c h l e a d t o c h e Boeing YC-14 AMST ( f i g . 8) wh ch w i l l make i t s i i r s t f l i g h t w i t h i n a few months and t o t h e NASA q u i e t s h o r t h a u l r e s e a r c h a i r c r a f t ( f i g . 9 ) which s h o u l d b e f l y i n g i n a b o u t 3 y e a r s .

As t h e c o n f e r e n c e p r o c e P d s , you w i l l n o t e t h a t t h e r e is s p e c i a l emphasis on t h e upper s u r f a c e blown f l a p , f o r t h i s is t h e c o n c e p t which h a s been r e s e a r c h e d most e x t e n s i v e l y s i n c e t h e l a s t NASA p o w e r e d - l i f t c o n f e r e n c e h e l d i n 1972, PERFORMANCE Now, l e t u s t u r n t o some g e n e r a l performance c o n s i d e r a t i o n s f o r powered- l i f t a i r c r a f t . The l a n d i n g performance w i l l be c o n s i d e r e d s i n c e i t i s g e n e r a l l y more c r i t i c a l t h a n t a k e - o f f performance f o r t h e s e a i r c r a f t . Some o f t h e f a c t o r s irivolved i n l a n d i n g - f i e l d l e n g t h a r e illustrates i n f i g u r e 10.

On t h i s p l o t o f wing l o a d i n g a g a i n s t a p p m a c h s p e e d and t h e c o r r e s p o n d i n g operational f i e l d l e n g t h , t h e r e is a f a m i l y of c u r v e s r e p r e s e n t i n g d i f f e r e n t approach l i f t c o e f f i - c i e n t s .

The band o f v a l d e s f o r 1 . 5 t o 1.8 is f o r c o n v e n t i o n a l a i r p l a n e s w i t h mechanical f l a p s .

Note t h a t t h e s e v a l u e s a r e approach l i f t c o c f i i c i e n t s which a r e c o n s i d e r a b l y lower t h a n maximum l i f t c o e f f i c i e n t s b e c a u s e of t h e v a r i o u s a n g l e - o f - a t t a c k an3 speed margins r e q u i r e d f o r s a f e t y o f c p t , r a t i o n .

The h a t c h e d a r e a r e p r e s e n t s t y p i c a l p o w e r e d - l i f t c o n d i t i o n s i n t h e higl:,rr wing l o a d i n g r a n g e and e x t e n d s from f i e l d l e n g t h s of a b o u t 609.6 m (2000 f t ) t t ~ bout 1371.6 m

(4500 f t ) . A i r c r a f t which u s e t h e s h o r t e r f i c l d l e n g t h s , 609.6 m (2000 f t ) t o a b o u t 1066.8 m (3500 f t ) , are u s u a l l y c l a s a i f 1 , s d as STOI, o r s h o r t t a k e - o f f an^ l a n d i n g a i r c r a f t ; whereas hose usl.ng t h e 1066 8- t o 1371.6-m (3500- t o 4500. . ) f i e l d l e n g t h s are termed RTOL, o r reduced t a k e - o f f and Landing a i r c r a f t . Tht approach l i f t c o e f f i c i e n t s can v a r y from v a l u e s as low as 2 f o r t h e RTOL t o v a l u e s o f 4 o r 5 f o r t h e STOL.

O f c o u r s e , lower wing l o a d i n g s c a n b e u s e d r a t h e r t h a n h i g h e r l i f t c o e f f i c i e n t s t o o b t a i n t h e s h o r t e r f i e l d l e n g t h s , b u t t h i s u s a g e c a n l e a d t o u n d e s i r a b l e r e d u c t i o n s I n c r u i s e performance and r i d e Now, c o n s i d e r t h e a d d i t i o n a l power which must b e i n s t a l l e d i n t h e a i r p l a n e t o o b t a i n powered l i f t . F i g u r e 11 shows t h e a j r p l a n e t h r u s t - w e i g h t r a t i o s r e q u i r e d t o produce c e r t a i n v a l u e s o f CL,,,, and a p p r o a c h liit c o e f f i c i e n t s f o r a n e x t e r n a l l y blown c o n c e p t . A s a n example o f a high-perfo-mance STOL c a s e , l e t u s t a k e a n a p p r o a c h o f 4 which g i v e s a l a n d i n g f i e l d l e n g t h of a b o u t CL 609.6 m (2000 i t ) w i t h a wing l o a d i n g of 3830 ~ / m ? (80 l b / f t 2 ) . The t h r u a t - weight r a t i o r e q u i r e d i n t h i s c a s e is a b o u t 0.5 o r a b o u t t w i c e t h e i n s t a l l e d t h r u s t - w e i g h t r a t i o f o r c o n v e n t i o n a l jet t r a n s p o r t s . Of c o u r s e , i f t h e l o w e r approach l i f , r e q u i r e d f o r RTOL a i r c f t is u s e d , t h e t h r u s t - w e i g h t r a t i o s r e q u i r e d a r e w ~ c h smaller. As h a s been i n d i c a t e d , t h e s e c u r v e s a r e f o r e x t e r - n a l l y blown f l , \ p s . The more e f f i c i e n t i n t e r n a l l y blown f l a p s r e q u i r e l e s s t h r u s t - w e i g h t r , a t i o , a s i n d i c a t e d i n f i g u r e 12 ( d a t a from r e f . 1 0 ) .

F i g u r e 12 stlows t h e s t a t i c t h r u s t - w e i g h t r a t i o r e q u t r e d a s a f u n c t i o n o f approach CL f o r i n t e r n ~ i l l y and e x t e r n a l l y blown f l a p s . The l o w e r t h r u s t r e q u i r e m e n t f o r t h e i n t e r n a l l y blown f l a p s is a p p a r e n t . However, i n o r d e r t o o b t a i n a m e a n i n g f u l comparison o f t h e power r e q u i r e m e n t s f o r t h e i n t e r n a l l y and e x t e r n a l l y blown f l a p s , it is n e c e s s a r y t o c o n s i d e r t h e c h a r a c t e r i s t i c s o f t h e e n g i n e s used w i t h t h e two f l a p s y s t e m s . T h i s p o i n t is i l l u s t r a t e d i n f i g u r e 1 3 by combining t h e d:lta o f f i g u r e 12 w i t h some e n g i n e i n f o r m a t i c n . The c u r v e a t t h e right: i l l u s t r a t e s t h e v a r i a t i o n w i t h e n g i n e f a n p r e s s u r e r a t i o ' t h e s t a t i c t h r u s t - w e l g h t r a t i o a v a i l a b l e w i t h a g i v e n d e s i g n c r u i s e t h r u s t . Trte e n g i n e s a p p r o p r i a t e f o r u s e w i t h e x t e r n a l l y blown f l a p s have a r e l a t i v e l y low f a n p r e s - s u r e r a t i o and, h e n c e , p r o v i d e m?~ch more s t a t i c t h r u s t t h a n t h e e n g i n e s f o r i n t e r n a l l y blown f l a p s d e s i g n e d f o r t h e some c r u i s e t h r u s t . The d a s h r i l i n e s w i t h a r r o w s i n d i c a t e t h a t t h i s d i f f e r e n c e i n e n g i n e c h a r a c t e r i s t i c s a l m o s t b a l a n c ? ~ o u t t h e d i f f e r e n c e i n f l a p e f f i c i e n c y s o t h a t t h e o v e r a l l performance, a s i n d i c a t e d by t h e a p p r o a c h CL o b t a i n e d w i t h a g i v e n c r u i s e t h r u s t , is n o t g r r ~ t l y d i f f e r e n t f o r t h e two f l a p s y s t t 3 .

Another i m p o r t a n t f a c t o r a f f e c t i n g t h e performance o f t h e e x t e r n a l l y blown I n t h e c a s e o f s y s t e m s is t h e re1 i t i o n s h i p o f t h e e n g i n e e x h a u s t t o t h e f l a p .

t h e e x t e r n a l l y blown f l a p (EBF), i t h a s been found t h a t t h e amount o f powered l i f t o b t a i n e d depends on h o ~ well t h e f l a p " c a p t u r e s " t h e e n g i n e e x h a u s t and t u r n s i t downward.

T h i s p o i n t is i l l u s t r a t e d i n f i g u r e 1 4 ( d a t a i r o n r e f . 1 1 ) which shows p o w e r e d - l i f t increment as a f u n c t i o n o f s l i p s t r e a m c a p t u r e t a t i o , z/D, where z/D is d e f i n e d b y t h e s k e t c h . The lift increment appears t o v a r y d i r e c t l y as t h e p r o p o r t i o n o f t h e slipstream c a p r . : e d ~ n d a c t u a l l y c o n t i n u e s t o i n c r e a s e LLbyond n z/D o f 1 where t h e bottom u.' t h e e n g i n e e x h a u s t would t h e o r e t i c a l l y

. . . .

c o i n c i d e w i t h t h e bottom of t h e f l a p . It h a s been found t h a t t h i s r e l a t i v e l y simple f a c t o r Z / D can satisfactorily account f o r chenges i n geometric d e s i g n f e a t u r e s such a s l o n g i t u d i n a l and v e r t i c a l p o s i t i o n o f t h e n a c e l l e , t h e i n c i - dence o f t h e n a c e l l e , and t h e r e l a t i v e s i z e of t h e f l a p and e n g i n e n o z z l e . A paper by D. R. Hoad ( r e f . 12) i n t h i s conference w i l l g i v e more i n f o r m a t i o n O,I t h i s s u b j e c t .

I n t h e c a s e of t h e upper s u r f a c e blown £ l a ? (USB), t h e r e a r e some o t h e r c r i t i c a l f a c t o r s involved i n t h e t u r n i n g of t h e j e t exhaust a s i n d i c a t e d by f i g u r e 1 5 (taken from r e f . 1 3 ) . On t h i s p l o t of e n g i n e f a n p r e s s u r e r a t i o a g a i n s t t h e r a t i o of jet t h i c k n e s s t o f l a p t u r n i n g r a d i u s , a boundary f o r good t u r n i n g is shown. The boundary i n d i c a t e s t h a t r e d u c t i o n s i n p r e s s u r e r a t i o permit t h i c k e r j e t s t o be used, b u t i t h a s been found t h a t even w i t h low-fan- p r e s s u r e - r a t i o e n g i n e s , some s p e c i a l f e a t u r e s a r e r e q u i r e d f o r s a t i s f a c t o r y t u r n i n g . These s p e c i a l f e a t u r e s i n c l u d e extreme f l a t t e n i n g of t h e exhaust nozzle, a downward d e f l e c t i o n of t h e n o z z l e , and t h e u s e of some f l o v c o n t r o l d e v i c e such a s boundary-layer c o n t r o l o r v o r t e x g e n e r a t o r s a t t h e knee of :he The YC-14 A M S T makes u s e of v o r t e x g e n e r a t o r s a l o n g w i t h a s m a l l n o z z l e f l a p .

An i l l u s t r a t i o n of t h e improvement i n d e f l e c t i o n a n g l e t o o b t a i n good t u r n i n g .

t u r n i n g o b t a i n e d w i t h rlozzle d e f l e c t i o n a n g l e is shown i n f i g u r e 16. Note t h e t f a v o r a b l e s h i f t i n t h e boundary w i t h t h e d e f l e c t e d n o z z l e . The s k e t c h e s i n ! I f i g u r e 17 ( t a k e n from r e f . 13) i l l u s t r a t e how t h e d e f l e c t e d n o z z l e f l a t t e n s t h e I j e t s h e e t t o produce b e t t e r t u r n i n g . S i n c e t h e j e t s h e e t a l s o 2preads o u t , i t covers a g r e a t e r T a r t of t h e f l a p span and r e s u l t s i n improved : f t performance.

!

It should be poinLed o u t t h a t t h e n o z z l e d e f l e c t i o n a n g l e i ' l u s t r a t e d i n f i g u r e s 16 and 17 m y be r e f e r r e d t o i n l a t e r papers i n t h e conference a s l e f l e c t o r a n g l e , kickdown a n g l e , o r n o z z l e roof ,agle. D ~ f i n i t ! ~ o n of t h e s e a n g l e s may d i f f e r i n d e t a i l b u t they a l l r e f e r t o a dow~ward d e f l e c t i o n of t h e exhaust over t h e t o p of t h e wing t o f l a t t e n t h e j e t s h e e t and make it t u r n STABILITY AND CONTROJ.

Now l e t u s t u r n from performance t o s t a b i l i t y and c o n t r o l c o n s i d e r a ~ i o n s .

/ : ; A c r i t i c a l problem i n t h i s a r e a f c r b o t h e x t e r n a l l y blown c o n c e p t s is maintain- i n g l a t e r a l t r i m w i t h an engine o u t . Of c o u r s e , an a t t e m p t is made i n t h e b a s i c I! /; d e s i g n of t h e a i r c r a f t t o minimize t h e p r o b l e n by l o c a t i n g t h e e n g i n e s a s f a r inboard on t h e wing a s p o s s i b l e ; but s p e c i a l p r o v i s i o n s a r e s t i l l r e q u i r e d t o o b t a i n l a t e r a l t r i m w i t h o u t p r o h i b i t i v e l o s s e s i n l i f t . T y p i c a l e ~ g i n e - o u t r o l l i n g monents measured on EBF and USB models ( r e f s . 1 3 and 14) a r e p r e s e n t e d i n f i g u r e 1 8 a s a f u n c t i o n of t h e engine-out l i f t l o s s . The s o l i d l i n e r e p r e - s e n t s t h e r o l l i n g moments o b t a i n e d by m u l t i p l y i n g t h e l o s s i n l i f t by t h e d i s - t a n c e o u t t o t h e dead engine ( y l b ) , whereas t h e d a t a p o i n t s s h ~ i t h e measured r o l l i n g moments, For both models, t h e f a c t t h a t t h e measured moments a r e s m a l l e r t h a n t h e calculates moments i n d i c a t e s t h a t t h e c e n t e r o f l!ft induced by an e n g i n e is somewhat inboard of t h e engine. These measured mome.ts, how- e v e r , a r e s t i l l very l a r g e and r e q u i r e s p e c i a l a t t e n t i o n on t h e p a r t of t a e i i : 1 , , I

I I 9 n e s a t i s f a c t o r y s o l u t i o n t o t h e p r o b l e m f o r t h e e x t e r n a l l y b l o b n f l a p is i l l u s t r a t e d i n f i g u r e 1 9 ( t a k e n f r o m r e f . 1 4 ) . Shown on a p l o t o f r o l l i n g - moment c o e f f i c i e n t a g a i n s t l i f t c o e f f i c i e n t are t h e b a s i c 4 - e n g i n e CL,max c o n d i t i o n , t h e e n g i n e - o u t c o n d i t i o n w i t h n o l a t e r a l tr!!., a n d t h e trimmed con- d i t i o n s o b t a i n e d w i t h m i d s p a n d i f f e r e n t i a l f J . a p s a n d s p o i l e r s . With b o t h t h e s p o i l e r s a n d f l a p s d e f l e c l e d , t h e r o l l i n g moment i s Inore t h a n a d e q u a t e f u r lateral trim; t h e r e f o r e , much o f t h e s p o i l e r e f f e c t i v e n e s s L s a v a i l a b l e f o r m a n e u v e r i n g i n r o l l .

T h i s s o l u t i o n t o t h e e n g i n e - o u t l a t e r a l trim p r o b l e m d i d n o t work f o r t h e

1 ;-

u p p e r s u r f a c e blown f l a p b e c a u s e o f a b a s i c d i f f e r e n c e i n t h e f l o w p a t t e r n s o v e r

L

t h e w i n g a n d f l a p , as i l l u s t r a t e d i n f i g u r e 20 ( t a k e n f r o ~ r e f . 1 3 ) . F o r t h e f e x t e r n a l l y blown f l a p , t h e f l o w i m p i n g e s on t h e b o t t o m s u r f a c e o f t h ~ f l . p s a n d

f

s p r e a d s o u t s p a n w i s e t h r o a g h t h e f l a p s l o t s s o t h a t t h c p o w e r e d - l i f t e f f e c t

I

e x t e n d s w e l l o u t b o a r d o f t h e e n g i n e s . F o r t h e u p p e r s u r f a c e blown f l a p , t h e !

jet e x h a u s t t e n d s t o r o l l up a n d c o r . t r a c t , a n d t h u s p u l l s t h e l o w e r v e l o c i t v f r e e - s t r e a m f l o w i n w a r d a l o n g t h e mid.;pan. The m i d s p a n f l , l p s e g m e n t is t h e r e - 1 ; .

f o r e n o t v e r y e f f e c t i v e f o r p r o v i d i n g r o l l t r i m . A much more e f f e c t i v e r o l l

!

trim f o r t h e USB c o n f i g u r ~ t i o n was f o u n d t o b e t h e u s e o f a s y m m e t r i c a l b o u n d a r y - l a y e r c o n t r o l ; t h a t is, t h e u s e o f BLC o n t h e l e a d i n g e d g e a ~ d a i l e r o n o f t h e I w i n g w i t h t h e e n g i n e o u t b u t n o t on t h ? c ) t h ? r w i n g . F i g u r e '71 shows some l a t e r a l 1 trim d a t a o b t a i n e d w i t h t h i s method ( r r f . 1 3 ) w h i c h l o o k v e r y s i m i l a r t o t h e r e s u l t s o b t a i n e d c n t h e LBF model w i t h t h e m i d s p a n d i f f e r e n t i a l f l a p and s p o l l e r .

T h i s p o i n t w i l l b e c o v e r e d i n more d e c l i l by A . E . P t l e l p s 111 a n d J . L . J o h n s o n i ' Y , ( r e f . 1 5 ) .

,

! _

A n o t h e r c r i t i c a l s t a b i l i t y and , ~ n t r o l p r o b l e m a r e a f o r p o w e r e d - l i f t a i r - , .

c r a f t L s t h e d e s i g n o f t h e h o r i z o n t a l t a i l f o r a d e q u a t e l o n g i t u d i n a l trim a n d : i I : s t a b i l i t y . L o n g i t u d i n a l trim is a p r o b l e m b e c a u s e o f t h e l a r g e nose-down p i t c h - i n g moments p r o d u c e d by p o w e r e d - l i f t f l a p s a t h i g h r h r u s t s e t t i n g s . The p r o b l e m 8 .

is i l l u s t r a t e d i n f i g u r e 2 2 ( d a t a f r o m r e f . 1 3 ) w h i c h shows t h e h o r i z o n t a l - t a i l s i z e r e q u i r e d 'c trim o u t t h e s e nose-down moments a t v a r i o u s l i f t c o e f f i c i e n t s .

t , - C u r v e s a r e shok , f o r a 27O s w e p t w i n g cind a n un:,wept w i n g h a v i n g U S B f l a p s .

I ( S i m i l a r r e s u l t s vrould b e e x p e c t e d w i t h t h e EBF c o n c e p t .)

A t a i l arm ( 1 t a i l / c ) o f f o i l r w i n g c h o r d s a n d a t a i l l i f t c o e f f i c i e n t ( i L , t a j l ) o f two h a v e b c c n $ assumed i n c a l c u l a t i n g t h e c u r v e s . I t is a p p . i r e n t t h a t v e r y l a r g e h o r i z o n t a l t a i l s a r e r e q u i r e d f o r trim a t ' t h e h i g h e r l i f t c o e r f i c i c n t s obt.-iined w i t h pow- e r e d l i f t , e s p e c i a l l y f o r t h e unswept w i n g . The trim r e q u i r e m e n t s a r e s m a l l e r i ~ r t h e s w e p t w i n y b e c a u s e w i t h t h e e n g i n e s l o c a t e d i n b o a r d , t h e p o w e r e d - l i f t l o a d s are a c t i n g f u r t h e r f o r w a r d w i t h r e s p e c t t o t h e c e n t e r o f g r a v i t y a n d t h e r e - f o r e p r o d u c e s m a l l e r nose-down moments. Even f o r t h e s w e p t w j n g , h o w e v e r , t h e t a i l s i z e s r e q u i r e d a t t h e h i g h e r l i f t c o e f f i c i e n t s a r e much l a r g e r t h a n t h e + ;irl-?.;l o f a b o u t 20 p e r c e n t u s u a l l y r e q u i r e d f o r c o n v e n t i o n a l t r a n s p o r t s . f T h i s l a r g e h o r i z o n t a l t a i l must a l s o b e p o s i t i o n e d p l o p e r l y on t h e a i r - c r a f t t o g i v e s a t i s f a c t o r y l o n g i t r l d i n a l ? t a b i l i t y , a s i l l u s t r a t t ? : , . I f i g u r e 2 3 4 1 : : ( t a k e n f r o m r e f . 1 4 ) . T h e s e p i t c h i n g - m o m e n t d a t a , f g r a powered-l i f t a p p r o a c h 1 ; ' , , c o n d i t i o n , shcw t h e u n s t a b l e t a i l - o f f c j r v e w i t h t h e l a r g e nose-down ~ n o m e n t s 1 : and two t a i l - o n c u r v e s . With t h e h i g h r e a r w a r d t a i l l o c a t i o n , t h e model is 1 : :orward i n t h e h i g h p o s i t i o n makes l o n g i t u d i n a l l y u n s t ; a b l e . Moving t h e t a i l F i g u r e 24 ( t a k e n 1 I t h e model s t a b l e , a t least o u t t o a n a n g l e o f a t t a c k o f 15".

, 1 '

I / I

i i t : + , # L - .

The trail- from r e f . 14) shows why moving t h e r a i l f o r v a r d helped t h e s t a b i l i t y .

i n g v o r t i c e s o r i g i n a t i n g a t t h e wing t i p o r outboard end of t h e f l a p move inward s o t h a t a r e a w a r d - l o c a t e d t a i l t e n d s t o move i n t o a r e g i o n of d e s t a b i l i z i n g downwash as a n g l e of a t t a c k is increased. L o c a t f c g t h e t a i l f a r t h e r f c w a r d g e t s it f a r t h e r away from t h e v o r t i c e s and i n t o a r e g i o n of l e s s d e s t a b i l i z i n g downwash. It is apparent from t h i s s k e t c h and t h e two preceding d a t a f i g u r e s t h a t s i z i n g and l o c a t i n g t h e h o r i z o n t a l t a i l f o r s a t i s f a c t o r y t r i m and s t a b i l i t y can be a c r i t i c a l d e s i g n problem For powered-lif t a i r c r a f t .

Another important s t a b i l i t y and c o n t r o l c o n s i d e r a t i c n f o r pow: r e d - l i f t air- c r a f t i s t h e s t a b i l i t y of :'-- Dutch r o l l c s c i l l a t i o n , as i l h s t r a t e d i n f i g c r e 25 (taken from r e f s . 1 3 and 1 4 ) . C a l c u l a t e d Dutch r o l l c h a r a c t e r i s t i c s f o r USB and EBF c o n f i g u r a t i o n s w i t h swept a l ~ d u n m e p t wings a r e shorn w2th boundaries taken from an AGARD p u b l i c a t i o n o a t l i n i n g STOL handling c r i t e r i a ( r e f . 1 6 ) . Tht: p l o t l e f t shows t h a t w i t h t h e swept wing, both t h e USB and EBF a i r c r a f t had on t h e u n s a t i s f a c t o r y Dutch r o l l s t a b i l i t y when t h e l i f t c o e f f i c j . e n t was i n c r e a s e d from 1.5 t o 5.0. S a t i s f a c t o r y dampiag could be o b t a i n e d by doubling t h e b a s i c r o l l . ' % * < 7 = : : and yaw damping of t h e EBF a i r c r a f t and t r i p l i n g t h e r o l l and yaw damping of t h e

.2

USB a i r c r a f t . I n c o n t r a s t , t h e p l o t on t h e r i g h t f o r t h e unswept wing shows t h a t :.<' i n c r e a s i n g t h e l i f t c o e f f i c i e n t from 1.5 t o 5.0 makes t h e Dutch r o l l s t c b i l i t y . - + I s a t i s f a c t o r y even w i t h t h e b a s i c r o l l and yaw damping.

The feet t h a t t h e unswept . .I wing looks s o good from t h e s t a n d p o i n t of Dutch r o l l , w h i l e t h e swept wing was shown t o r e q u i r e a much s m a l l e r h o r i z o n t a l t a i i f o r l o n g i t u d i n a l t r i m ( f i g . 22) ACOUSTICS AND LOWS t ..': .: I - ; f ..:4 The a r e a s of powered-lift a c o u s t i c s and i o a d s w i l l r.:w b e considered.

A good i i l u s t r a t i o n of t h e s e v e r i t y of t h e n o i s e problem f o r powered-lift STOL a i r c r a f t

1 . I' ~2

'.-.I is shown i n f i g u r e 26 (from r e f . 17) which compares t h e n o i s e requirements f o r

t:...: !

STOL and CTOL (conventional take-off and l a n d i n g ) a i r c r a f t . F i r s t , t h e b a r s a t

/-<-:j

t h e l e f t show t h e p r e s e n t and proposed F e d e r a l Aviation A d m i n i s t r a t i o n (FAA) I - : . , a s i d e l i n e n o i s e c o n s t r a i n t s (103 t o 98 EPNdB) f o r a s i d e l i n e d i s t a n c e of 0.56 !cm (0.35 mile) o r 643 m (2100 f t ) . I f t h e s e v a l u e s a r e converted t o a s i d e l i n e d i s t a n c e of 151 m (500 f t ) they become 124 and 119 EPNdB. The b a r a t t h e r i g h t shows t h a t t h s t e n t a t i v e STOL n o i s e goal f o r t h i s same 151-m (500-ft) s i d e l i n e d i s t a n c e is 95 EPNdB, which means t h e STOL must be 24 t o 29 EPNdB q u i e t e r than , conventional a i r p l a n e . T h i s s t r i n g e n t requirement. cf c o u r s e , sterns from t h e f a c t t h a t STOL a i r c t a f t a r e intended t o o p e r a t e from a i r p o r t s which a r e c l o s e r t o populated a r e a s .

Although t h e STOL is r e q u i r e d t o b e much q u i e t e r than a CTOL, it is a c t u a l l y p o t e n t i a l l y n o i s i e r because i t h a s a much h i g h e r i t i s t a l l e d t h r u s t and o p e r a t e s a t high khrust v a l u e s d u r i n g approach and 1andiv.g.

The s o l u t i o n t o t h i s problem is obviously t h e use of a very q u i e t e n g i r e ; and pronlising r e s e a r c h and develop- ment have been going on i n t h i s a r e a .

Unfortunately, t h e e x t e r n a l l y blown f l a p produces a d d i t i o n a l n o i s e which cqmpounds t h e problem, as i l l u s t r a t e d i n f i g - u r e 27 (from r e f . 18).

Noise r a d i a t i o n p a t t e r n s a r e shown f o r engine a l o n e , f o r

I

r

f l a p s r e t r a c t e d , and f o r a take-off f l a p d e f l e c t i o n . There is a small i n c r e a s e i n n o i s e l e v e l even w i t h f l a p r e t r a c t e d , and a very l a r g e i n c r e a s e when t h e f l a p s a r e extended down i n t o t h e jet e x h a u s t .

A s i n d i c a t e d e a r l i e r , it was t h e s e v e r i t y o f t h i s f l a p impingement n o i s e w i t h t h e EBF which r e s u l t e d i n renewed i n t e r e s t i n upper s u r f a c e blowing.

The b e n e f i t t o b e gained by having t h e exhaust flow above t h e wing t o t a k e advantage of t h e s h i e l d i n g e f f e c t of t h e wing is i l l u s t r a t e d i n f i g u r e 28 (from r e f . 1 8 ) which compares r o i s e r a d i a t i o n p a t t e r n s and n o i s e l e v e l s f o r EBF and USB f l a p systems w i t h a l a n d i n g f l a p d e f l e c t i o n . The p l o t a t t h e l e f t shows t h a t t h e USB produces more n o i s e above t h e wing but produces much l e s s n o i s e below t h e wing, which is, of c o u r s e , t h e important d i r e c t i o n . The p l o t a t t h e r i g h t shows t h e v a r i a t i o n of n o i s e w i t h n o z z l e exhaust v e l o c i t y f o r t h e two concepts. A sub- s t a n t i s l r e d u c t i o n i n exhaust v e l o c i t y is r e q u i r e d w i t 1 1 t h e EBF t o g i v e compa- r a b l e n o i s e l e v e l s w i t h t h e USB; t o o b t a i n t h i s lower exhaust v e l o c i t y , a n e n g i n e w i t h a lower f a n p r e s s u r e r a t i o is r e q u i r e d w i t h t h e EBF. Recent 2zvelap- ments which i n d i c a t e s o l u t i o n s t o t h e n o i s e problems of b o t h t h e s e c o n c e p t s a r e covcred i n l a t e r s e s s i o n s of t h e conference.

i

I n t h e a r e a c f aerodynamic l o a d s , one of t h e problems i n h e r e n t i n t h e e x t e r n a l l y blown concepts is t h e lame s t a t i c l o a d s produced on t h e f l a p s a s i l l u s t r a t e d i n f i g u r e 29 ( t a k e n from r e f . 1 9 ) . On t h i s p l o t of t h e spanwise v a r i a t i o n - o f f l a p normal f o r c e on t h e t h r e e . f l a p segments, t h e peak l o a d s a r e o b t a i n e d d i r e c t l y behind each engine. These r e s u l t s were o b t a i n e d on a n EBF model; b u t s i m i l a r peak v a l u e s behind t h e e n g i n e s 0ccr.r f o r USB c o n f i g u r a - t i o n s a s w i l l be s e e n i n a subsequent paper by B. P e r r y III and M. R. Mendenhall ( r e f . 20). Another l o a d s problem f o r b o t h of t h e s e cortcepts is h i g h - i n t e n s i t y f l u c t u a t i n g l o a d s which can induce high v i b r a t i o n l e v e l s and s o n i c f a t i g u e .

F i g u r e 30 ( t a k e n from r e f . 21) i l l u s t r a t e s t h e p r i n c i p a l s o u r c e s of t u r b u l e n t p r e s s u r e f l u c t u a t i o n s f o r hoth e x t e r n a l l y blown c c n c e p t s . These p r e s s u r e f l u c - t u a t i o ~ ~ s can be generated w i t h i n t h e e n g i n e by combustion, i n t h e mixing r e g i o n of t h e c o r e o r bypass exhaust j e t , o r i n t h e flow impingement r e g i o n by boundary l a y e r s o r s e p a r a t e d flow. The s i g n i f i c a n c e of t h e dynamic l o a d s induced by t h e s e p r e s s u r e f l u c t u s t i o n s is i l l u s t r a t e d i n f i g u r e 31 ( t a k e n from r e f . 22).

The sound p r e s s u r e l e v e l s of s e v e r a l s o u r c e s of a c o u s t i c l o a d i n g on a i r c r a f t a r e compared i n b a r graph form. For sound :bressure l e v e l s ;above about 130 dd, s o n i c f a t i g u e f a i l u r e s of l i g h t secondary : . . r u c t u r e s have become a problem w i t h t h e t o p f o a r s o u r c e s shown. It is t h e r e f o r e expncted t h a t blown f l a p s ( b o t h EBF and US&) w i l l a l s o be s u b j e c t t o s o n i c f a t t p u e and t h a t s p e c i a l a t t e n t i o n must be given t o t h i s problem i n t h e detk ed d e s i g n of t h e powered-lift system.

OTHER POWERED-LIFT CONCEPTS Some o t h e r powered-lift c o n c e p t s which have r e c e n t l y bren r e c e i v i n g ' a t t e n - t i o n a r c i l l u s t r a t e d i n f i g u r c 32. F i r s t , a t t h e top of t h e f i g u r e i~ t h e over- p o t e n t i a l a p p l i c a t i o n t o c o n v e n t i o n a l the-wing blowing arrangement which has subsonic t r a n s p o r t s and s u p e r s o n i c t r a n s p o r t s . T h i s c o n ~ * ~ > p t d i f f e r s f ron upper s u r f a c e blowing i n t h a t t h e engine exhaust i n c r u i s i n g f l i g h t does not touch t h e upper s u r f a c e of t h e wing. Thus, s c r u b b i n g d r a g LS avoided and i t might be

possible to position the engine so that the 2xhaust produces a favorable rather than a detrimental interference drag. For low-speed flight, tail-pipe deflectors turn the exhaust downward against the top of the wing. Research results on this concept will be given in a rubsequent paper by P. L . Coe and P. G. Fournier Another concept, illustrated at the lower left, is spanwise blowing, a technique in w ! : ' -11 a jet of air is blown out along the upper surface of the wing in a direction essentially parallel to the leading edge in order to enhance the leading-edge vortex and thereby delay vortex breakdown and wing stall to higher angles of attack. (See ref. 24.) This concept appears to be promising as a means of increasing the maneuverability of fight~r aircraft. Another means of increasing fighter maneuverabi!ity, which has also been studied recently, is the use of powered-lift maneuvering flaps s u ~ , , .is illustrated at the lower right of figure 32. Flaps of this type can provide the substantial increase in lift desired for better maneuvering capability.

CONCLUDING REMARKS In t h i ~ overview of powered-lilt technology, an attempt has been made to present in a very condensed form, an objective view of both the potential and the problems of powered lift. The papers to be presented during the remainder of the conference will complete the picture and will cover some of the latest developments in the field.

REFERENCES 2 . Davidson, I . M . : The Jet Flap. J. Roy. Aeronaut. Sac., vol. 60, no. 541, Jsn. 1956, pp. 25-50.

Preprint No. 715, S.M.F. Fund Paper, Inst. Aeronaut. Sci., Jan. 1957.

4. Malavard, L.; Poisson-Quinton, Ph.; and Jousserandot, P. (T. M. Berthoff and D. C . Hazen, transl.): Theoretical and Experimental Investigations of Circulation Control. Rep. No. 358, Dept. Aero. Eng., Princeton Univ., Assessment of 5. Harris, K. D . : The Hunting H.125 Jet-Flap Research Aircraft.

Lift Augmentation Devices, AGARD-LS-43-71, Feb. 1971.

6. Quigley; Hervey C.; and Innis, Robert C.: A Flight Investigation of the STOL Characteristics of an Augmented Jet Flap STOL Research Aircraft.

NASA TM X-62334, 1974.

7 . Campbell, John P.; and Johnson, Joseph L., Jr.: Wind-Tunnel Investigatic of an External-Flow Jet-Augmented Slotted Flap Suitable for Application to Airplanes With Pod-Mounred Jet Engines. NACA TN 3898, 1956.

8. Turner, Thomas R.; Davenport, Edwin E.; and Riebe, John M.: Low-Speed Inves- tigation of Blowing From Nacelles Mounted Inboard and on the Upper Surface of an Aspect-Ratio-7.u 35O Swept Wing With Fuselage and Various Tail Arrangements. NASA MEMO 5-1-59L, 1959.

9. Maglieri, Domenic J.; and Hubbard, Harvey h , : Preliminary Measurements of NASA the Noise Characceristics of Some Jet-Augmented-Flap Configurations.

MEMO 12-4-58L, 1959.

10. Hoau, Danny R . : Co~parison of Aerodynamic Performance of Several STOL Con- cepts. S T O ! , Technology, NASA SP-320, 1972, pp. 111-119.

11. k?s, Marshall H.: Air Force STOL Tactical Aircraft Investigation: Evalua- @reprinq 730914, Soc. Automot. Eng. , tion of Exte nally Blown Flaps.

Oct. 1973.

1.2. Hoad, Danny R . : Externally Blown Flap Impingement Parameter. Powered-Lift Aerodynamics and Acoustics, NASA SP-406, 1976. (Paper no. 7 of this compilation. ) 13. Johnson, Joseph L., Jr.; and Phelps, Arthur E., 111: Low-Speed ~erodynamics [~re~rint) 740470, Soc. Automot* Eng* 9 of the Upper-Surface Blown Jet Flap.

Apr.-May 1974.

14. Parl.ecr, Lysle P . : Stability and Control of Externally Blown Flap Configu- rutFa;zs. STOL Technology, NASA SP-320, 1972, pp, 55-69.

15. P b e l ~ . ~ , Arthur E., 111; Johnson, Joseph L., Jr.; and Margason, Richard J.: Su ! ry of Low-Speed Aerodynamic Characteristics of Uppc%r-Surf ace-i3lown J e .Flay Configurations. Powered-Lift Aerodynamics and Acoustics, NASA SP- ,O6, 1976. (Paper no. 4 of this compilation.]

16. V;ST1) Handling. I - Criteria and Discussion. AGARD Rep. No. 577, Dec. 1970.

17, R;ll{s, Raymond 2 STOL Noise Sources and Fan Noise Treatment. Aircraft E l g l ; l e Noise Red~rtion, NASA SP--711, 1972, pp. 247-258.

18. Dortc' , Robert C.; and Reshotko, Meyer: EBF Noise Tests With Engine Under-

cha Wing and Over-the-Wing Configurations. STOL Technology, NASA SP-320, 197;!, pp. 455-473.

I ! ) . Gl.et!ne-, George C.; and Perry, Boyd, 111: Aerodynamic Loads Measurements on Externally Blown Flap STOL Models. STOL Technology, NASA SP-320, 1972, pp. 121-130.

20. Pe?.ry, Boyd, 111; and Mendenhall, Michael R. : Measured and Calculated .teed.!? Aeradynnmic Loads on a Large-Scale Upper-Surface-Blown Model.

rowerod-Lf f t Aerodynamics and Acoustics, NASA SP-406, 1976. (Paper ro. 2b of this compilation.)

2i. Miz ;on, John S.; Shotlnstcr, James A. ; and Willis, Conrad M. : Fluctuating P it.h~sllres on Aircrnf t Wing and Flap Surfaces Associated With Powered-Lif t System:3. A I M . ' d p e r 75-472, Mar. 1975.

22. Lansinlt, Donald L.; Mixson, John S.; Brown, Thomas J.; and Drischler, Joseph A. : Externally Blown Flap Dynamic Loads. STOL Technology, NASA SY-3?0, 1972, pp. 131-142.

23. Coe, Paul L., Jr,; and Fournier, Paul C.: Application of Powered-Lift Concepts for Improvcj Cruise Efficiency of Long-Range Aircraft. Powered- Lift Aerodynami:~ and Acoustics, NASA SP-406, 1976. (Paper no, 5 of this compill.a':ion.

24. Campl~etl, J : ?s F.: Augmentatior~ of Vortex Lift by Spanwise Blowing.

A T A A P.~ner No. ;5-993, Aug. 1975.

-

-

-

I I I 1 1900 1920 1940 1960 1980 YEAR .

Figure 1.- Maximum l i f t history.

THRUST DEFLECTION POWER-OFF I I FT -.---- .. . ." , -. .- - .-....-- ^^ THRUST Figure 2.- C o m p o n e n t s of p o w e r e d lift.

A F I R S T f11Gl-R A F1 RST RESEARCH

BLOWING BLC -\ A BOEl NG 367-80

---- -A A HUNTING JET F M P

JET F I A P A l RPLANE AUCMENTOR WlNG WlNG EXERNALLY 0tOlVM FLAP

UPPER SURFACE (

BLOWN FLbP Figure 3.- Powered-lif t chronology.

I

Figure 4 . - l u e i n g 367-80 BLC airplane.

I

Figure 5 . - Hunting j e t - f l a p a i r p l a n e , Figure 6.- C - 8 augmantur wing z i t p l a n e .

I t I

ORIGWAC PAGE TS OF POOR QUALITY

S T A T I C T/W A V A l l A B L E W I T H

---

F U P s ~ INTERNALLY BLOWN FLAPS

I

___lt 0 2 4 6 APPROACH C L FAN PRESSURE R A T I O Figure 13.- Performance comparison including engine characteristics. Externally and internally blown flap systems.

POWERED-1 I F T 1'

EXHACST BOUND." R Y

l t //

SLIPSTREAM CAPTURE RATIO, z l D Externally blown flap.

Figure 14.- Slipstream capture.

EARLY CONFIGURATIONS

F

PRESSURE( RECENT CONFIGUHATIONS R A T I O JET THICKNESS TURNING R A D I U S Figure 1 5 . - S t a t i c t u r n i c g .

PRESSURE RATIO JET TH I C KNESS TURNING R A D I U S Figure 16.- E f f e c t of nozzle d e f l e c t i o n a n g l e .

NACELLE W i T H STRAIGHT NOZZLE NACELLE W l TH i)CFLFCTED NOZZLE Figure 17.- Flow characteristics behind nacelles.

EXTERNALLY BLOWN FlAP UPPER SURFACE BLOWN FlAP 4 -ENG I N E 2-ENGINE ENGINE - OUT ENGINE - OUT L I F T LOSS, b C L Figure 18.- Engine-out rolling moments.

*=I.,-- .

T H I S ENG 1 NE OUT

_--- ----

--

\

L I D S P A N DIFFERENTIAL FLAP AND s P o m

--

M I u S P A N . D I F F E R E N T I A L FLAP LIFT COEFFICIENT. CL Figure 19.- Engine-out l a t e r a l trim.

F z t e r n a l l y blown f l a p .

Figure 20.- Comparison of f l ~ w patterns for EBF and USB models.

T H l S ENGl NE OUT BLC A N D SPOILER

ROLLING- --

LIFT COEFFICIENT. CL Figure 2 1 . - Engine-out l a t e r a l trim.

Upper surface blown flap.

, ..

. .

UNSWEPT W I N G . .

\. . 1 .

. ~ V T ..,.,$\?.-:.

. .+..~'. ~. .

H O R I ZONTAL-TAI L AREA .- ., .,. , I' ,30

. . -

, ~ a ; . .

I .

W I N G AREA .

\

I I 1 1

L

0 2 4 b 8 W I N G LlFT COEFFICIENT Figure 2 2 . - Horizontal-tail s i z e required for trim.

= 2 .

USB flap; ~ t e i l f rn 4 ' ' I . , t a i l

PITCH ING-MOMENT H ICH REARWARD COEFFICIENT.

Cm

-

0 ,

-0.5 -

HIGH FORWARD /--+- HORIZONTAL TAIL

-1.0 -

.

-1.5 -5 0 5 10 15 M 25 ANGLE 3 F ATTACK, deg Figure 23.- Longitudinal stability.

Figure 24.- Wing v o r t e x flow.

U S B LBF CL 0 1.5 0 5.0 0 5.0 (ROLL AND Y A W D A M P I N G DOUBLED) A 5.0 (ROLL AND YAW D A M P I N G TR IPLED!

SWEPT N I NG UNSWEPT W I N G

r

r

DAMPING .2 - SATISFACTORY PARAMETER,

-

T112

L UNACCEPTABLE UNACCEPTABLE

-. 2

Id I I I A 0 -4 . 8 1 . 2 0 .4 . 8 1.2 FREQUENCY PARAMETER. wd Figure 25.- Dutch roll characteristics.

FAA REGULATI ONS EXTENDED TO 151 m (500 it) E R E C T I V E ' I 0 FAA REGULATIONS PERCEIVED NO1 SE LEVEL.

EPNdB TENTAT l VE STOL SIDELINE 2 100 ft 500 ft 500 ft D l STANCE 643 m 151 m 151 m (0.35 m i l e ) Figure 26.- CTOL and STOL noise requirements.

672 000 N (150 000 lb) aircraft.

I I A , PERCEIVED NOISE LEVEL, PNdB k'igure 27.- EBF noise radiation patterns.

30.4 m (100 f t) from noise source.

NOISE RADIATION PAllERNS NOISE LEVELS EXHAUST VELOCITY. 207 rn sec (680 ft ,'secl 270' I + EBF

* USB

ft ' sec : I 151 183 213 344 274 3@ m 'sec

! /

NOZZLE EXHAUST VELOCITY ! i ' 1 Figure 28.- Comparison of EBF and USB n o i s e .

151 m (500 ft) from n o i s e source; 30°/6~" f l a p s . , .

1 I / /

I

1 . 1

I

i

r,i.-,l-;A~TT -;-, . . -* - .. -

1 ; ----, . . . . , . -. , .- , . . - * , - , ,-

, i

1 ..I I . I j 1 . I 1 L . l " I 8

3 O r

25 -

2 0 -

FLAP

NORMAL 15 -

FORCE

-

-

0 L 1 I I I I I I I 1 I 1 . I . 2 - 3 .4 .5 .6 .7 .8 .9 S E M I S P A N P O S I T I O N Figure 29.- Spanwise variation of flap normal force.

4-engine EBF, landing flaps (15°/150/550) ; a = 1 6 ' ; C,, = 4.0.

EXTERNALLY BLOWN FLAP BYPASS TURBULENT M I X I N G REGION . J

\ EDGE FLOW

UPPER SURFACE BLOWN FLAP \ C O R E A I R T ~ R B ~ L E N T M I X I N G REGION TURBULENCE Figure 30.- Sources of fluctuating pressure on blown flaps.

FATIGUE -

W FAILURES 1:o 111:) 170 100 210 S O l r V D PRESSURE LEVtL, dB Figure 31.- Sound ptesaurc l e v ~ l s of accrustfc l o a d l n f : on a i r c r a f t structures.

OVER -THE-\IV IMG BLO\V I NG 5 UPERSnN FC TRANSPORTS SUIBSONIC TRANSPORTS POWt RED-1 If I MAN€UVER!NG F l A P S SPANllrl SE BLO\jlING TO D E l A Y STALL F i ~ u r e 32 .- Other pawpred-lif c conrepts.

- . . -. *

UPPER-SURFACE-BLOWING FLOW-TURNING PERFORMANCE William C. Sleeman, Jr.

NASA Langley Research Center Arthur E. Phelps 111 Langley Directorate, U. S. Army Air Mobility R&D Laboratory SUMMARY Jet-exhaust flow-turning characteristics were determined for systematic variations in USE (upper-surface blowing) exhaust nozzles and trailing-edge flap conf iguration variables from experimental wind-off (static) flow studies.

For conditions with parallel flow exhausting from the nozzle, jet height (as indicated by nozzle exit height) and flap radius were found to be the most important parameters relating to flow turning. Nonparallel flow from the nozzle, as obtained from an internal roof angle andlor side spread angle, had a large favorable effect on flow turning.

Comparisons made between static turning results and wind-tunnel aero- dynamic studies of identical configurations indicated that static flou- turning results can be indicative of wind-on powered-lift performance for both good and poor nozzle-flap combinations but, for aarginal designs, can lead to overly optimistic assessment of powered-lift potential.

INTRODUCTIO-i The need for systematic study of upper-surface-blowing (USB) nozzle and USB nozzle and flap design variables were undertaken to fulfill this need for basic USB des~.gn information, as well as to improve understanding of the flow ' I phenomena associated with generation of powered lift.

[ ! . .; I l i Static flow-turning studies offer a relatively simple and inexpensive ' . / I . ; means of evaluating the high-lift performance potential of a range of USB ' 1 t I .

configuration vari6bles. It is recognized that other factors, such as thrust : recovery efficiency, are important to powered-lift performance; however, with- out good flow turning, a configuration has little chance of developing .

acceptable powered-lift characteristics. Static flow-turning d a t ~ for a broad range of configurations can identify promising nozzle-flap configurations for , subsequent powered-lif t evaluation with forva speed effects in wind-tunnel I tests. Past experience generally has shown that configurations with good static flow turning also provide appreciable lift increments due to power,

' I

whereas c o n f i g u r a t i o n s t h a t showed poor flow t u r n i n g a t s t a t i c c o n d i t i o n s I a l s o had v e r y l i t t l e g a i n i n l i f t due t o power. One of t h e o b j e c , ~ v e s of t h e work presented h e r e i n was t o determine e f f e c t s of a broad range of n o z z l e - f l a p c o n f i g u r a t i o n s on s t a t i c flow-turning a n g l e s . The second o b j e c t i v e was t o determine how w e l l t h e wind-on l i f t c h a r a c t e r i s t i c s of USB c o n f i g u r a t i o n s could be i n f e r r e d from s t a t i c f l o w - t u n i n g c h a r a c t e r i s t i c s .

S t a t i c tests were conducted on a USB nozzle-flap model t h a t provided f o r s y s t e m a t i c v a r i a t i o n s of b a s i c d e s i g n parameters over a nozzle-pressure- r a t i o range from about 1 t o 3. E i g h t v a l u e s , o f n o z z l e h e i g h t , seven v a l u e s of f l a p r a d i u s , and f i v e v a l u e s of run l e n g t h ahead of t h e f l a p were i n v e s t i - g a t e d f o r p a r a l l e l flow from t h e exhaust nozzle. The b a s i c n o z z l e was modified i n t e r n a l l y t o provide f o u r i n t e r n a l roof a n g l e s and two s i d e s p r e a d a n g l e s i n a d d i t i o n t o t h e b a s i c p a r a l l e l flow (0') a n g l e s . Both s t a t i c d a t a and wind-on tests i n t h e L m g l e y V/STOL wind t u n n e l of a few complete models w i t h r e c t a n g u l a r exhaust n o z z l e s and w i t h full.-span leading-edge blowing pro- vided comparative information f o r a s s e s s i n g t h e a p p l i c a b i l i t y of s t a t i c flow- t u r n i n g r e s u l t s .

SYMBOLS L i f t

1lft p n c = C l e l a n t -

T h r u s t

t h r u s t c o e f f i c i e n t , -

q s H n o z z l e h e i g h t , h e i g h t of n o z z l e roof above nozzle f l o o r a t e x i t

I I I

9 t e s t dynamic p r e s s u r e : I ; f l a p r a d i u s , e f i ~ c t i v e t u r n i n g r a d i u s of USB f l a p a t t a n g e n t p o i n t t o upper s u r f a c e of wing a i r f o i l reference, w L . 6 a r e a a n g l e of a t t a c k of wing chord l i n e f l a p d e f l e c t i o n , a n g l e of f l a p chord l i n e a t t r a i l i n g edge w i t h r e s p e c t t o wing chord l i n e USB upper-surface blowing Notation: Run l e n g t h l e n g t h of s t r a i g h t flow run downstream of nozzle e x i t t o beginnicg of f l a p c u r v a t u r e

Flow-turning angle e f f e c t i v e s t a t i c turning angle of j e t flow a f t e r f o r c e measurements, Axial force Nozzle pressure r a t i o r a t i o of t o t a l pressure i n exhaust nozzle t o ambient t o t a l pressure i n t e r n a l angle of nozzle roof with respect t o Nozzie roof angle nozzle f l o o r Nozzle spread angle i n t e r n a l spread angle of nozzle s i d e s measured from nozzle center l i n e (slanted t o spread exhaust flow . l a t e r a l l y over wing and f l a p ) STATIC TESTS OF NOZZLE-FLAP VARIABLES Model Description A photograph of t h e s t a t i c t e s t model f.s shown i n f i g u r e 1 with h a l f of ir-supply hookup.

h t ranging from or a l l smaller hat formed t h e The width of t h e exhaust nozzle remained constant while other nozzle parameters such as spread angle and nozzle height were varied. The span of the f l a p was s u f f i c i e n t l y l a r g e t o cor.rain the flow f o r a l l conditions of A l l t e s t s of the model were con- t h e exhaust nozzle i n v e s t i g a t e d ( f i g . 2 ) .

ducted over a range of nozzle pressure r a t i o s which varied from about 1.1 t o 3.1 f o r nozzle heights equal t o o r l e s s than 2.54 c m (nozzle aspect r a t i o of 7) ; f o r the l a r g e s t nozzle height of 5.08 cm (nozzle aspect r a t i o of 3.5) ,

t h e maximum p r e s s u r e r a t i o t h a t could be o b t a i n e d was l i m i t e d t o about 1.6 because of mass-flow and t o t a l - p r e s s u r e l i m i t a t i o n s of t h e a i r - s u p p l y system.

E f f e c t of P r e s s u r e R a t i o on Flow Turning The f l a p and n o z z l e v a r i a b l e s i n v e s t i g a t e d i n t h e p r e s e n t s t u d y provided over 300 d i f f e r e n t c o n f i g u r a t i o n s which were t e s t e d o v e r a range of n o z z l e T h i s paper p r e s e n t s o n l y s e l e c t e d p o r t i o n s of t h e v a s t p r e s s u r e r a t i o s .

amount of d a t a o b t a i n e d , t o i l l u s t r a t e t h e p r i n c i p a l f i n d i n g s and t o summarize t h e r e s u l t s .

Data showing e f f e c t s of n o z z l e p r e s s u r e r a t i o on flow t u r n i n g for v a r i c u s F l a p r a d i u s , f l a p t u r n i n g n o z z l e a s p e c t r a t i o s a r e p r e s e n t e d i n f i g u r e 3.

The j e t flow t u r n i n g a s a n g l e (90°), and run l e n g t h were h e l d c o n s t a n t .

i l l u s t r a t e d i n f i g u r e 3 is t h e a n g l e of t h e j e t flow a f t e r i t p a s s e s over t h e wing and f l a p . The t e s t r e s u l t s show t u r n i n g a n g l e s t h a t extended t o about so0, which i n d i c a t e s t h a t t h e j e t flow dfd n o t adhere t o t.he 90' f l a p a l l a l o n g t h e f l a p b u t s e p a r a t e d a t some p o i n t ahead of t h e t r a i l i n g edge.

The n o z z l e a s p e c t r a t i o was changed by v a r y i n g t h e j e t e x i t h e i g h t , and a d e c r e a s e in t h e n o z z l e a s p e c t r a t i o was accompanied by an i n c r e a s e in the h e i g h t and a t t e n d a n t flow t h i c k n e s s . The r e s u l t s of f;gure 3 shoy an ex- pected r e d u c t i o n i n flow t u r n i n g as t h e j e t thickened (decreased n o z z l e a s p e c t Also, a s t h e j e t h e i g h t i n c r e a s e d , c o n d i t i o n s were reached where t h e r a t i o ) .

flow :auld no longer n e g o t i a t e t h e t u r n over t h e f l a p a t h i g h e r p r e s s u r e r a t i o s and t h e a b r u p t l o s s i n £ l o b t u r n i n g sliown f o r sorce cur\lcs i n d i c a t e s sudden detachmer.t of t h e flow. P a s t e x p e r i e n c e h a s shown tb,qt such s u d d e ; ~ detachment I 1 !

can occur i f t h e c o r n e r is t o o s h a r p ( s m a l l t u r n i n g r a d i u s ) , t h e j e t is t o o I I , t h i c k , o r t h e p r e s s u r e r a t i o is t o o high.

i l l

1 i i : E f f e c t s of p r e s s u r e r a t i o on flow t u r n i n g f o r a range of f l a p r ~ d i . 2 ~ I r e shown i n f i g u r e 4 , where f l a p r a d i u s is expressed nondimensionally a s a f u n c t i o n 1 1 , .

of n o z z l e h e i g h t . For t h e s e t e s t s , t h e n o z z l e a s p e c t r a t i o (and n o z z l e h e i g h t ) was held c o n s t a n t , a s was t h e f l a p a n g l e of 90°. P a s t e x p e r i e n c e would I 8 l e a d one t o expect i n c r e n s e s i n flow t u r n i n g w i t h i n c r e a s i n g t u r n i n g r a d i u s ( r e f . 1 ) ; however, t h e d a t a of f i g u r e 4 show p r o g r e d b i v ~ d e c r e a s e s i n t u r n i n g . .

w i t h i n c r e a s i n g r a d i u s a t p r e s s u r e r ~ t i o s up t o 2.2. A s mentioned p r e v i o u s l y ,

j

I I the s m a l l e s t r a d i u s was t o o s h a r p f o r flow t u r n i n g a t h i g h p r e s s u r e r a t i o s , A d d i t i o n a l and a b r c p t detachment was shown f o r p r e s s u r e r a t i o s above 2.2.

d e t a i l s on e f f e c t s of f l a p r a d i u s , p a r t i c u l a r l y i n t h e lower r a n g e , w i l l be discussed l a t e r .

The t e s t r e s u l t s over a range o i p r e s s u r e r a t i o p r e s e n t e d i n f i g u r e s 3 1 1 I ;

j I 1 -

and 4 a r e f a i r l y t y p i c a l of tllc n a t u r e of t h e c h a r a c t e r i s t i c s o b t a i n e d f o r a I ! - wide range of n o z z l e s and f l a p geometry, and most of t h e d a t a showed only minor v a r i a t i o n s i n t u r n i n g ang1.e w i t h p r e s s u r e rat!.o a t low and moderate p r e s s u r e r a t i o s .

The t e s t r e s u l t s a t low v a l u e s (c1.5) of n o z z l e p r e s s u r e

r a t i o were found t o b e i n d i c a t i v e of t h e maximum flow t u r n i n g t o b e expected from a given c o n f i g u r a t i o n . A n o z z l e p r e s s u r e r a t i o of 1.4 h a s been s e l e c t e d a s r e p r e s e n t a t i v e of a q u i e t , high-bypass-ratio p r o p u l s i o n system m d w i l l b e u t i l i z e d f o r t h e remaining d i s c u s s i o n of t h e s t a t i c tetlts of n o z z i e - t l a p v a r i a b l e s .

E f f e c t of Flap Radius on Fl.ow Turning Data which show e f f e c t s of f l a p r a d i u s on flow t u r i . 1 g f o r s m a l l i n c r e - .mts over a broad r n n g e of n o z z l e a s p e c t r a t i o s a r e p r e s e n t e d i n f i g u r e 5.

These r e s u l t s a r e f o r a n o z z l e p r e s s u r e r a t i o o I 1.4 and a c o n s t a n t run The f l a p r a d i u s is nondimensionalized by t h e n o z z l e h e i g h t .

length.

While r a d i u s - h e i g h t r a t i o s up t o 32 were i n c l u d e d i r t h e r e s u l t s of f i g u r e 4 , t h e r e s u l t s of f i g u r e 5 a r e concerned only w i t k t h e lower end of t h e r a d i u s range ( t o v a l u e s of about 8).

The r e s u l t s of f i g u r e 5 show t h e expected i n c r e a s e i n f l ~ w t c m i n g k i t h i n c r e a s i n g r a d i ~ s a t very low v a l u e s of radius-height r a t i o , b u t a breakaway p o i n t is e v i d e n t beyond which l i t t l e o r no i n c r e a s e iil t ~ . r n i n g c c c u r s as t h e r a d i u s inci-eases ( f o r each n o z z l e a s p e c t r a t i o ) . Each n o z z l e a s p e c t r a t i o had i t s own breakaway p o i n t e x c e p t p o s s i b l y f o r t h e asper:t-ratio-28 n o z z l e , f o r which d a t a were l a c k i n g . l h e s e t e s t r e s u l t s a l s o s b ~ w t h a t t h e r e i s some l i m i t t o t h e flow-turning a n g l e t h a t c ~ l l be o b t a i n e d w i t h a p a r t i c u l a r upper p a r a l l e l - f low nozzle'/ f l a p combination.

The d a t a of figur,? 5 were g e n e r a l l y r e p r e s e n t a t i v e of d a t a o b t a i n e d a t o t h e r p r e s s u r e r a t i o s and run l e n g t h s . I n c r e a s i n g t h e n o z z l e p r e s s u r e r a t i o from t h e 1.4 v a l u e used f o r f i g u r e 5 t o a v a l u e of 2.0 callsed only a s l i g h t r e d u c t i o n i n flow t u r n i n g , w h i l e v a r i a t i o n s i n flow t u r n i n g w i t h run l e n g t h s g r e a t e r t h v l z e r c werr r e l a t i v e l y s m a l l . Most of t h e dat.a a v a i l a b l e from t h i s i n v e s ~ i g a t i o n weze used t o develop t h e Following f i g u r e ( f i g . 6 ) which accounts f o r second-order s f f e c t s by shaded a r e a s whicli r e p l a c e d t h e d i s c r e t e l i n e s of f i g u r e 5. Breakaway p o i n t s f o r s e v e r a l n o z z l e a s p e c t r a t i o s a r e i n d i c a t e d by t h e numbers shown on t h e r i s i n g shaded band i n f i g u r e 6 . T h i s ;ing shart.;-4 band s e p a r a t e s t h e r e g i o n where flow t u r n i n g i s a v a i l a b l e from .

, , ~ e r e g i o n viis:r-6 flow t u r n i n g is no'- a v a i l a b l e with p a r a l l e l fl3w n o z z l e s and The r i s i n g b a l d e x t e n d s t o about 60' flow t u r n i n g f o r a f l a p r a d i u s 90' f l a p s .

r a t i o of around 4 , b u t t h i s good t u r n i n g is o b t a i n e d only w i t h a v e r y l a r g e n o z z l e a s p e c t r a t i o .

/ Experience with USB n o z z l e - f l a p c o n f i g u r a t i o n s i n o t h e r s t u d i e s h a s I I demonstrated t h a t f low-turning a n g l e s e q u a l t o c r b e t t e r than t h e 60 shown i n f i g u r e 6 have been o b t a i n e d on models w i t h r e l a t i v e l y l o w - a s p e c t - r a t i o 1 ! 1 ; : n o z z l e s . The q u e s t i o n n a t u r a l l y a r i s e s a s t o why t h e s e low-aspect-razio I ! : < .

n o z z l e s were a b l e t o provide such h i g h t u r n i n g . Since t h e h i g h - a s p e c t - r a t i o I I n o z z l e of t h e p r e s e n t s t u d y p r o v i d e s a very t h i n j e t , t h e e f f e c t i v e j e t h e i g h t I ' i a t t h e s t a r t of t u r n i n g is perhaps of more fundamental importance than n o z z l e . .

' .!

i ' l r l i i . ) . .. .

!

, . . - * . ' . , . ' % . . . I ' , , , * % . .,.. 4 - < . - - I .

h e i g h t . L o r a s p e c t - r a t i o n o z z l e s c a n turn f l o w i f t h e r e is some means of S e v e r a l means f o r c a u s i n g t h e e x i t f l o w t o t h i n n i n g and s p r e a d i n g t h e flow.

t h i n a f t e r l e a v i n g t h e n o z z l e habe been used s u c c e s s f u l l y and i c c l u d e t h e u s e of e x t e r n a l d e f l e c t o r s , a l a r g e i n t e r n a l r o o f a n g l e , and a p p r e c i a b l e n o z z l e s i d e f l a r e .

E E f e c t s o f Nozzle I n t e r n a l Angles The p r e s e n t n o z z l e - f l ~ p s t u d y i n c l u d e d s y s t e m a t i c v a r i a t i o n s i n i . * t e r n a l roof a n g l e and s i d e s p r e a d a n g l e a s shown s c h e m a t i c a l l y i n t h e s k e t c h e s i n f i g u r e 7. F l a p r a d i u s and r u n l e n g t h were varied-blAL t h e n o z z l e a s p e c t r a t i o of 7.0 was h e l d c o n s t a n t . T y p i c a l t u r r i n g - a n g l e rest-ts a r e a l s c p r ~ s e n t e d i n f i g u r e 7 a s a f u n c t i o n of i n t e r n a l roof a n a l e f o r t h r e e v a l u e s of s p r e a d a n g l e .

These t e s t r e s u l t s a r e f o r c n o z z l e a s p e c t r a t i o of 7 ard J r a d i u s , - h c i g h r r a t i o of 4. For p a r a l l e l e x i t f l o w ( s e e f i g . 6 ) . t h e f l o w t u r n i n g a t ttiese v a l u e s of a s p e c t r a t i o and r a d i u s would b e 25' t o 30°, which i s c o n s i s t e n t w i t h t h e f l o w t c r n i n g shown i n f i g u r e 7 a t 0' r o o f a n g l e and 0" s p r e a d a n g l e s . . + ! : t a t * - ment o f 60° t u r n i n g can b e o b t a i n e d e i t h e r by i n c r e a s i n g o n i y t h e roof s n g l e t o 30' o r by u s i n g 20' of s p r ~ a d a n g l e and O0 roof a n g l e & f i g . 7 ) . Combina- : i ~ n s o f roof a n g l e and s p r a n g l e can p r o k i d e up t a 80 of f l o w t u r n i n g .

The t r a d e - o f f s i n roof and s p r e a d a n g l e , a s i l l u s t r a t e d i n f i g u r e , , p r o v i d e t h e d e s i g n e r w i t h some freedom of s e l e c t i o c t o mi-lmiee probl.cms s u c h a s h i g h c r u i s e d r a g a s s o c i a t e d w i t h h i g h b o a t t a i ' . / r o o f a n g l e s .

dIE\'D-T11NNEI. TESTS OF CS)E!PLETE 1'SB \!t;DELS The p r e c e d i n g d i s c u s s i o n h a s been concerned w i t h w i a u - a . ~ - - a t i , t u r n i n g and d e s i g n v a r i a b l e s t h a t i n f l u e n c e f l o w t u r n i n g f o r r e c t a n g u l a r n o z i l e s . 7t1e second p a r t of t h i s paper d e a l s w i t h forward speed e f f e c t s f o r complste Ca.3 model c o n f i g u r a t i o n s i n wind-tunnel t e s t s . Emphasis is g i v e n t o t h e d e t e m i n a - t i o n of how w e l l t h e wind-on l i f t c h a r a c t e r i s t i c : : can be i n f e r r e d i r o n t!;* s t a t i c f low-turning c h a r a c t e r i s t i c s .

Four-Engine Elodel With Radius F l a p s A photograph of t h e f o u r - e n g i n e LlSB c o n f i g u r a t i o n i n t h e L ~ n g l e y V/STOL t u n n e l is shown i n f i g c ~ r e 8. The wing h.ld s u p e r c r i t i c a l a i r f o i i s e c t i o n s w i t h a maximum t h i c k n e s s af 9 . 3 p e r c e n t c h o r d , a nominal quarter-clir-rd sweep iillgle of 30°, an a s p e c t r a t i o of 7 . 4 8 , and a t a p e r r a t i o o f 0.147. T e s t r e s u l t s a d e t a F l e d d e s c r i p t i o n of t h i J model a r e g i v e n i n r e f e r e n c e 2 . T h i s r l > d e l p r o v i d e s a v e r y good t i e - i n w i t h t h e n o z z l e - f l a p s t u d y j u s t d i s c u s s e d inasmuch a s i t used a s p e c t - r a t i o - 6 r e c t a n g u l a r e x h a u s t n o z z l e s and a 90° r,idius i i a p .

.It shoi*,ld b e n o t e d t h a t b o t h t h e chord and a r e a of t h e f l a p decre,isc?d . i s t h e d e f l e c t i o n d e c r e a s e d from 90"; t h e a r c l e n g t h was p r o : ~ o r t i ~ ~ n a l t o f l a p d e i l e c - t i o n a n g l e ( i . e . , t h e a r c l e n g t h o t t h e 450 f l a p was h a l f o f t h a t f o r t h e 90' f l a p ) .

i i / i 7 ..:

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0 . .

Both s t a t i c flow-turning and companion forward speed test r e s u l t s a t 0 , .

, angle of a t t a c k a r e given i n f i g u r e 9 f o r a range of f l a p d e f l e c t i o n s from S t a t i c t e s t r e s u l t s p r e s e n t e d i n f i g u r e 9 show c o n s i s t e n t in- 45O t o 90°.

c r e a s e s i n flow t u r n i n g w i t h i n c r e a s i n g f l a p d e f l e c t i o n , and t h e t u r n i n g angle:: were e s s e n t i a l l y i n v a r i a n t w i t h t h r u s t . The l e v e l of flow t u r n i n ; w a s f a i r l y good; t h a t is, t h e nozzl2-flap c o n f i g u r a t i c n turned t h e flow about two-thirds of t h e f l a p daf l e c t i o n (60' t u r n i n g f o r 90' f l a p ) . C h a r a c t e r i s t i c s of t h e wind-on l i f t t h a t could b e i n f e r r e d from t h e s e f a i r l y good s t a t i c flow- t u r n i n g d a t a a r e t h a t l i f t should i n c r e a s e i n a s a o o t h and s t e a d y manner w i t h i n c r e a s e d t h r u s t , and t h a t l i f t should i n c r e a s e f o r a given t h r u s t , proportion- a t e t o changes i n f l a p d e f l e c t i o n . These c h a r a c t e r i s t i c s a r e c e r t a i n l y e v i d e n t i n t h e wind-on d a t a of f i g u r e 9 , and t h e s t a t i c t u r a i n g d a t a can be considered i n d i c a t i v e of t h e wind-on l i f t c h a r a c t e r i s t i c s t o be ex2ected.

Four-Engine Model With Modtfied S l o t t e d F l a p R e p r e s e n t a t i v e test r e s z l t s on t h e same b a s i c model j u s t d i s c u s s e d b u t w i t h d i f f e r e n t h i g h - l i f t f l a p s a r e given in f i g u r e 10. The modified f l a p system is shown a t t h r e e d i f f e r e n t d e f l e c t i o n s i n t h e s k e t c h e s a t t h e t o p of f i g u r e 10. The f l a p wzs o r i g i n a l l y a d o u b l e - s l o t t e d f l a p f o r a n e x t e r n a l l y blown f l a p n2del; t h e f l a p was modified f o r upper-surface blowing by f i l l i n g i n t h e gaps between f l a p elements. The r e s u l t i n g flap-radius-nozzle-height r a t i o s shown i n f i g u r e 10 a r e judged t o be t h e approximate e f f e c t i v e v a l u e s .

The s t a t i c flow t u r n i n g a t low and moderate t h r u s t appears t o b e f a c r l y good, i n t h a t t h e flow a n g l e was about e q u a l t o t h e f l a p d e f l e c t i o n and was r e l a t L v e l y i n v a r i a n t w i t h t h r u s t except f o r t h e f i r s t few d a t a p o i n t s a t t h e lower Flap angles. I n t h e middle-to-high t h r u s t range, however, a b r u p t de- tachment of t h e flow occurred f o r t h e h i g h e s t f l a p d e f l e c t i o n , a s evidenced by t h e l a r g e ~ e c r e a s e i n flow turning. It could be i n f e r r e d from t h e s e s t a t i c r e s u l t s t h a t t h e wind-an l i f t a t low t h r u s t would vary i n p r o p o r t i o n t o t h e f l a p d e f l e c t i o n b u t a t high t h r u s t would s u f f e r a l o s s i n l i f t i n going t o t h e h i g h e s t f l a p d e f l e c t i o n . The a c t u a l wind--on l i f t d a t a , however, show t h a t l i f t c o e f f i c i e n t s f o r t z e h i g h e s t f l a p d e f l e c t i o n were always lower than those f o r t h e lower f l a p d e f l e c t i o n s . The s t a t i c d a t a t h e r e f o r e a r e o v e r l y o p t i m i s t i c and n o t d i r e c t l y i n d i c a t i v e of t h e wind-on c h a r a c t e r i s t i c s obtained .

Previous experienc-. w i t h upper-surface-blown c o n f i g u r a t i o n s hab i a d i c a t e d t h a t poor wind-on powei-ed-lift c h a r a c t e r i s t i c s almost always accompany poor s t e t i c flow-turning c h a r a c t e r i s t i c s . Likewise, gogd i n v a r i a n t s t a t i c f l a w t u r n i n g g e n e r a l l y can be taken t o i n d i c a t e correspondingly good wind-on l i f t c h a r a c t e r i s t i c s . Considerable u n c e r t a i n t y can e x i s t between t h e s e extremes of very good and poor s t a t i c t u r n i n g ; however, where a p p r e c i a b l e v a r i a t i o n s i n flow t u r n i n g w i t h s t a t i c t h r u s t occur, anomalies i n t h e wind-on l i f t can The sudden detachment of t h e flow shown f o r t h e 65' g e n e r a l l y b e expected.

f l a p d e f l e c t i o n i n f i g u r e 10 s u g g e s t s t h a t t h e flow a t low t h r u s t was only

marginally attached, and the a d d i t i o n of more t h r ~ s t and/or addition of t h e free-stream v e l o c i t y caused the j e t flow t o detach. I n t h i s l a s t example, s t a t i c flow-turning r e s u l t s gave some i n d i c a t i o n t h a t t h e j e t flow was only marginally attached f o r some configurations and, with forward speed, would The question then a r i s e s a s t o t h e likelihood of a be expected t o detach.

s i t u a t i o n &ere the s t a t i c flow-turning d a t a gave no i n d i c a t i o n of marginal forward speed e f f e c t s would cause the flow t o de- attachment, while i n f a c t , tach. The discuasion i n the next s e c t i o n d e a l s with such a s i t u a t i o n .

- 1 , , .

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! 1 . ' USB Model With Full-Span Leading-Edge Blowing 1 i i . . A A photograph of t h e leading-edge-blowing model is shown i n f i g u r e 11. The model had blowing over the upper surface of the wing from a small full-span The model had a double-hinge p l a i n s l o t located near the wing leading edge.

A sketch of the model f l a p with a v e r j small radius on the forward element.

a i r f o i l , showing the blowing s l o t a t 19 percent chord and the deflected f l a p , f l a p allowed a is gixen a t the top of f i g u r e 12. The use of a double-hinged matrix of f l a p combinations t o be investigated by s e t t i n g the r e a r f l a p and varying the d e f l e c t i o n of t h e forward elercent over a range of angles. I n order t o cover more model configurations chan could be e a s i l y handled by the d a t a presentation used with the previous f i g u r e s , the d a t a of f i g u r e 12 a r e presented a s a function of t o t a l f l a p deflection.

.I -

Flow-turning angles and wind-on l i f t c o e f f i c i e n t s a r e presented i n f i g u r e 1 2 a s functions of t o t a l f l a p d e f l e c t i o n f o r Cp = 2.0 wizh the wind on and f o r @ + - t i c thrubt corresponding t o Cp = 2.0 with the wind o f f . None of the basic wind-off flow-turning data showed unusual v a r i a t i o n s with s t a t i c S t a t i c t e s t r e s u l t s of f i g u r e 12 show t h a t flow ~ u r n i n g increased a s t h r u s t .

the t o t a l f l a p d e f l e c t i o n increased, and turning angles up t o 85' were obtained a t t h e highest f l a p deflection. The wind-on l i f t data of f i g u r e 12 show in- creases i n l i f t with f l a p d e f l e c t i o n up t o a point and then a decrease i n l i f t .

It is i n t e r e s t i n g t o observe t h a t t h e l i f t peak always occurred a t 45' d e f l e c t i o n of t h e f r o n t f l a p (45' f r o n t + 0' r e a r = 4S0, 45' !rant + 15' r e a r = 60°, and 45O f r o n t + 3 0 ' r e a r = 75'). These r e s u l t s i n d i c a t e t h a t good s t a t i c flow turning may not always b e i n u ~ c a t i v e of correspondingly good power-on l i f t c h a r a c t e r i s t i c s . An assessment of t h e l i f t p o t e n t i a l based on t h e s t a t i c turn- , ; . .I ; . ! ...! , .

ing d a t a f o r t h i s leading-edge-blowing model would be o p t i m i s t i c f o r flow- turning angi s i n excess of about 60'. a - .

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. , .!

I ..:; - j 1 ;.; CONCLUDING REMARKS , .

S t a t i c t e s t r e s u l t s obtained i n an i n v e s t i g a t i o n of parallel-exit-flow USE nozzle and f l a p geometric v a r i a t i o n s over a range of nozzle pressure r a t i o from about 1 t o 3 showed very l i t t l e e f f e c t of pressure r a t i o on s t a t i c flow turning f o r configurations with well-established flow.

Test r e s u l t s obtained a t low preasure r a t i o s were i n d i c a t i v e of the maximum flow-turnicg performance t o be expected.

Some gains i n flow turning were even realized by the use of a

emall (2.54 cm) run length ahead of the f l a p , but f u r t h e r i n c r e a s e s i n run length produced r e l a t i v e l y amall changes i n turning. Flow-turning angles in- creased when t h e j 2 t height decreased and increased with f l a p r a d i u s up t o a p o i n t , beyond which l i t t l e o r no i n c r e a s e w a s evident.

The j e t h e i g h t , a s indicated by t h e nozzle e x i t h e i g h t , and t h e f l a p radius were t h e most important parameters r e l a t i n g t o flow turning f o r p a r a l l e l flow from t h e nozzle. Non- p a r a l l e l flow from t h e nozzle, a s obtained from an i n t e r n a l roof angle and/or s i d e spread angles, had a l a r g e favorable e f f e c t on flow-turning angles.

S t a t i c tests and wind-or? tests of complete USB models i n a wind tunnel indicated t h a t t h e s t a t i c flow-turning r e s u l t s can be i n d i c a t i v e of wind-on powered-lift performance f o r both good and poor conditions. For marginal conditions, t h e s t a t i c r e s u l t s can lead t o an overly o p t i m i s t i c assessment of wind-on powered-lift p o t e n t i a l .

REFERENCES 1. Phelps, Arthur E., 111: Aerodynamics of the Upper Surface Blown Flap.

STOL Technology, N A S A SP-320, 1972, pp. 97-110.

2. Sleeman, William C., Jr.; and Hohlweg, William C.: Low-Speed Wind-Tunnel I n v e s t i g a t i o n of a Four-Engine Upper Surface Blown Model Having a Swept Wing and Rectangular and D-Shaped Exhaust Nozzles. N A S A T N D-8061, 1975.

Figlre 1 . - Model used in static t e s t s of USB nozzle and f l a p vsriables.

NOZZLE HEI GHT FLAP RP,GIUS RUN LENGTH AHFAD OF FLAP r F10!1: S C R F E Y S

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TRANS l T1ON S K T t ON -' LENGTH

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FLAP RADl!!SJ

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1 " Figure 2.- PIodel conf i g u r a t i o r , ~ a r i a b l e s for p a r n l l el exf t flow,

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\ JET F I OW NOZZLE ASPECT RATIO (W I DTHIHE I GHT)

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28.0

FLOW 1

ANGLE, deg 14.0

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1.0 1.5 2.0 2.5 3.0 3.5 NOZZLE PRESSURE RAT!O Figure 3.- Effect of pressure r a t i o on flow turning for various nozzle aspect r a t i o s . Constant f l a p radius.

FLAP R A D I U S NOZZLE HE1 GHT FLOW TURNING ANGLE, deg - 0 1.5 2.0 2.5 3.0 NOZZLE PRESSURE RATIO Figure 4 . - Effect of pressure r a t i o on flow turning for a range of Nozzle aspect ratio of 2 8 .

f l a p radius.

' , L - . -.d-- A!'. .- - ...--- --.- NOZZLE ASPECT R A T I O

0-------_

---v 28 FLOW TURN l N G ANGLE. deg 1 - 1 I 1 1 I I I 0 1 2 3 4 5 6 7 8 FLAP R A D 1US:NOZZLE HEIGHT ' i g u r e 5.- E f f e c t of f l a p r a d i u s on flow t u r n i n g f o r a p r e s s u r e r a t i o of 1 . 4 and a c o n s t a n t run l e n g t h .

NOZZLE ASPECT R A T I O 60 r NOT A V A I LA BLE W I T H B A S I C NOZZLE (PARALLEL FLOW)

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M A X InilUM FLG\"I T U R N I h G

ANGLE, deg r 7

A V A I L A B L E W I T H B A S I C NOZZLE (PARALLEL FLOW1 0 1 2 3 4 5 FLAP R A D I U S I N O Z Z L E HEIGHT Figure 6 . - E f f e c t of f l a p r a d i u s on maximum flow t u r n i n g f o r a range of p r e s s u r e r a t i o and r u n l e n g t h .

NOZZLE S P R F A D ANGLE, d q SPREAD ANGLE

lo e

TOP V I E W F LO\Y TURNING FLAP R A D I U S - ~ Z Z L E HE I GHT ANGLE. deg NOZZLE A 5 PECT R A T I@. 7 ROOF ANGLE

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0 20 4r) % SIDE VIEW

NOZZLE ROOF AWGLE, deg Figure 7 . - Effects of nuzzle roof angle and spread angle on flow turning.

Figure 8. - Four-engPne USB model with modesate-radius f Laps in the L a n ~ l e y V/STOL tunnel.

23' ROOF ANGLE FLAP R A D I U S 29O SPREAD ANGLE =E- : 3 . 2 W l N D O N

-

STATIC b 4 , deq -

Figure 9 . - S t a t i c turning and wind-on l i f t for moderate-radius flap. a = 0'.

W l N D O N S T A T I C b f . deg b f , deg

6o -

FLOW T U R N I N G ANGLE. deg 0 0 THRUST 1 2 3 4 C~ Figure 10.- Effect of f l a p radius and deflection on s t a t i c turning and wind-on l i f t .

a = 0'.

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Figure 11. - Full-span leading-edge-b lowing model with small-radius

I f l a p s in the Langley V/STOL tunnel.

i 0 20 40 60 80 1000 2 0 40 80 1100 TOTAL FIAP DEFLECTION, deg Figure 12.- Static turning arc! wind-on lift for small-radius f l a p .

a = O O .

h18-24049' .. . - a

RESULTS OF STATIC TESTS OF A SCALE M O D E L

OF THE BOEING YC-14 POWERED-LIFT SYSTEM 1:: James L. H a s s e l l , Jr.

NASA Langley Research Center S U M M A R Y One-quarter-scale s t a t i c ground t e s t s of t h e Boeing YC-14 p o w e r e d - l i f t f f u l l - s c a l e t e s t r e s u l t s . The 1/4- system were conducted f o r c o r r e l a t i o n w i t h s c a l e model u t i l i z e d a JT-15D turbofan engine t o r e p r e s e n t t h e CF6-50D e n g i n e The employed on t h e YC-14 advanced medium STOL t r a n s p o r t p r o t o t y p e a i r c r a f t .

t e s t s included e v a l u a t i o n of s t a t i c t u r n i n g performance, s t a t i c s u r f a c e p r e s s u r e and temperature d i s t r i b u t i o n s , f l u c t u a t i n g l o a d s , and a c c e l e r a t i o n s of p o r t i o n s of t h e wing, f l a p s , and f u s e l a g e . R e s u l t s a r e p r e s e n t e d f o r t h e range of f a n p r e s s u r e r a t i o a s l a n d i n g f l a p c o n f i g u r a t i o n o v e r an a p p r o p r i a t e and v o r t e x g e n e r a t o r a f f e c t e d by s e v e r a l v a r i a b l e s i n c l u d i n g ground h e i g h t m o d i f i c a t i o n s . S t a t i c t u r n i n g a n g l e s of t h e o r d e r of 60' were obtained. The h i g h e s t s u r f a c e p r e s s u r e s and temperatures were c o n c e n t r a t e d o v e r t h e upper s u r f a c e of t h e f l a p s i n t h e region immediately a f t of t h e USB n o z z l e .

INTRODUCTION P a s t NASA r e s e a r c h on t h e upper-surface blowing (USB) concept h a s pro- g r e s s e d from t e s t s of s m a l l - s c a l e powered models f o r e v a l u a t i o n s of low-speed 1 ; powered-lift performance, s t a b i l i t y , and c o n c r o l ( r e f s . 1 t o 4) t o t e s t s of l a r g e - s c a l e models powered by r e a l t u r b o f a n e n g i n e s f o r e v a l u a t i o n s of t h e o p e r a t i o n e l environment produced on t h e upper wirg and f l a p s u r f a c e s by t h e !I i The d a t a b a s e provided by t h i s r e s e a r c h h a s provided concept ( r e f s . 5 t o 7).

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a v a l u a b l e foundation f o r f u r t h e r development of t h e USB concept. A s a r e s u l t

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1 , < , of i n c r e a s e d i n t e r e s t i n USB and t h e performance p o t e n t i a l i n d i c a t e d f o r such t h e Boeing Company r e c e n t l y i n c o r p o r a t e d t h e USB concept i n t h e a system, J YC-14 advanced medium STOL t r a n s p o r t p r o t o t y p e . A s i n d i c a t e d i n f i g u r e 1, t h e YC-14 u t i l i z e s Coanda f l a p s i n c o n j u n c t i o n w i t h USB n o z z l e s t o provide low- , speed powered l i f t .

It w i l l be noted, however, t h a t t h e a i r p l a n e u s e s D- n o z z l e s t o provide e f f i c i e n t performance during low speed o p e r a t i o n s and a t c r u i s e .

The D-nozzle, which h a s a s e m i e l l i p t i c a l e x i t shape, r e p r e s e n t s a marked change i n design from t h e high-aspect-ratio, r e c t a n g u l a r n o z z l e s pre- v i o v s l y s t u d i e d by NASA i n l a r g e - s c a l e t e s t s ( r e f s . 5 and 6 ) .

The development program f o r t h e YC-14 included f u l l - s c a l e s t a t i c t e s t s of the powered-lift system components shown i n f i g u r e 2.

These components included t h e CF6-50D engine, t h e USB n c z z l e , a s t u b wing, t h e USB f l a p s , ana

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, .

an a d j a c e n t b o i l e r p l a t e s e c t i o n of t h e f u s e l a g e . Some uf t h e r e s u l t s of t h e s e f u l l - s c a l e t e s t s a t Boeing's T u l a l i p t e s t f a c i l i t y a r e p r e s e n t e d i n r e f e r e n c e 8.

I !

Boeing's f u l l - s c a l e s t a t i c t e s t of t h e YC-14 p o n v e r e d - l i f t system provided t h e impetus f o r reduced-scale t e s t s a t Lsngley u s i n g a s m a l l t u r b o f a n engine - !

t h e primary o b j e c t i v e b e i n g t o e v a l u a t e s c a l i n g r e l a t i o n s h i p s f o r t h e v a r i o u s t e c h n o l o g i e s involved. Other o b j e c t i v e s were t o conduct e x p l o r a t o r y powered- I l i f t r e s e e r c h w i t h a c r u i s e - c o n f i g u r e d n o z z l e and t o e v a l u a t e a broader range I of v a r i e b l e s t h a n p r a c t i c a l a t T u l a l i p because of t h e c o n s t r a i n t s i n time and C O S t .

t h e E v a l u a t i o n o i t h e s c a l i n g r e l a t i o n s h i p s w i l l be accomplished l a t e r ; p r e s e n t paper is l i m i t e d t o 8 f!,scussion of some of t h e 1 / 4 - s c a l e t e s t r e s u l t s .

geometric a r e a of secondary ( f a n flow) d u c t a t mixing p l a n e , Asecondary cm2 ( i n 2 ) I

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s u r f a c e s t a t i c p r e s s u r e r a t i o .

Incremental l o c a l s u r f a c e s t a t i c p r e s s u r e Ambient p r e s s u r e ambient t e m p e r a t u r e , "C (OF) nondimensi3nal spanwise l o c a t i o n , Spanwise d i s t a n c e from a i r p l a n e c e n t e r l i n e A i r p l a n e wing semispan D i s t a n c e nol-ma1 to l o c a l s u r f a c e nondimensional normal l o c a t i ~ n ,

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tangency a n g l e of t h e upper s u r f a c e of t h e af t f l a p a t t h e t r a i l i n g edge measured w i t h r e s p e c t t o wing chord p l a n e , deg

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t h r u s t recovery e f f i c i e n c y , Thrust upper-surface blowing DESCRIPTION OF M O D E L The Langlcy 1 / 4 - s c a l e model of t h e YC-14 powered-lift system and t h e s t a t i c ground t e s t a p p a r a t u s is shown i n f i g u r e 3. The model is powered by a JZ-15D t u r b o f a n engine r a t e d st 9786 N (2200 l b f ) s t a t i c t h r u s t ( a s cqnpared w i t h 222.4 k c (50 000 l b f ) t h r u s t f o r t h e CF6-50D engine of t h e f u l l - s c a l e YC-14).

The s i d e door on t h e USB n o z z l e was t e s t e d o n l y i n t h e "door open" c o n f i g u r a - t i o n corresponding t o t h a t used f o r powered-lift o p e r a t i o n . T h i s door has two f u n c t i o n s : F i r s t , i t provides t h e n e c e s s a r y n o z z l e a r e a v a r i a t i o n t o o b t a i n engine match c o n d i t i o n s f o r b c t h take-off and c r u i s e , and second, i n t h e open p o s i t i o n f o r powered-lift o p e r a t i e n , i t h e l p s s p r e a d t h e exh,:ust gas flow ovt- board over t h e Coanda i l a p span. E s s e n t i a l l y a l l geometric d e t a i l s of t h e n o ~ z l e , wing, f l a p s , v o r t e x g e n e r a t o r s , and f u s e l a g e a r e s c a l e d from t h e YC-14.

I n t h i s view t h e model is vounted a t a s c a l e d h e i g h t above t h e graund c o r r e - sponding t o t h e f u l l - s c a l e t e s t a t T u l a l i p ; t h e upper s ~ r f a c e of t h e wing is 1.45 m (4.75 f t ) above t h e s u r f a c e , compared w i t h 5.80 m (19 f t ) f o r t h e f u l l - s c a l e c o n f i g u r a t i o n a t T u l a l i p .

Figure 4 sbows how t h e YC-14 USB nozzle was adapted t o t h e JT-15D e n g i n e , which h a s a b y ~ s s s r a t i o of 3 . 3 and a maximum fan p r e s s u r e r a t i o of 1.4.

(The CF6-50D engine has a bypass r a t i o of 4.4 and a maximum f a n p r e s s u r e r a t i o of 1 . 6 . ) Thc .e d i f f e r e n c e s i n engine c h a r a c t e r i s t i c s f o r c e d a compromise i n t h e 114-scale nozzle d e s i g n ; t h e primary and secondary a r e a s a t t h e n i x i n g p l a n e were a d j u s t e d t o t h e v a l u e s shown i n f i g u r e 4 r a t h e r than b e i n g geo- m e t r i c a l l y s c a l e d . B a s i c a l l y , t h i s design f e a t u r e of t h e n o z z l e was ccmpro- mised i n o r d e r t o match s t a t i c p r e s s u r e s from t h e f a n and c o r e flows. A l l o t h e r geometric d e t a i l s of t h e YC- 14 n o - z l e , such a s t h e skewed plug primary, were r e t a i n e d i n t h e 114-scale r ~ c l . The purpose of t h e skewed pldg primary is t o d i r e c t t h e h o t c o r e flow t o t h e top of t h e mixed flow n o z z l e a r d t h e r e b y minimize thermal problems on t h e wing and f l a p s .

Some of t h e c h a r a c t e r i s t i c s of t h e JT-15D engine a r e compared w i t h t h e CF6-5013 engine i n t a b l e 1. Aside from t h e very l a r g e d i f f o r e n c e i n t h r u s t , t h e s e two t u r b o f a n e n g i n e s have somt.ahat s i m i i a t c h a r a ~ , . . e r i s t i c s . With b o t h e n g i n e s o p e r a t i n g a t a f a n p r e s s u r e r a t i o of 1.4, t h e fan dnd c o r e v e l o c i t i e s of t h e two e n g i n e s correspond f a i r l y w e l i . T h i s o p e r a t i n g c o n d i t i o n i s abgut 66 p e r c e n t maximum t h r u s t f o r t h e CF6-50D e n g i n e , However, a s a r e s u l t of t h e 4 7

1 i

, I : ' , \'..I , 3' I , d i f i e r e n c e i n b y p a s s r a t i o s , t h e e x h a u s t £1. i n t h e l / 4 - s c a l e t e s t s i s com- i b posed of a l a r g e r p r o p o r t i d n o f c o r e f l o w , r e s u l t i n g i n t h e h i g h e r n r a k tempera- i ' !

t u r e a t t h i s o p e r a t i n g c o n d i t i o n . The t e m p e r a t u r e d a t a of t h e i / 4 - s ~ a l e t e s t s s h o u l d t h e r e f o r e g i v e c o n s e r v a t i v e r e s u l t s a s r e l a t e + i t o t h e YC-14 a i r p l a n e .

I

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A s k e t c h oi t h e i l 4 - s c a l e a p p a r a t u s is F -9sen:ed i n f i g u r e 5 t o p r o v i d e a d e s c r i p t i o n o r t h e f o r c e measurement system. The e n t i r e model, i n c l u 6 i n g t h e b e l l m o u t h i r . l e t , e n g i n e , n o z z l e , wing, f l . i p , and f u s e l a g e , Ls s u p p o r t e d on a f l o a t i n g f r m e which has s:i,in g a g e s fcn;ircr ltld a f t t o measure normal f o r c e and J s i m p l e rod-rigged l o a d c e l l t o measure a x i a l f o r c e . These measurements p r o v i d e t h e magnitude and d i r e c t i o n of t h e r e s c l t a n t f o r c e . The r e s u l t a n t f o r c e is t h e t i r u s t r e c o v e r e d , and t h e d i r e c t i o n of t h e r e s u l t a n t f o r c e is i h e j e t t u r n i n g a n g l e 6 , . I n o r d e r t o d e t e r m i n e t h r u s t , t h e f o i w a r d e n g i n e mounts a r e p i v o t e d i n p i l l o w b l o c k s which a r e i n s t r u m e n t e d w i t h s t r a i n g a g e s t o measure e n g i n e a x i a l f o r c e . The a f t e n g i n e mount is s u p p o r t e d by an a d j u s t a t ? . e ll.r,k w i t h rod end b e a r i n g s a t t h e t o p and battom. The USB n o z z l e i s mounted t o t h e e n g i n e c a s e b u t does n o t Eoucn t h e wiilg, t h e f u s e l a g e , o r any o t h e r p a r t of t h e f l o a t i n g frame s t r u c t u r e . A r u b b e r s e a l i s o l a t e s t h e f a c e of t h e e n g i n e from t h e bellmoutr: i n l e t . The o u t p u t uf t h e e n g i n e a x i a l - f o r c e s t r a i n gage was ca , i b r a t e d a g a i n s t measured t h r u s t w i t h t h e wtng, f l a p s , and f u s e l a g e reffioved, and t h e c a l i b r a t i o n o b t a i n e d was used t o d e t e r m i n e measured t h r u s t d k - r i n g subsequelit t e s t i n g w i t h t h e wing, f l a p s , and f u s e l a g e i n p l a c e . T h r u s t r e c o v e r y e f f i c i e n c y r\ is t h e n c'efined as t 1 - i ~ r a t i o o f t h e t h r u s t r e c o v e r e d t o t h e measured t h r u s t .

Sensor l o c a t i o n s f o r o t h e r t y p e s of d a t a a r e shown i n f i g u r e 6 . S t a t i c p r c s s u r e p o r t s and thernocoup!es were d i s t r i b u t e d u n i f o r m l y o v e r t h e wings and f l a p s and a l s o q v e r a p p r ~ p r i a t e a r e a s of t h e f u s e l a g e . S u r f a c e micto- phones and a c c e l e r o m e t e r s w e r a a l s o p o s j t i o n e d or. t h e wing, f l a p s , end f u s e l a g e , and a r e d e s c r i b e d i n r e f e r e n c e 9.

I n a d d i t i o n , p r e s s u r e and t e m p e r a t u r e r a k e s were i n s t a l l e d i n b o t h t h e e n g i n e f a n d u c t s and i n t h e primary n o z z l e , and t h e bellmouth i n l r t was i n s t r u - mented t o measure i n l e t mass flow. T h i s e v g i n e i n s t t m e n t a t i o n provided f o r d e t e r m i n a t i o n of i d e a l t h r u s t a n d , when r e l a t e d t o t h e measured t h r u s L , p e r m i t t e d e v a l u a t i o n of t h e n o z z l e v e i o c i t y c o e f f i c i e n t .

The f l u c t u a t i n g l o a d s and a c c e l e r a t i o n d a t a a r e c u r r e n t l y b e i n g a n a l y z e d and w i l l n o t be p r e s e n t e d . The d a t a d i s c u s s e d w i l l t h e r e f o r e be l i m i t e d t o s t a t i c t u r n i n g ? e r f o r u a n c e , s u r f a c e p r e s s u r e s and t e m p e r a t u r e s , and f l o w s c r v e y s a s a f f e c t e d b y s e v p r a l of t h e more i m p c r t a n t v a r i a b l e s t o r t h e l a n d i n g f l a p c o n f i g u r a t i o n which h a s a t r a i l i n g - e d g e u p p e r - s u r f a c e a n g l e of € ! 6 . s 0 , The t e s t v a r i a b l e s i n c l u d e d a range of f l a p d e f l e c t i o n s , t h r u s t c o r r e - sponding t o f a r pressure r a t i o s f r o n 1.1 t o i . 4 , and h e i g h t s above t h e ground c o r r e s p o n d i n g tc wheel c o n t a c t h e i g h t (5.80 m (19 I t ) f u l l s c a l e ) , t o an a i r b o r n e h e i g h t (9.14 T (30 f t ) f u l l s c a l e ) , and t o a f r e e - a i r c o n d i t i o n .

4 8

The e f f e c t i v e n e s s of v o r t e x g e n e r a t o r s i n a i d i n g flow attachment is c r i t i c a l i n t h i s powered-lift concept where t h e d e s i g n e r h a s e l e c t e d t o avoid a h i g h nozzle kickdown ax~gle b e c a u ~ e of c r u i s e performance c o n s i d e r a t i o n s ; t h e v o r t e x g e n e r a t o r = o n f i g u r a t i o n t h e r e f o r e was considered t o be an important v a r i a b l e .

STATIC TURNING PERFORMANCE Figure 7 shows t h e e f f e c t of vortex-generator deployment on s t a t i c t u r n i r ? performance f o r t h e l a n d i n g f l a p r m f i g u r a t i o n . The d a t a a r e pre- and t h e t h r u s t recovery s e n t e d i n terms of t h e j e t t u r n i n g a n g l e tij e f f i c i e n c y as f u n c t i o n s of f a n p r e s s u r e r a t i o . The p r e s s u r e r a t i o range * of l n t e r e s t f o r t h e YC-14 o p e r a t i n g i.n t h e l a n d i n g approach is between 1.25 I i; 3 ! and 1.4. The r e s u l t s p r e s e n t e d were obtained f o r an h / b v a l u e of 0.147, C - 1

.- . _

t h e h e i g h t corresponding t o t h e YC-14 w i t h wheels on che g r o m d . For t h i s - : . p a r t i c u l a r comparison t h e gap between t h e n o z z l e and t h e wing war unsealed.

. - k d : . . . ...

The r e s u l t s i n d i c a t e deployment of t h e b a s i c v o r t e x g e n e r a t o r s p r o v i d e s about :.

.

. _' .. 12' improvement i n j e t t u r n i n g a n g l e w i t h a d e c r e a s e i n t h r u s t recovery - e f f i c i e n c y of about 5 t o 6 p e r c e n t . Also, v a l u e s of jet t u r n i n g a n g l e tend These r e s u l t s demonstrate t o decrease w i t h i n c r e a s i n g f ~ n p r e s s u r e r a t i o .

t h e e f f e c t i v e n e s s of r e t r a c t a b l e v o r t e x g e n e r a t o r s i n providing improved j e t t u r n i n g a n g l e without r e s o r t i n g t o a high nozzle kickdown a n g l e which n i g h t unduly compromise c r u i s e e f f i c i e n c y . A t t h i s p o i n t it should b e acknowledged t h a t t h e f u l l - s c z l e s t a t i c t e s t s a t T u l a l i p produced 5' t o 6O b e t t e r s t a t i c t u r n i n g angle ~ 5 t h v c r t e x g e n e r a t o r s r e t r a c t e d and about 8 b e t t e r t u r n i n g : angle w i t h v o r t e x g e n e r a t o r s deployed. These d i s c r e p a n c i e s a r e p o s s i b l y re- . . l a t e d t o t h e j u n c t ~ r e between t h e ITSR nozzle and t h e wing upper s u r f a c e . On

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f t h e f u l i - s c a l e 1;-14 t h e r e is no j u n c t u r e ; t h e f l o o r of t h e n o z z l e i s formei?

by t h e upper s u r f a c e of t h e wing.

Figure 8 shows t h e e f f e c t of h e i g h t above t h e ground on s t s t i c t u r n i n g perfo--mance. For t h i s co~npa: i s o n t h t b a s i c v o r t e x g e n e r a t o r s a r e i n t h e up position,and t h e c l e a r a n c e gap between t h e edge of t h e USB n o z z l e a?d t h e upper s u r f ace of t h e wing h a s been s e a l e d t o b e t t e r r e p r e s e n t t h e f u l - - s c a l e PC-14 nozzle arrangement. R e s u l t s a r e p r e s e n t e d f o r two ground h e i g h t s : h/b v a l u e s of 0.147, corresponding t o a wheel-on-the-grouild c o n d i t i o n , and 0.232, c o r r e s l o n d i n g t o a c o n d i t i o n wherein t h e YC-14 wing is a 2 a h e i g h t of 9.14 m (30 f t ) above t h e runway.

The r e s u l t s i n d i c a t e only a5out a lo i m - provement i n j e t t u r n i n g angle and a few r e r c e n t b e t t e r t h r u s t recovery e f f i c i e n c y a t t h e h i g h e r h e i g h t . The i n f i r i t e ground h e i g h t c o n d i t i o n is y e t t o be t e s t e d .

!

I n o r d e r t o a s s u r e an adequate l e v e l of powered-lift performance i n t h e

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landing approach c o n d i t i o n , f u r t h e r improvement i n s t a t i c t u ~ n i n g performance t o vrilues approaching 60' was needed, e s p e c i a l l y a t t h e h i g h e r f a n p r e s s u r e

I ! ! . I <

j r a t i o s . A s shown i n f i g u r e 9 , two m o d i f i c a t i o n s t o t h e b a s i c v o r t e x g e n e r a t o r s , . , were made i n an attempt t o improve -ae t u r n i n g . The f i r s t m o d i f i c a t i o n simply ! i;

. . : 1 . 4

doubled t h e span of t h e o u t s i d e p a i r of v o r t e x g e n e r a t o r s s o t h a t more of t h e j e t e f f l u x would b e a f f e c t e d and theyeby r e e n e r g i z e t h e boundary l a y e r a t t h e knee o f t h e f l a p b e t t e r t h a n w i t h t h e b a s i c v o r t e x g e n e r a t o r s . The second m o d i f i c a t i o n r e t a i n e d t h e same i n c r e a s e d span and a l s o i n c r e a s e d t h e inc!dence a n g l e 10' on t h e o u t s i d e p a i r . The r a t i o n a l e f o r t h i s second m o d i f i c a t i o n was t o i n t e n s i f y t h e v o r t i c e s produced by t h e o u t s i d e p a i r o h o r t e x genera- t o r s s o a s t o f u r t h e r e n e r g i z e t h e boundary l a y e r and t h s r e b y promote b e t t e r f l o w attachment toward t h e t r a i l i n g edge of t h e f l a p s .

Figure 1 0 p r e s e n t s t h e s t a t i c t u r n i n g performance o b t a i n e d w i t h t h - two v o r t e x g e n e r a t o r m ~ d i f i c a t i o n s a s compared w i t h t h e b a s i c v o r t e x g e n e r a t o r s .

Both m o d i f i c a t i o n s improved t h e j e t t u r n i n g a n g l e , e s p e c i a l l y a t t h e i m p o r t s o t h i g h e r f a n p r e s s u r e r a t i o s . The second m o d i f i c a t i o n provided a s i g n i f i c a n t So o r 6' i m p r o v ~ n e n t i n j e t t u r n i n g a n g l e a t a c o s t of about 4 p e r c e n t i n t h r u s t recovery c t ' f i ~ i e t i i y . i

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SURFACE PLSSURE AND TEMPERATUS DISTRIBUTIONS

!

, .

F i g u r e 11 p r o v i d e s a b a s i s f o r understanding t h e flow c h a r a c t e r i s t i c s on t h e u p p e r ' s u r f a c e of t h e wing and f l a p s . S u r f a c e p r e s s u r e r a t i o c o n t m r s a r e

1 ; ! t#>

shown p l o t t e d on t h e upper s u r f a c e ~ f t h e wing and f l a p s i n t h e r e g i o n a f t of t h e USB n o z z l e Negative v a l u e s of p r e s s u r e r e t i o i n d i c s t e s u c t i o n p r e s s u r e s

I 1 '

and p o s i t i v e v a l u e s i n d i c a t e impsct p r e c s u r e s . I n a d d i t i o n t o p r o v i d i n g i n d i - c a t i o n s of good o r poor flow a t t a c h m e n t , t h i s type of d a t a is a l s o r a l u a b l e 1 ' 1 .

tc t h e d e s i g n e r f o r t h e d e t e r n i n a t i o n of s t e a d y - s t a t e s t a t i c l o a d s o v e r p o r t i o n s , I of t h e wing and f l a p s . The c a s e s h o ~ n i n f i g u r e 11 is f o r t h e l a n d i n g f l a p f !

!

a , c o n f i g u r a t i o n w i t h t h e b z s i c v o r t e x g e n e r a t o r s deployed, f o r a v a l u e of h/b

e

of 0.147 and f o r a f a n p r e s s u r e r a t i o of 1.36. Along t h e c e n t e r l i n e of t h e

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t h r u s t a x i s , s u c t i o n p r e s s u r e s occur almost t o t h e L r a i l i n g edge i n d i c a t i n g 1 I : : !

e x c e l l e n t f l o v a t t a c h m e n t , b u t r e g i o n s O F poor f l a w attachment a r e i n d i c a t e d f i i b o t h inboard a d j a c e n t t o t h e f u s e l a g e ana outboard toward t h e t i p of t h e f l a p . , , i I ' .

A c r o s s p l o t of t h e s u r f a c e p r e s s u r e r a t i o s along t h e spanwise dashed l i n e i n d i c a t e d on t h e a f t f l a p i n f i g u r e 11 p r o v i d e s a b e t t e r indication of t h e degree of flow attachment.

Figure 12 p r e s e n t s t h e c r o s s p l o t of a'? f l a p p r e s s u r e s t o show t h e e f f e c t o f v o r t e x g e n e r a t o r m o d i f i c a t i c n s . S u r f a c e p r e s s u r e r a t i o Ap/pamb is p l o t t e d a g a i n s t t h e n o n d i m e n s i ~ n a l s p m p a r a a s t e r 2- w i t h n e g a t i v e b / 2 The l o c a t i o n s of t h e c e n t e r l i n e and t h e v a l u e s of p r e s s u r e r a t i o upward.

s i d e of t h e f u s e l a g e a r e i n d i c a t e d i n t i e f i g u r e . The n c s t dramatic improve- ment i n flow attachment due t o t h e v o r t e x g e n e r a t o r m o d l f i c a t i o n s a c c u r s i n t h e inbcard a r e a of t h e f l a p s a d j a c e n t t o t h e f u s e l a g e . ~ i ~ e j e t t u r n i n g a n g l e s f o r t h i s p r e s s u r e r a t i o of 1.4 ( r e f e r t o f i g . 10) a r e a l s o i n d i c a t e d i n f i g u r e 12 f o r t h e b a s i c v o r t e x g e n e r a t o r s and f o r each of t h e m o d i f i c a t i o n s .

These p r e s s u r e d a t a t h e r e f o r e i n d i c a t e t h a t a l l t h e improvement i n s t a t i c t u r n i n g a n g l e r e s u l t e d from t h e irtboard v o r t e x geiierator ruodification. The outboard modified v o r t e x g e n e r a t o r only p e n a l i z e d t h e c h r u s t recovery e f f i c i e n c y .

Surface-temperature c o n t o u r s o v e r t h e wing and f l a p s f o r t h e c a s e w i t h b a s i c v o r t e x g e n e r a t o r s deployed a r e p r e s e n t e d i n f i g u r e 13. The h i g h e r temperatures from t h e primary flow a r e c o n c e n t r a t e d n e a r t h e c e n t e r l i n e , w i t h peak v a l u e s s l i g h t l y above 204O C (40"' F) on t h e f l a p system. F l a p s t r u c t u r e s of s u i t a b l e h i g h - t e m p e r a t u r e - r e s i s t a n t a l l o y s such as s t a i n l e s s s t e e l should e x p e r i e n c e no problem i n t h i s environment. The s u r f a c e temperatures i n t h e r e g i o n of wing s t r u c t u r e p r e s e n t no problem, even f o r aluminum a l l o y s . Although n o t p r e s e 2 t e d i n i i g u r e 13, t e m p e r r t u r e s on t h e s i d e of t h e f u s e l a g e never ex- ceeded 60 C (140 F).

FLOW SURVEYS Various flow surveys were made t o determine t h e e x t e n t of t h e j e t e f f l u x d i s p e r s i o n and t o e v a l u a t e l o c a l flow c o n d i t i o n s r e l a t e d t3 s p e c i f i c micro- phone l o c a t i o n s . I n o r d e r t o o b t a i n d e t a i l e d v e l o c i t y p r o f i l e s of t h e flow a d j a c e n t t o t h e wing and f l a p upper s u r f a c e s , a s u r v e y rake about 48.3 cm (19 i n . ) l o n g having 25 t o t a l p r e s s u r e probes and fewer s t a t i c p r e s s u r e and thermocouple probes was used. D i s c r e t e v e l o c i t y measurements were o b t a i n e d a t each of t h e r o t a 1 p r e s s u r e probe l o c a t i o n s .

F i g u r e 14 shows t h e r e s u l t s of some o f t h e flow surveys f o r t h e c a s e of v o r t e x g e n e r a t o r s deployed a t t h e h i g h e s t f a n p r e s s u r e r a t i o . V e l o c i t y pro- f i l e s n e a r t h e c e n t e r l i n e a r e shown a t t h e n o z z l e e x i t , a t two l o c a t i o n s on t h e f l a p s corresponding t o microphone p o s i t i o n s , and a f t of t h e t r a i l i n g edge of t h e f l a p s . The rake was positio-led normal t o t h e l o c a l s u r f a c e f o r each l o c a t i o n . The peak v e l o c i t y a t the n o z z l e e x i t w a s a b o u t 375 m/sec (1230 f t / s e c ) and decayed t o about 251 m/sec (825 f t / s e c ) a t t h e f l a p t r a i l i n g edge. The p r o f i l e b e g i n s t o change a p p r e c i a b l y on t h e forward f l a p and then shows c o n s i d e r a b l e t h i c k e n i n g over t h e a f t f l a p element and a t rhe t r a i l i n g edge.

Figure 15 shows t h e r e s u l t s of rake s u r v e y s made over t h e span of t h e n o z z l e e x i t and t h e f l a p t r a i l i n g edge w i t h modified v o r t e x g e n e r a t o r s deployed ( m o d i f i c a t i o n 2 ) . The d a t a were a l s o o b t a i n e d a t t h e h i g h e s t v a l u e of f a n p r e s s u r e r a t i o of 1.4. I n t h e s e p l o t s t h e v e l o c i t y d a t a a r e p r e s e n t e d i n terms of c o n s t a n t v e l o c i t y c o n t o u r s measured i n t h e two p l a n e s ' i n d i c a t e d - a t t h e n o z z l e e x i t and a t t h e f l a p i - a i l i n g edge. The p l o t s a r e o r i e n t e d t o r e l a t e t h e spanwise e x t e n t of t h e ITSB n o z z l e t o t h e f l a p span. T h e d a t a show t b ? t t h e h i g h e r v e l o c i t y flow from t h e primary n o z z l e is concentrii.-cl i n t h e upper c e n t e r o f t h e USB n o z z l e , w i t h some i n d i c a t i o n c f s p r e a d i n g of t h e lower v e l o c i t y f a n flow beyond the p r o j e c t i o n o f t h e n o z z l e due t o t h e "door open" geometry. Much of t h e flow p a t t e r n a t t h e n o z z l e e x i t is a p p a r e n t a t t h e f l a p t r a i l i n g edge, and i t should be noted t h a t t h c h i g h e r v e l o c i t y flow i s s t i l l c o n c e n t r a t e d along a p r o j e c t i o n of t h e n o z z l e c e n t e r l i n e . The flow was thickened 3 t o 4 times t h e nozzle depth b u t h a s s p r e a d v e r y w e l l a c r o s s t h e f l a p span. Of p a r t i c u l a r i n t e r e s t i s t h e f u r t h e r i n d i c a t i o n of good flow attachment a t t h e f l a p t r a i l i n g edge n e a r t h e s i d e of t h e f u s e l a g e which was brought about by t h e modified v o r t e x g e n e r a t o r s .

CONCLUDING RETlARKS P r e l i m i n a r y r e s u l t s of s t a t i c t e s t s of t h e 114-scale Boeing YC-14 USB model i n d i c a t e t h a t t h e s t a t i c t u r n i n g performance of t h e l a n d i n g f l a p configu- r a t i o n was improved a p p r e c i a b l y by t h e use of t h e b a s i c v o r t e x g e n e r a t o r design.

Regions of poor flow attachment were noted near t h e t r a i l i n g edge of t h e f l a p Improved both inboard a d j a c e n t t o t h e f u s e l a g e and outboard toward t h e t i p .

flow attachment was o b t a i n e d by t a i l o r i n g t h e v o r t e x g e n e r a t o r s which r e s u l t e d Peak v a l u e s of s u r f a c e i n f u r t h e r improvement. i n s t a t i c t u r n i n g performance.

s t a t i c p r e s s a r e s and t e m p e r a t u r e s were concentrated over t h e upper s u r f a c e of t h e f l a p s along t h e c e n t e r l i n e of t h e t h r u s t a x i s . These r e s u l t s t o g e t h e r w i t h r e s u l t s d e a l i n g w i t h f l u c t u a t i n g l o a d s , c u r r e n t l y under a n a l y s i s , w i l l be c o r r e l a t e d w i t h f u l l - s c a l e YC-14 s t a t i c d a t a w i t h an e v e n t u a l o b j e c t i v e of e s t a b l i s h i n g a p p r o p r i h t e s c a l i n g laws f o r t h e v a r i . > u s t e c h n o l o g i e s involved.

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: i . 1 . ) . , . , - - . . . a + .

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REFERENCES . , 1 . Phelps, Arthur E . ; Letko, William; and Henderson, Robert L. : Low-Speed i i iJind-Tunnel Investigation oi a Semispan STOL Jet Transport Wing-Body With I .

an Upper-Surface Blown Jet Flap. NASA TN D-7183, 1973. 1 : , .

2. Phelps, Arthur E., 111; and Smith, Charles C., Jr.: Wind-Tunnel Investiga- tion of an Upper Surface Blown Jet-Flap Powered-Lift Configuration.

NASA TN D-7399, 1973.

3 . Sleeman, "illiam C . , Jr. ; and Hohlweg, William C. : Low-Speed Wind-Tunnel

Invest~gation of a Four-Engine Upper Surface Blown Model Having a Swept win^ an? Xectangular and D-Shaped Exhaust Nozzles. NASA TN D-806l, 1975.

! 1 i

I 4 . Phelps, Artli-*r E., 111: Wind-Tunnel Investigation of a Twin-Engine Srraight- ! .

Wing Upper-S~rface Blown Jet-Flap Configuration. NASA TN 0-7778, 1975. j I I i 5 . Staff of the Langley Research Center: Wind-Tunnel Investigation of the Aerodynamic Performance, Steady and Vibratory Loads, Surface Temperatures and Acoustic Characteristics of a Large-Scale Twin-Engine Upper-Surface Blown Jet-Flap Configuration. NASA TM X-72794, 1975.

6. Shivers, James P. ; and Smith, Charles C., Jr. : Static Tests of a Sim~lated Upper Surface Blown Jet-Flap Configuration Utilizing a Fuil-Size Turbofan Engine. NASA TN D-7816, 1975.

7 . Carros, Robert J.; Boissevain, Alfred G.; and Aoyagi, Kiyoshi: Aerodynamic Characteristics of a Large-Scale Hybrid Upper Surface Blown Flap Model

i-I

Having Four Engines. NASA TM X-62460,.1975.

. ! . .

8 . Sussman, M. B.; Harkonen, D. L.; and Reed, J. B . : USB Environment Measure- , ments Based on Full-scale Static Engine Ground Tests. Powered-Lift Aero- dynamics and Acoustics, NASA SP-406, 1976. (Paper no. 30 of this compilation.)

! I : 9. Schoenster, James A.; Willis, Conrad M.; Schroeder, James C.; and ! I : . , .

' i Mixson, John S.: Acoustic-Loads Research for Powered-Lift Configurations.

Powered-Lif t Aerodynamics and Acoustics, NASA SP-406, 1976. (Paper no. 26 I , _ i ;~ of this compilation.)

- .

Table 1.- Comparison of Engine C h a r a c t e r i s t i c s CF6-50D JT-15D . . . . . . . . . . . . .

Rated t h r u s t , kN ( l b f ) 222.4 (50 000) 9.79 (2200)

B y p a s s r a t i o . . . . . . . . . . . . . . . . . . 4.4 3.3

Fan p r e s s u r e r a t i o . . . . . . . . . . . . . . . 1.6 1.4

Fan v e l o c i t y , m/sec ( f t / s e c ) . . . . . . . . . . a255 (836) 255 (836)

a

Core v e l o c i t y , m/sec (f t l s e c ) . . . . . . . . . 347 (1140) 375 (1230)

a700 (1260) 872 (1570) Peak exhaust g a s temperature, K (OR) . . . . . .

9 a l u e s a t 66 p e r c e n t maximum t h r u s t , corresponding t o a f a n p r e s s u r e r a t i o of 1.4.

Figure 1 .- Doeing YC- 14 adv ~ n c e d rne6inm STOL transport.

f

Figure 3. - 1/4-scale YC-14 ground t e s t apparatus.

-- 1 &/--rl

I

I /-- , 12.0~

--.

I - -l r - - 7 -

1 -4

L ------

1 I

- - - -

T R I C A R E A S A T h l l X I N G PLANE

_ _ - -

*primary 711.0 cm2 I 1 1 C . 2 II I I Figsre 4 . - YC-14 USE nozzle matched w i t h JT-15D turbofan engine.

Bypass ratio of 3 . 3 ; fan pressure ratfo of 1 . 4 ,

BELLMOUTH/ RUBBER SEAL

r . I / r T H R U S T GAGE JT-15D ENGINE

_,-- YC-14 NOZZLE rn 1 Figure 5 .- Force measurement system for 114-scale ground t e s t s .

c S T A T I C PRESSURE PORTS ONLY S T A T I C PRESSURE PORTS A N D THERMOCOUPLES Figure 6 . - Sensor locations on wing, f l a p s , and fusel-age.

r\ " VORTEX GENERATORS

a- -- -a-

6 ., deg

I -0 -,--- --

I \- RETRACTED

. 7 UL-U

1.05 1.10 1.15 1.20 1.25 1.30 1.35 1.40 FAN PRESSURE R A T I O Figure 7 . - Effect of basic vortex generators on s t a t i c turning performance.

h/b = 0,147; nozzle-wing gap unsealed.

---

\ hlb 7P

rl

0- - -9 g - ---P-- ---a

1.10 1.15 1.20 1.25 1.30 1.35 1.40 1.45 FAN PRESSURE R A T I O Figure 8,- Effect of height above ground on s t a t i c turning performance.

Basic vortex generators up.

M O D I F I C A T I O N NO. 1 DOlJ BLE S P A N ON OUTS I DE P A I R M O D I F I C A T I O N NO. 2 DOUBLE S P A N A N D INCREASE INCIDENCE ON OUTS I DE PA I R Figure 9 . - Vortex generator modifications.

M O D I F I C A T I O N 1 M O D I F I C A T I O N 2 "1.10 1.15 1;20 1.25 1.30 1.35 1.40 1.45 FAN PRESSURE R A T I O Figure 10.- Effect of vortex generator modifications on static turn performance.

Figure 11.- Surface pressure r a t i o contours with basic vortex gener = 86.5O; h/b = 0.147; fail pressure r a t i o of 1 . 3 6 .

Up' ' U S B -.14 r M O D I F I C A T I O N 2 - S I DE OF FUSELAGE - -.02

4 FUSELAGE 7

I 1 I 1 I I I I I

-.40 -.36 -.32 -.28 -.24 -.2C - . I 6 - 2 -.a - 0

,Y!

bR Figure 12.- Effect of vortex generator modifications.

Figure 13.- Surface temperature contours with basic vortex generators up.

h/b = 0.147; = 1 1 ' C (52' F) ; f a n pressurn r a t i o GUSB = 86.5'; Tamb o f 1.36.

174.9 m/ sec -

F2'K\, ( 1230 ft'sec) . C\ ,-. \ , , /25 1.5 m/sec - ' Figure 14.- Velocity p r o f i l e s from nozzle e x i t t o t r a i l i n g edge, = 86.S0; basic vortex generators up.

5~~~

F L A P T R A I L I N G EDGE S l l R V E Y F i g u r e 15.- V e l o c i t y c o n t o u r s a t n o z z l e e x i t and f l a p t r a i l i n g edse w i t h m c d i f i e d v o r t e x g e n e r a t o r s ( m o d i f i c a t i o n 2 ) . Fan p r e s s u r e r a t i o of 1 . 4 .

SUPWRY OF LOW-SPEED AERODYNAMIC CHARACTERISTICS OF UPPER-SURFACE-BLOWN JET-E'LAP CONFIGURATIONS Arthur E. Phelps I11 Langley Directorate, U.S. Army Air Mobility R&D Laboratory Joseph L . Johnson, Jr., and Richard J. Margason NASA Langley Research Center SUMMARY The results of recent wind-tunnel investigations to provide fundamental information on the upper-surface-blown (USB) jet-flap concept demonstrated that the USB concept prcvides good high-lift performance. The low-speed performance appears to be mainly dependent upon the jet turning angle and turning efficiency and on the use of proper leading- and trailing-edge treatnent to prevent pre- mature flow separation. The best means of achieving good turning performance in anv particular USB application nust be determtned from overall operational con- ~tderations in which high-speed performance, structures and noise, as vell as low-speed performance, are e~laluated. The large divi3g moments generated at high lift coefficients can bc trimmed satisfactorily with a l ~ r g e , conventional horizontal tail; a high tail position iq best from longitudinal stability con- siderations. Large rolling and yawing aoments are introduced with the loss of an engine, but these moments can be trimmed satisfactorily through the use of asymmetrical boundary-layer control and thrcugh the use of spoiler and rudder deflection as needed.

INTRODUCTION In recent years, considerable effort has been directed toward studies of the aerodynamic and acoustic characteristics of upper-surface-blown (USB) jet-flap configurations (refs. 1 to 7 ) . The results of past aerodynamic investigations have Indicate3 that the USB concapt can provide the high lift necessary for effi- citnt 3TOL operation; acoustic studies have indicated that the USB concept mav provide beneficial noise reduction during flyovers because the wing shields ground observers from the noise produced at the engine exhaust nozzle. More recent studies have provided solutions to stability and control problems such as pitch trim, longitudinal stability at high lift, and lateral trim for engine- out conditions.

The present paper has beer ~repared to summarize some of the more important characteristics of USB configurations in the areas of performance, longitudinal stability and trim, la1 :ral-directional stability, engine-out lateral trim, and dynamic stabilicy and coytrol. Although the discussion is directed toward JSB coniigurations, certain problems such as pitch trim and longitudinal stability ar-1 conmon to all powered-lift STOL systems; the problem of enr,. ~e-out lateral trim is common to other powe-od-lift concepts utilizing discrete blowing, such as the externally blown flap (EBF) arrangement. Therefore, the data presented

- in the present paper for USB configurations may also be generally applicable to other powered-lift concepts.

SYMBOLS aspect ratio wing span lift coefficient

CL

power-induced lift coefficient

C ~ r

maximum lift coefficient CL , max rolling-moment coefficient

effective dihedral parameter, ac1/aB

pitching-moment coefficient

longitudinal stability parameter, a & / a a a

yawing-moment coefficient directional stability parameter, a $ / a B gross thrust coefficient, T/qS wing mean aerodynamic chord axial force normal force tail length thrust time to damp to half-amplitude weight body axes coordinates angle of attack flight path angle flap deflection

jet deflection downwash angle Dutch-roll damped frequency parameter Abbreviatinns : boundary-layer control externally blown flap USB upper-surface blown V.G. vortex generator PERFORMANCE In a previous paper by William C. Sleeman and Arthur E. ?helps (ref. 8 ) , it was shown that good static turning could be achieved with the USB concept.

Figure 1 su~mnariz~s the static turning performance of a number of different USB configurations in terms af the ratio of normal force to thrust plotted against the ratio of axial force to thrust. The shaded band in figure 1 indi- cates representative values of static turning performance obtained with the USB concept and shows that efficiencies from about 80 to 90 percent can be obtained with high flap settings. For lower flap settings, efficiencies are generally much higher (95 percent or greater for flap angles below abodt 4 0 ° ) and turning angles are within a few degrees of the upper surface tangency angle It has been determined from previous experience that a USB configuration with marginal static turning performance caused by regions of sepa- >red or par- tially separated flow will almost certainly exhibit poor lift -1erfi -ce i r ?

forward flight. Good static turning characteristics, on the other ; , have been shown to be a reliable indicator of good lift performance in f?rwa:-d flight.

Figure 2 illustrates the effect of forward speed on the static pressure distri- bution and surface temperatures of a large-scale USB model with turbofan engines (ref. 3 ) . Static turning testc of the configuration indicated good static per- formance. As shown in figure 2, forward speed had little effect on the magni- tudes of the static pressures and surface temperatures along the engine center- line. Forward speed caused only slightly higher suction pressures and slightly cooler temperatures over the flap. Based on these results, it appears that structural and thermal design information may be determined for USB wing-flap systems on the basis of static tests which might be conducted with outdoor static rigs utilizing f ill-scale engines, nacelles, and wing-flap hardware.

Although tests have shown that the two major extsrnally blown powered-lift concepts (EBF and USB) are generally comparable in overall performance, there are some fundamental differences in the exbust jet flow fields between the two concepts at forward speed conditions. For example. 'I the FBF system the jet impinges on the lower surface of the flap and spreads spanwise, covering most , 65

of the flap span. In the USB system, however, a different, more localized flow behavior occu1:s as indicated in figurc 3. The sides, or edges, of the jet sheet produced by the engine exhaust roll up into vortices which enlarge and tend to a : . I thicken the jet as it turns over the trailing-edge flap. A number of factors I . . ' . i influence the formation of these vortices, but the ratio oi jet velocity to free- . . -. I : -; !

stream velocity and the thickness of the jet prcduced by the engine exhaust seem . , i to be the most powerful. In addition to the vortex rollup of the jet sheet,

; , . ;;I

there may be additional vortices produced by the external shape of the USB noz- . .

I: This is especially true for a sharp-cornered rectangular nozzle; the nozzle zle. I . i .

". : /

vortices can be minimized by using a well-rounded or D-shaped nczzle.

The over- * j ~ . ' ,- : : all effect of this vortex formation is to confine the jet influence to a highly I .. .: ; localized region near the jet. In fact, the jet may actually entrain free-stream 1 . . -'i . - r , 8 % .

air in such a way as to cause spanwise flow ~nboard, rather than outboard. The significance of this flow characteristic will be discussed in subsequent sections of this paper.

. I , i , -!

, * 1 .

Figure 4 presents high-lift data for a two-engine straight-wing USB configu- , . ..I !

ration (ref. 4 ) over a range of R~vrolds number and for power-off and power-on

j

cases. The data show the anticipated large influence of Reynolds number on lift

1 ..-I

for the power-off case, but for the power-on case ( c ~ = 3 ) , the Cata show very 1 1 . , I e

1 : :.!-

little effect of free-stream Reynolds number. These results have also been observed in other investigations, indicating that for moderate to high thrust ' - 1 .

coefficients, the lift characteristics of the complete configuration are pre- ! 'I.

dominantly influenced by the exhausc jet rather than the free-stream flow; this i l ? : factor may be related to the high level of turbulence of the jet as it impinges

:A

i on the wing. Fron these results, it appears that small-model data may oe used t !

i 1: with confidence in the design of ;>repulsive-lift systems.

- I Shown in figure 5 are ddta illustrating the effect of a leading-edge Krueger flap on the lift characteristics of a USB model with a high jet turning angle (Sf = 600). The trailing-edge jet tt~rning angle generates a strong upwash

i i !

field ahead of the wing, and the need for leadj-ng-edge devices for adequate pro- , -!

tection against leading-edge stall is clea:~y demonstrated by the data. As can i I , i be seen, a marked increase in maximum lift coefficient and in stall angle of !

attack resulted from the installation of a leading-edge Krueger flap.

One interesting point noted in tests of US6 configurations is that close I

: I i attention must be given to leading-edge stall in the vicinity of the nacelle.

I , " The very powerful upwash at the wing leading edge can pose serious problems ! i - when the nacelle is close to the fuselage or another nacelle (as in a four- ' . .

engine arrangement). Figure 6 illustrates this r-obiem for a two-engine i : . I straight-wing configuration and a four-cngine swept-wing configur2tion. Both ! ' .I configurations were large-scale wind-tunnel models powered by JT15D-1 turbofan I , . , engines to provide a more realistic operational envtroliment than that produced , , . ,

by small, cold jets. During tests of the two-engin<,straight-wing model (ref. 3 ) , t 1

1 - a very strong upwash field w ~ i s observed between the nacelle and fuselage during power-on conditions. Without leading-edge treatnent in this region (whjch was only about 2 percent of the wing span) the wing inboard of the jet and the entire top surface of the fuselage between the nacelles was badly stalled. Recc.1tour- ing the lowt surface of the nacelle to provide smoother flow transiti ~n and flap with blowing BLC between the nacell,. and adding a leaaing-edse Krueger

I+i -.-.

, , fuselage resulted in a significant improvement in the flow quality over the

1: ; ' -:: . I '

fuselage. The left side of figure 6 shows the lift improvements resulting from i . i the modifications to the origlczl wing.

! .

I A similar problem was ellcountered in tests of the large four-engine swept- wing configuration of reference 9 which exhibited severe separaticn along the leading edge between the fuselage and inboard nacelle and between the inboard In this case, unsweeplng the leading edges, recon- and the outboard nacelles.

touring the lower surface of the nacelles, and adding blown leading-edge Krueger flaps resulted in the improvements shown on the right side of figure 6. These data indicate a significant increase in both and stall angle of attack C L , , , , for the modified model.

!

The effects of partial- and full-span flaps on the lift characteristics of a two-engine U S B configuration are presented in figure 7. The dzta show that a large increase in lift coefficient is obtained by extending the trajling-edge flap to full span. In order to determine the proportion of this lift increment due to power effects, the data were analyzed in terms of power-induced circula- as a function of thrust coefficient, and the results tion lift cgefficient C L ~ The data of figure 8 show that the benefit of power- are presented in figure 8.

induced circulation lift on the lift of a U S B configuration with full-span flaps is mininal. Also presented in figure 8 are data for an internally blown jet flap in which the exhaust flow is distributed uniformly along the entire wing span (ref. 1 0 ) . Generally speaking, t.he localized-flow USB configuration pro- d ~ ~ c e s about 65 to 70 percent of the power-induced circulation lift available from an inzernally blown system.

Presented in figure 9 is a plot af the spanwise distribution of normal- force coefficient for the same model used to obtain the lift data of figure 8.

The data of figure 9 are presented for power-off, power-on, and engine-out con- ditions at zero angle of attack. These data indicate that the influence of the

f

exhaust jet on the wing is conta!~-d in a region extending approximately 1.75 noz- zle widths cutboard of the nozzle. In fact, for spanwise locations outboard of the 65-percent-semispan station, the power-off and power-on load distributions are very nearly the same. Of course, the actual spanwise location at which the engine-induced loads diminish to the power-off levels is configuration dcpendent, I but arrangements in which the engines are located well inboard on the wing will generally exhibit this characteristic. From figure 9 it appears-that the amount

, I I

of lift to be gained by deflecting a flap on the outboard portion of the wing 1 !

(beyond about 70 percent of the s-I ispan) is primarily the lift available from

! !

unpowered flow conditions, and it ! s not likely to be greatly influenced by power-induced effects.

Figcre 10 presents the lift characteristics of a number of USB configura- tions having different nozzle designs. Included are data for rectangular noz- zles of three different width-height ratios, a fairly high kickdown D-nozzle; a low kickdown D-nozzle with vortex generators (ref. ll), and a D-nozzle with B L i (hybrid U S B , ref. 12). The data presented have been plotted for a jet turn- ing angle of 5 0 ° . These data indicate that at low to moderate thrust coeffi- cients, such as those used on approach, there is very little difference in lift for the configurations tested. Thus, it appears that the lift characteristics

of USB configurations may be primarily a function of the jet turning pcrfor- mancr, assuming adequate leading- and trailing-edge treatment to prevent pre- mature flow separation. It has been shown (ref. 13) that the geometric nozzle characteristics which are desirable for eood turning (such as high-aspect-ratio rectangular nozzles, large kickdown angles, aild flare hngles) are detrimental to cruise performar 2 . It a,>pears, therefore, that variabl? geometry features, such as nozzle {lectors or vortex generators, or BLC ma;- be required to achieve optimum performance for bljth the high-speed and low-speed flight conditions.

In any event, the design nf a single nozzie to satisfactorily fulf<ll both the high- speed and low-spe6.d requirements represents a significant challenge to the designer.

The foregoing alscussion has centered on thc lift performance of the USB concept. However, drag characteristics are also important from an operational viewpoint. A flight envelope relating glide path, lift coefficient, thrust- weight ratio, and angle of attack is very useful in relating lift and drag to overall performance because it serves to establish power requirements and speed margins for 3 given configuration. Figure 11 presents trimmed flight envelopes for a two-engine straight-wing U S B configuration and for a four-engine swept- wing U S B configuration. The lift data for these two configurations are con- tained within the bands shown on figure 10 and are therefore generally repre- sentative of U S B configurations tested.

At the present time, there are no certified. requirements for approach per- !i formance of powered-lift airplanes. For the daza of figure 11, it is assumed 1 ;

I '

that the aircraft will fly a ?.5O glide slope a t a lift coefficient of 4.0. In the event of an engine failure, it must be pos;ible to arrest the descent with

:(i

I full power on the remaining engines without changing flap setting or lift coef- f'eient. -- From the data on the left side of figure 11 it can be seen that, for the two-engine configuration, the landing approach can be flown at a thrust- i , .

!

weight ratio of 0.21 with a stall margin of about 14O. Ic order to arrest the descent at this same flap setting, a thrust-weight ratio of 0.35 is required.

It has been found with this model and with other U S B (and EBF) models that such performance envelopes aye almost the same with one engine out as with all engines operating. Hence it can be concluded that, for this two-engine configuration, a total installed thrust-weight ratio of about 0.70 is required.

A similar analysis of the data for the four-engine configuration shown on the right side of figure 11 indicates an approach thrust-weight ratio of 0.25, and an engine-out thrust-weight ratio requirement of 0.45. In the case of the four-engine configuration, only 25 percent of the available thrust is l ~ s t in an engine failure, so the four-engine aircraft reqdires an installed thrust-weight ratio of 0.60.

Based on such analysis, it appears that somewhat higher values of thrust- weight ratio are required Idr the two-engine configuration than for the four- engine configuration, as would be expected. The data of figure 11 have been found to be generally repressntative of both U S B and EBF configurations; gen- erally, both concepts require higher installed thrust-weight ratios than inter- nally blown flap concepts (such as the distributed blowing concept which distriuutes the jet uniformly along the wing span). The simplicity of the

discrete-blowing, powered-lift systems, however, make them attractive for appli- cation to pcwered-lift STOL aircraft, as indicated by the selection of the USB and EBF concepts for use in the powered-lift prototype aircraft of 'the U.S. Air Force Advanced Medium STOL Transport program.

LONGITUDINAL STASILITY AND 'iZIM It is a well-known fact that provision of adequate trim in pitch is a seri- ous problem for powered-lift airplane configurations as illustrated in figure i2.

The data show the variation of pitching moment with angle of attack for several thrust levels at high lift conditions for a swept-wing, two-engine USB config- uration with the horizontal tail off. It should be noted that high thrust increases the longitudi~al instability as well as the diving moments of the configuration.

nction of tail lift c,3ef ficient fcllow- 14. The tail trim requirements were ficient of 8, a tail length of 3.5E. and tail size required for trim is about 37 percent of the wing area, a value nearly The use of a slotted elevator double that required by conventional airplanes.

can reduce the required tail size to about 30 percent of the wing nrea, but a considerably higher tail lift coefficient would be required to reduce the size of the tail to that for convention?l airplanes.

The foregoing data have shown that the powered-lift configurn~ion exhibits large pitching moments and that a largz tail is required for trim. It should be pointed out th?t one very important factor which must be considered in sizing the tail is the tail location. As in the case of other powered-lift systems, results in very high downwash angles, the high lift generated in the USB cmcept particularly directly behind the engines. For this reason, care must be taken in locating the horizontal. tail so as to avoid the high downwash region in which the tail could become ineffective. As an example of the flow characteristics behind a USB configuration, the variation of the downwash factor 1 - a with . f : acr ! .i ! .

for the two-engine configuration for three vertical locations of the hori- CL I . : e .

zontal tail is presented in figure 14. The data in figure 14 are for low-angle- of-attack conditions, and the increase in lift coefficient is obtained by an increase 1-1 thrust rather than by an increase in angle of attack. The data of figure 14 show that regardless of the tail location, the tail lost effective- Flow survey work for USE configurations has not been as extensive as for tions showed tha: with engines located inboard on the wing, a strong downwash z: '- 1 - r;, ., .I,

field was produced along the rear of the fuselage which made the low tail a:rangement undesirable from stability considerations. Also, it was found that

i . (

I - ; vortices shed from the wing tips and flaps did not trzil ~traight backward but , - were drawn in shar?ly toward the centerline of the airplane. At high angles

- )i

of attack, a horizontal tail located relatively far rearward and low would enter I the vortex flow and become ineffective. For this reason, the horizontal tail : 1 was generally located high and forward to retain its stabilizing contributing :, for higher thrust levels and higher angles of attack. Limited flow survey work , .

with the USB concept has demonstrated downwash characteristics similar to those' ' 1 i - .I of EBF concepts, and similar high-forward horizontal tail locations have proven desirable from longitudinal stability considerations.

In order to illustrate how the downwash data of figure 14 affect the con- tribution of the horizontal tail to stability, calculations have been made t~ determine the contribution of the tail at the high and low tail locations

%

of figure 14. The results a** shown in figure 15 tcgether with tail-off data.

The a i l contributions are based on a tail size of 35 percent of the wing area and a tail length of 3.5Z. The data show that at low lift coefficients, the high tail position provided adequate stability. The low tail, however, pro- vided very little stability at low lift coefficients and, as the lift coeffi- cient was increased by increasing power, the combined effects cf increased instability of the wing-fuselage combination and reduced tail effectiveness resulted in a very unstable configuration at high power settings.

LATERAL-DIRECTIONAL STABILITY The lateral-directional stability discussion presented herein is based on results obtained for two USB model configurations which were flight tested in the Langley full-scale ~nnel. Ptiotographs of the two models mounted for static force zests are presented in figure 16. One model, with an unswept wing and two engines mounted close inboard to the fuselage, represented a 115-scale model of the large-scale USB Aero Commander configuration recently tested in the Langley full-scale tunnel. The second model was a four-engine, swept-wing USB configu- ration. Although the two models were different in planform and englne arrange- ment, the powered-lift ~haracteristics for the two configurations were generally The lateral-directional very similar, as illustrated by the data in figure 17.

stability characteristics for the two models, however, were considerably dif- ferent, as shown in plots of the directional stability derivative Cng and the in figure 18. As shown b y the data, the effective dihedral derivative C 1 ~ swept-wing configuration had relatively large values of positive effective dihedral which increased as lift coefficier.~ increased. The directional sta- bility was also relatively high and ihere was an increase in directional sta- bility with increasing lift coefficient. The data for the straight-wing con- figuration show that the dihedral effect was relatively small and actually decreased with increasing power, but the directional stability increased very rapidly as power was applied. The differences in dihedral effect for the two configurations can probably be attributed to the differences in wing sweep angle.

The differences in directional stability were primarily a result of differences

in the effect of the engine exhaust wake on the vertical tail. The engines were located much closer to the fuselage on the straight-wing configuration than on the swept-wing configuration, and the vertical tail was influenced much more by sidewash than on the swept-wing arrangement.

In order cu illustrate how these differences in the static lateral- directional stability derivatives affect dynamic lateral-directional stability characteristics, period and damping characteristics were determlned by using three-degree-of-freedom calculations for the two models; the results are pre- sented in figure 19. The data show Dutch-roll characteristics in terms of the

daluping parameter - and the damped frequency parameter ad. The handling

T1/2 quality boundaries were taken from an AGARD bblication for STOL handling cri- teria (ref. 15). The plot on the left of figure 19 shows that the Dutch-roll oscillation for the swept-wing configuration was unstable and would be consid-

ered unacceptable with power off (cL = 1.5) or on (cL = 5 . 0 ) . In order to

achieve acceptable Dutch-roll characteristics, the damping in both roll and yaw would have to be doubled; even higher artificial damping would be required for satisfactory characteristics. In contrast to these results, the plot on the right side of figure 19 shows that the Dutch-roll mode for the unswept- wing configuration was stable and that increasing power resulted in increased Dutch-roll damping, with the result that satisfactory characteristics could be achieved without artificial stabilization.

ENGINE-OUT LATERAL TRIM One of the major problems associated with powered-lift systems utilizi-ng discrete blowing is that of restoring lateral trim in the event of the failure of one engine. This problem involves both roll and yaw, with : h e roll require- -nent being the more critical in an approach condition. One major objective of recent research on the USB concept was to determine effective means cf provid- ing roll trim for the engine-out condition. One method found to bs effectivz for the USB configuration was that of asymmetrical boundary-layer control, that t ' i e aileron of the engine- is, boundary-layer contrcl on the leading edge and on out wing. Some lateral trim datn obtained with this rnelhod of trim are shown in figures . 1 and 21 for the swept-wing and straight-wing flying models.

Figure 20 is a plbt of yawing-momeat coefficient and rolling-moment coeffi- cient plotted against lift coefficient for the four-engine, swept-wing flying model. In a four-engine operation, the rolling moments were essentially zero and a maximum trimmed lift coefficient of 10 was achieved. With the fail~re of an outborrd engine, the maximum lift coefficient decreased to about 8 and large out-of-trim rolling and yawing moments were introduced. By applying boundary- layer control to the failed engine side, it was possible to simultaneously pro- vide roll and yaw trim. Additional moments produced by spoiler deflection could then be used for maneuver contrcl.

The data of figure 21 show the engine-out rolling-moment and yawing-soment ccefficients plotted against lift coefficient for the two-engine, straight-wing 7 1

flying model. In this case, the application of boundary-layer control to the failed-engine side reduced the out-of-trim rolling and yawing moments, but it was necessary to employ spoiler deflection and a blown rudder to achieve roll and yaw trim. Additional spoiler and rudder deflection were available for maneuver control.

MODEL FLIGHT TEST RESULTS As part of the basic research program of USB jet-flap configurations, dynamic stability and control investigations have been made of the four-engine swept-wing configuration and the two-engine straight-wing configuration by using the free-flight model technique. This technique has proved to be useful in pre- vious research in pointing out problem areas which might have been overlooked in conventional testing. The swept-wiug model, shown in flight in figure 22, had a span of 3.05 m (10 ft) and was powered by four 13.97-cm-diameter (5.5 in.)

turbofan engine simulators driven by compressed air. The horizontal tail incor- porated a Krueger flap, and the elevator was deflected upward 5 0 ° . Longitudinal control was provided by deflecting the entire horizontal tail, and lateral- directional control was provided by spoilers and rudder. The two-engine straight- wing model was powered by turbofan engine simul~tors similar to those used on the swept-wing model and h ~ d a similar control system.

The free-flight technique is illustrated in figure 23. This figure shows a model being flown without restraint in the 9- by 18-m (30- by 60-ft) open- throat test section of the Langley full-scale tunnel and remotely controlled about all three axes by human pilots. Control surfaces are operated by remotely contralled pneumatic actuators. Pneumatic power and electric control signals are supplied to the model through a flexible trailing cable made up of electri- cal conductors and light plastic hoses.

The results of the free-flight model tests showed that with all engines operating and with artificial damping about the roll and yaw axes, the models were easy to fly even at lift coefficients up to 8.0. Without artificial sta- bilization, however, the swept-wing model exhibited a lightly damped Dutch-roll oscillation which made flging'difficult. The straight-wing model exhibited good damping characteristics and was easy to fly without artificial stabilization.

These results were in good agreement with the previously discussed dynamic sta- bility calculations. With one engine inoperative, the models were trimmed laterally through the use of boundary-layer control on the leading edge and aileron of the engine-out wing and through the use of spoiler and rudder deflec- tion as needed for additional lateral-directions1 trim. With artificial damping, the models were flown up to high lift coefficients, and the dynamic behavior with one engine inoperative was found to be generally similar to that for all engines operating.

CONCLUDING REMARKS The results of recent wind-tunnel investigations to provide fundamental information on the upper-surface-blown (USB) jet-flap concept demonstrated that the USB concept provides good high-lift performance. The low-speed performance appears to be mainl;? dependent upon the jet turning angle and turning efficienc ?nd on the use of proper leading- and trailing-edge treatment to prevent prema- ture flow separation. The best means of achieving good turning performance in any particular USB application must be determined from overall operational con- siderations in which high-speed performance, struct~res and noise, as well as 102-speed performance, are evaluated. The large diving moments generated at high lift coefficients can be trimmed satisfactorily with a large, conventional horizontal tail; a high tail position is best from longitudinal stability con- siderations. Large rolling and yawing moments are introduced with the loss of an engine, but these moments can be trimmed sati:;factorily through thc use of asymmetrical boundary-layer control arld throtigh the use of spoiler and rudder deflection as needed.

REFERENCES 1. P h e l p s , A r t h u r E . ; L e t k o , W i l l i a n ~ ; and Henderson, Robert L.: Low-Speed Wind Tunnel I n v e s t i ~ a t i o n o f a Semispan STOL J e t T r a n s p o r t Wing-Body With a n Upper-Surface Blown J e t F l a p . NASA TN D-7183, 1973.

2 , tlt,elps, A r t h u r E . , 111; and Smith, C h a r l e s C . , J r . : Wind-T\.,nnel I n v e s t i g a - : t i o n o f a n Upper S u t f a c e Blown J e t - F l a p Powered-Lift C o n f i g u r a t i o n . NASA TN D-7399, 1973. f 3. S t a f f o f t h e Langley Research C e n t c c : Wind-Tunnel I n v e s t i g a t i o n of t h e Aero- dynamic Eerformance, S t e a d y and i i b r a t o r y Loads, S c r f a c e Temperatures and A c o u s t i c C h a r a c t e r i s t i c s of a ~ a r g e - S c a l Twin-Engine Upper-Surface Blown J e t - F l a p C o n f i g b r a t i o n . NASA T M X-72794, 1975. I 4. S m i ~ h , C h a r l e s C . , Jr. ; P h i l p s , A r t h u r E . , 111; and Copeland, W . Latham: Wind-Tunnel I n v e s t i g a t i o n d f a Large-Scale Semispan Model With an Unswept

I

, Wing and a n Upper-Surface Blow'~ J e t F 1 . 7 ~ . NASA T N D-7526, 1974.

I r

5. Sleeman, W i l l i a m C . , J r . ; and Hohlweg, W i l l i a m C . : Low-Speed Wind-Tu. el I n v e s t i g a t i o n of a Four-Engine Upper S u r f a c e Blown Model Havind a Swept- Wing and R e c t a n g u l a r and D-Shaped Exhaust Kozzles. NASA TN P-8061. 1975.

6. Reshotko, Meyer; O l s e n , Willfam A . ; and Dorsch, Robert G . : P r e l i m i n a r y Noisei T e s t s o f t h e Engine-Over-the-Wing C o ~ c e p t .

I. 30'-60' F l a p P o s i t i o n . . $SA ' j ; ' C .

T M S-68032, 1972.

7. Reshotk?, Meyer; O l s e n , William A . ; and Dorsch, Robert G . : P r e l i - i n a r y T e s t s o f t h e Engiae-Over-the-Wing Concept. 11. lo0-20° F l a p P o s i t i o n .

NASA T M X-68104, 1972.

8. Sleeman, W i l l i a m C . , J r . ; and P h r l p s , A r t h u r E . , 111. : Flow-Turning Performance. Powered-1,ift Aerodynamics and A c o u s t i c s , NASA SP-406, 1976. (Paper no. 2 o f t h i s c o m p i l a t i o n . ) 9. Koenig, Cavid C . ; and Aoyagi, K i y o s h i : Maximum L i f t of Upper S u r f a c e Slowing STOL A i r c r a f t With Swept Wings. A I A A Paper No. 75-868, J u n e 1975.

I I 10. Lowry, John G . ; R i e b ~ , Sohn M . ; and Campbell, John P . : The Jet-Augmented F l a p . P r e p r i n t Nc. 7i5, S.?I.F. rund P a p e r , I n s t . Aeronaut. S c i . , J a n .

1957.

I 11. Wimpress, John K . : Upper S u r f a c e Blowing Technologv a s Applied t o t h e Yi-14 A i r p l a n e . [ ~ r e ~ r i n t l 730916, Soc. Automot. Eng., Oct. 1973.

C a r r o s , Robert J . ; C o i s s e v a i n , A l f r e d G . ; and Aoyogi, K i y o s h i : Aerodynamic C h a r a c t e c i ~ t i ~ 5 of a Large-Scale Hybrid Upper S u r f a c e Blown Model Having Four Engines. NASA T M X - f 2460, 1975.

o QCSEE. D-NOZZLE IUNPUBLI'SHED) O B O E I N G Y C - 1 4 (REF. 11) A RECT., A = 2 + DEFLECTOR (REF. 4 ) d RECT., A = 4 1.0 Y a R I D (REF. 12) -8 .o F~

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T .2 0 .2 .4 .6 8 1.0 Fiylre 1 . - Static turning characteristics of several US3 c-nfigurations.

-

W I N D OFF

----

W I N D ON SURFACE TEMPERATURE Figurc < . - pressure and temperatLre profiles.

Figure 3 . - USB flow c.har,~cterist i c s .

C 0 C 3 P C: REYfUOLDS Nllh\BER RFY NOLDS NUh\BER FIgurc 4 . - tf f c c t nf Hryualds ~~llmhcr on - i f t c l ~ : l r n c t c ; l s t i ~ ~ for two-cnginc straight-win;, llSB conf igurot ion.

LEADING-EDGE KRUEGER FLAP LIFT COEFFICIENT, '4 "CLEAN LEA31NG EDGE L 3 L Figuie 5.- Effect of leading-edge high-lift devices on lift. 6, = 60°; C , , = 2.0.

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B A S I C ' h103;:iEO j v I I I I I -10 0 10 20 30 40 -10 0 10 20 31) 40 a, deg a, deg .gure 6.- Leading-edge treatment for two-engine straight-wing and four-engine swept-wing configurations.

rCs-, ,r B LC

FULL- "A\ SPAN

P A R T I A L - SPAN F l AP FLAP

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LIFT COEFFICIENT, 0 1 2 3 4 THRUST COEFFICIENT. Cv

Figure 8 . - Power-induced l i f t c h i ~ r a c t e r ist i c s cf two-t>ngine

sti..:::-.ht-wine LISB configuration. c F f = 60°; iu = O0.

FUSELAGE CENTERLI NE SECTION N O R M A L FORCE COEFF I C I ENT.

C" NOZZLE NOZZLE I I 1 I I I

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( R I G H T W I N G ) !LEFT W I N G Figure 9.- Spanwise normal-force distribution for two-engine straight-wing USB configuraiion. B f = 6 0 ' ; a = OO.

THRUST COEFFICIENT C ,, NOZZLE L I F T 6 0 RECT., A = 2 + DEFLECTOR

COEFF I C IENT. 1

(REF. 4 ) 0 RECT., A = 4 0 RECT., A = 6 QCSEE D-NOZZLE (UNPUB b YC-14, D-NOZZLE W I T Y V.Gis (REF. 11) D HY BR I D (REF. 12) 2 - fi RECT.. A = 6 (REF. 31 O RECT,, A = 6 (REF. 5 ) 0 I I I I

-10 o 10 2 0 30 d o

a, deg Figure 10.- Effect of nozzle geometry on lift for a number of USB configurations.

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. 5 PLA I N ELEVATOR SLOTTED ELEVATOR .4 T A I L SIZE REQU l RED .3 FOR TR lM HOR IZONTAL-TA I L AREA W I N G AREA .2 CONVENT1 O N A L TA I L L I F T COEFFICIENT. CL, TA I Figure 13.- Horizontal-tail requirements.

CL,wing = 8.0; l t = 3.5c; 10-percent static margin.

DOWNLYAS FACTOR.

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1 - - da L I F T COEFFICIENT. C L f : Figure 14.- Variation of downwash factor. A f = 6g0.

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HOR IZONTAL-TA I I AREA - 0+35 \'/I NG AREA tONCI TUDINAL STJ 9ILETY PARAMETER. 0 d C m - STABLE & -. 05 L I F T 20EFF1ClENT, C L Figure 15.- Static longitudinal s t a b i l i t y .

F o u r - e n g i n e swept v iilg Figure 1 6 . - Photographs of two-engine and four-engjne UbS modzls i n s t a l l e d i n the L a n g l ~ y f u l l - s c a l e runnel.

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L k;, 0, -10 0 10 20 30 4 0 a, deg Figure 17.- Lift characteristics of two-engine and four-engine US5 models.

4-ENGINE SWEPT WING, 2-ENG INE STRAIGHT WINC lt = 69' B f = 65 EFFECT1 VE D l HEDRAL,

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0 2 4 6 0 2 4 6 L I F T COEFFICIEF!T, CL Figuie 18.- Lateral-directional stability characteristics of two-r.~gine and four-engine USB models.

L I F T C O F i F I C I E N T 0 1.5 0 5.0 0 5.0 (ROLL AND Y A W D A M P I N G DOUBLED) A 5.0 (ROLL .4ND YP.W D A M P I N G TRIPLED) 2-ENG INE STRAIGHT WING. b f = 6 9 ' 4-ENGINE SWEPT WING. b f = 65' -4

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0 - 4 .8 1.2 0 -4 .8 1.2 FREQUENCY PARAMETER, wd Figure 19.- Dutch-roll characteristics of two-engine 92d four-engine USB models.

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LIFT COEFFICIENT. CL Figure 21.- Lateral trim characteristics for two-engine straight-wing USB model.

FTgure 22.- Photograph of f o u r - e n g i n e s w e p t - w i n g free-flight US& model in the Langley full-scale t u n n e l .

()111c,INAL PAGE rS

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APPLICATION OF POWERED-LIZ CONCEPTS FOR IMPROVED CRUISE FE'FiCIENCY OF LONG-RANGE AIRCRAFT

Paul L. Coe, Jr. , and Paul G. FG .,'nier

N A S A Langley Research Center S b W I The p r e s e n t paper summarizes r e s u l t s of r e c e c t stcldies conducted a t t h e NASA Langley ~ e s e a r c h Center t o e x p l o r e t h e use of powered-lift concepts £01.

improved Iew-speed performance of long-range subsonic and s u p e r s o n i c c L L 3 i s ~ v e h i c l e s . The r e s u l t s i n d i c a t e t h a t powered l i f t can provide s i g n i f i c a r ' t improvementb i n low-speed ~"rfo-rmanc:, a s b e l l a s s u b s t a n t i a l increase:.

I 1

I c r u i s e e f f i c i e n c y and ; m g e jr b o t h s u b s o . ~ i c and s u p e r s o n i c c r u i s e 1 c o n f i g u r a t i o n s .

INTRODUCTION -.

ltte NASA Langiey Kesearch Center is c u r r e n t i y i n v e s t i g a t i ~ i g t h e iise of powered-lift concepts f o r irrproved low-speed performance o; long-rana,e sub- T h i s r e s e a r c h h a s been d t r e c t e d toward s o n i c and s u p e r s o n i c c r u i s e a i r c r a f t . L concept, which may provide s u b s t a n t i a l i n c r e a s e s i n l i f t f o r iaproved rake-off I and l a n d i n g performance and, f u r t h e r , which may provide b e t t e r engine-airframe : i m a ~ c h i n g f o r improved c r u i s e e f f i c i e n c y and range.

, T . ' The p r e s e n t paper sumn~arizes r e s u l t s of recent. s t u d i e s of powered-lift concepts, conducted i n tZle Langiey V/STCL and .!-scale t u n n e l s . I n p a r t i c - u l a r , t h e paper d i s c u s s e s (1) t h e a p p l i c a t i o n o t t1.e wer-i1.2-wing blowing (O'IWB) concept t o an advanced subsonic c r u i s e con<lv,it:?tlon and ( 2 ) t h s d p p l i - c a t i o n of OTWB and t h r u s t v e c t o r i n g concepts t o an a:ivanced s u p e r s o n i c c r u i s e c o n f i g u r a t i o n .

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height of engine nozzle above wing (see fig. 5) incidence of canard tail area aircraft weight angle of attack flap deflection angle .%breviat ions : I i boundary-layer concrol i I over-the-wing bl~wing upper-surface blowing POTENTIAL BENEFITS DERIVED FROM POWERED-LIFT CONCEPTS One of the fundamental considerations in the design of a cruise efficient aircraft is the sizing of the configuration with regard to wing area and installed thrust requirements.

It is recoznized that the sizing process fnvolves considerable compromise, and that low-speed performance plays a key part in the trade-off.

Presented in figure 1 is a classical "thumb print" plot which shows the variation of range with installed thrust-to-weight ratio T/W and wing load- ing W/S. Also shown in this figure is a typical take-off field length con- straint which emphasizes the iinpact that low-speed performance has on engine- airframe sizing.

The important point illustrated by figure 1 is the fact that for a specified configuration, optimum range is obtained with relatively low values of T/W and relatively high values of W/S, and that increasing T/W ; !

or reducing W/S from the optimum values in order to m e t the take-off field Figure 2 illustrates the influence of take-off lift coefficient, T/W, and W/S on take-off field length requirements. This relationship was obtained from an empirical study and is discussed in detail in reference 1.

t is seen that a specified take-off field length can be obtained, with relatively low values of T/W and relatively high values of w/S, provided that sufficiently high values of take-off lift coefficient can be obtained.

Therefore, the successful application of powered-lift concepts, which yield lmprovea low-speed performance, wiil allc.; acceptable take-off field lengths to h values of T/W and W/S sized to obtain optimum cruise POWERED-LIFT CONCEPTS INVESTIGATED The powered-lift concepts considered herein are described and discussed individually. Although the details differ, the fundamental consideration is the same for both subsonic and supersonic cruise vehicles; namely, to allow the wing area and installed throst to be sized to provide optimum cruise rffi- ciency whilt using powered-lift concepts to meet the low-speed operational requirements associated with conveational aircraft.

One particularly promising povzred-lift concept, which may have near-term applications for long-haul sdbsonic transports, is over-the-wing blowing (On>).

Figure 3 shows a photograph and a sketch of +he concept applied tc a subsonic transport configuration with an aspect-ratio-7.48 wicg and c leading-edge sweep of 33.6O. The conf;guration is equipped with four, pylon-mounted, upper-surface engines with deflectable exhaust nozzles. Reference 2, which combined the ana- lytical resul.ts of reference 3 and the experimental results of references 4 and 5, has shown that the OTWB concept with undeflected exhaust nozzles can pro- vide substantial reductions in induced drag. The reduction in induced drag is provided by the jet exhaust which induces an u~wash on the wing. The upwash rotates the wing force vector forward and effectively produces a negative incre- ment in induced drag. Therefore, if it is possible to produce addiiional circu- lation lift by deflection of the exhaust flow downward onto the wing surface during take-off and landing, such an arrangement wouid provide not only improved low-speed performance but also improved cruise performance.

As was sentimed pr!viously, powertd-lift concepts h ~ v e also been applied tc superson-c cruise veh~cles. Figure 4 shows a photograph of a large-scale advanced sitperzcnic cruise arrow-king configuratior. which has an aspect ratio of 1.72 and au i!,board leading-edge sweep of 7 4 O .

The powered-lift concepts investigated for improved low-speed performance of this configuration are also sketched in figure 4 and include (1) bouniary-layer control (BLC) for enhanced flap cffectivenesa ?nd prevention of flow separation at high flap deflections, (2; OTWB for additional circulation lift (this concept also has another advan- tage in that the trailing-edge flap system may be continuous, rather than the segmented system necessitated by the use of the conventional underslung engines), and (3) thrust vectoring which provides increased lift by a combination of the direct vector ~Impcnent of thrust arid by the additional circulation lift pro- duccd by flow zriL r x nmen t .

LIFT CHARACTERISTICS Subsonic Cruisc Vehicles Figure 5 shows the variation of additional circulation lift CL,~, obtained with the OTWB concept, as a function of the racio of the height of the engine h/D. The data are presented for a = OC above the ving to the engine diameter and bf = 4 5 ' . It should be noted that the exit nozzle deflection varied with so that the jet would impinge at approximately the same chorawise location.

h/D From figure 5 it is seen that, as with other powered-lift concepts, relatively small values of thrust coefficient result in significant levels of addi- C,, tional circulation lift and that further incrGases in result in more gradual Cu

increases in CL, r. Furthermore, from figure 5 it can be seen that there is only

a slight increase in C L * ~ . as h/D is increased from 0.5 L 1.0.

Figure 6 presents a comparison of the lift characteristics ~btained for the (USB) concept applied to a OTWB concept with those for an upper-surface blowing configuration comparable with that used in the OTWB investigation (see ref. 6 ) .

The USB concept used rectangular exhaust nozzles having as aspect ratio of 6.

On the basis of a comparison of the data for the OTWB and USB concepts, it would appear that both concepts produce essentially the same level of additional circulation lift. However, the reduction in induced drag provided by the OTWB concept in the cruise configuration indicates that such a configuration may have a higher level of cruise efficiency than a tor-figuration with the USB concept.

In light of these considerations, the KASA Langley Research Center will be conducting tests with a large-scale model of the advanced OTWB subsonic trans- port configuration sllon in figure 7 . The configuration uses a supercritical airfoil with an aspect-ratio-12 wing arid is designed for efficient cruise at Mach numbers of about 0 . 8 . Curing the take-off and landing phases of flight, the exhaust is deflected downward onto the wing surface to provide the desired high lift for improved law-speed performance.

Supersonic Cruise Vehicles Figure 8 summarizes the improvements in lift obtain'ed with the various powered-lift concepts investigated for the advanced supersonic cruise vehicle.

From figure 8 it is seen that the increment in lift provided by the plain trailing-edge flap is reduced for flap deflections above 200, as a result of flow separation at the higher flap deflections. As would be expected, the application of BLC provides enhanced flap effectiveness and eliminates flap stall for flap defiections up to 4 0 ' . Figure 8 also shows that thrust vector- ing provides an additional increment in lift; however, this increment was limited to the vector component of thrust. The fact that thrust vectoring failed to ?tovide additional circulation lift is attributed to the relatively far aft position of the exhaust nozzles for the particular engine location considered. The data show further that the OT,.!B concept provides substantial additional increases in lift. This result is attributed to the continuous

I trailing-edge flap system, permitted by the upper-surface-mounted engines, and the additional circulation lift produced by the concept.

The potential benefits obtainable from the application of powered-lift concepts to the supersonic cruise vehicle are illustrated in figure 9 . The relatively low value of lift curve slope, associated with the low-aspect-ratio, highly swept, arrow wing, and the tail scrape angle are seen to constrain the I take-off lift coefficient of the basic concept to values of only about 0.55.

Furthermore, the relatively high angle of attack associated with this lift coefficient results in substantial drag which penalizes the low-speed perfor- mance. In addition, the relatively high rotation angle requires the use of a visor nose for acceptable pilot visibility and also requires an elongated land- ing gear installation which results in a weight and volume penalty.

Results of airframe-engine sizing studies have indicated that significant improvements in supersonic cruise efficiency and range can be obtained by reduc- ing the wing size of this ccnf3-guration by about 25 percent. Therefore, to obtain acceptable take-off field lengths, the resized vehicle would require approximately a 25-percent increase in lift coefficient, which corresponds to CL 2 9.7.

Figure 9 shows that thrust vectoring in combination with BLC provides the desired lift coefficient of 0.7 at a 2 fjO, whereas OTWB and BLC provides a lift coefficient of 0.7 at a : l.SO. The use of OTWB or thrust vectoring therefore permits operation at reduced angle of attack which would result in a significant reduction in drag and thereby provide improved low-speed performance. Further- more, the reduced angle of attack would allow a reduction in landing gear length and may also eliminate the requirement for a visor nose, both representing a significant weight savings. However, the most significant point is that with the increased value of lift, the wing size can be reduced toward the optimum size for increased supersonic range while maintaining acceptable take-off and landing performance.

It should be noted that the OTWB concept appears to provide better low- speed performance than the particular thrust vectoring concept investigated.

However, it is constdered that, through proper design, the thrust vectoring concept might be as efficient as the OTWB concept and that the thrust vector- ing concept may offer some advantages over other high-lift concepts for achiev- ing improved lateral control.

For example, one promising thrust vectoring concept which uses a two-dimensional exit nozzle design is illustrated in fig- ure 10. Such an arrangement should provide improved powered-lift characteris- tics and also allow significant, and needed, improvements in roll control by the introduction of drfferential thrust vectoring.

LONGITUDINAL TRIM It should be noted that the data presented in the previous section corre- spond to untrimmed values of lift coefficient, and that extremely large nose- down pitching moments accompany the increases in lift provided by the powered- lift concepts. The problem of providing pitch trim for subsonic transport

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configurations has been discussed in reference 7; therefore, the present dis- cussion is limited to the problem of providing pitch tria; for supersonic cruise .- , ,-ZCI .

vehicles.

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Figure 11 shows the pitching moments produced by application of powered- ! . .+/ ..

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lift concepts to the supersonic cruise vehicle. Relative merits of various f %.

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means for providing pitch trim have been investigated for this configuration

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and the results are praented in figure 12. The analysis was conducted for a - . . . . . .-,,* 2 trim lift coefficient of 0.7 and a static margin (d%/dcL) of 3 percent. As . . " 3 would be expected, the use of a conventional aft tail for trim requires a down- - .- '.I

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load, whereas the canard concepts require an upload. Furthermore, both the con- , : .-.$ ,.. ' venticlal aft tail and the fixed canard require relatively high values of tail I , ..e4

lift c ,efficient ( c ~ , ~ ~ ~ ~ ) and would therefore probably require a sophisticated

hlgh-lift system.

It is possible to achieve the favorable upload of the canard and to elim- I ! - I inate the requirement for a high tail lift coefficient by introducing a gearing

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so that the canard angle of attack is reduced as the aircraft angle of attack .I ! i .

is increased. It should be poiated out that the eared canard requires the lowest lift coefficient per tail area ratio (St/Sf and may not require the sophisticated high-lift devices which would be associated with either the fixed canard or the conventional aft tail.

CONCLUDING REMARKS The application of powered-lift concepts to advanced long-range subsonic and supersonic cruise vehicles appears promising for providing significant improvements in low-speed performance. The increased lift provided by the powered-lift concepts allows a reduction in both wing size and installed thrust requirements which yields a better engine-airframe match for improved cruise I . . : efficiency and range. The powered-lift benefits appear to be particularly sig- * i.

nificant for the supersonic cruise vehicle because of the inherently poor low- speed lift characteristics associated with the low-aspect-ratio, highly swept wing required for supersonic flight.

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REFERENCES Weirich, Robert L . : Analytical Determination of the Take-Off Performance of Some Representative Supersonic Transport Configurntions. NASA TN I ! - 2 3 0 8 , 1964.

Bower, Robert E . : Opportunities for Aerodynamic-Drag Rcd~~ction.

NASAfUniversity Conference on Aeronautics, NASA SP-372, 1975, pp. 323-352.

Putnam, Lawrence E . : An Analytical Study of the Effects of Jets Located More Than One Jet Diameter Above a Wing at Subsonic Speeds. NASA ' i ' N D-7754, 1974.

Putnam, Lawrence E . : Exploratory Investigation at Mach Numbers Frcm C.40 to 0.95 of the Effects d i Jets Blown Over a Wing. NASA TN D-7367, 1973.

bald, B . : Airframe-Engine Interaction for Engine Configurations Mounted Above the W?ng. Part 11: Engine Jet Simulation Problems in Wind Tunnel Tests. Airframe/?ropulsion Interference, ACARD-CP-150, Mar. 1975, pp. 26-16 - 26-32.

Sleeman, William C., Jr.; and Hoblweg, William C.: Low-Speed Wir..-Tunnel Investigation of a Four-Engine Upper Surface Blown Model Having a Sdeyt Wing and Rectangular and D-Shaped Exhaust Nozzles. NASA TN D-8061, 1975.

Phelps, Arthur E., 111; Johnson, Joseph L., Jr.; and Margason, Rieha-d J.: Summary of Low-Speed Aerodynamic Performance of the Upper-Surface blown Jet-Flap CL\ncept Powered-Lift Aerodynamics and Acoustics, NASA SP-406, 1976. (Paper no. 4 of this compilation.)

CRUISE TA KE-OFF AUD lAND I NG \ Figure 3. - OTWR concept npp 1 i e d t o srlhsonic transport.

PLAIN FLAP + & L C

OVER-THE-WING BLOWING THRUST VECTORING Figure 4 . - Powered-lift concepts i n v c a t l ~ n t e d lor s u p c r s o ~ l i c c r u i s e conf igurnt ion.

Figilre 5 . - Additional c i r c u l a t i o n l i f t produced by OTWB for subsonic c r u i s e configuration. €if = 4!i0.

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Figure 6 . - Comparison o f low-speed I l f t c h t \ r a c t e r i s t i c s f o r OTWB and USB concepts applied t o subsonic c r u t s e conFiguretions.

--- Figure 7.- Sketch of advanced long-haul subsonic transport with OTWB.

OVER-THE-W I N G BLOWING + BLC VECTOR I N G + BLC Figure 8.- Lift improvements due to powered-lift concepts for advanced supersonic cruise vehicle. a - O O .

OVER -THE -W I N G BLOWING + BLC r- THRUST VECTOR I N G + BLC a, deg Figure 9.- P o t e n t i a l b e n e f i t of a p p l i c a t i c n o f p o s e r e d - l i f t concepts t o supersonic c r u i s e v e h i c l e . 5 f = 30".

Figure 10.- Ske. . I of r12vised thrust vectoring concept for a w r s o n l c c r u i s e v e h i c l e .

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COMPARISON OF AERODYNMIC THEORY AND EXPERIMENT I 3 FOR JET-FLAP WINGS Thomas G. Gainer, Long P. Yip, and Raymond D. Vogler NASA Langley Research Center SUMMARY This paper compares aerodynamic theory predictions made for a jet-fl.'pped wing with experimental data obtained in a fairly extecsive series of tcasts in the Langley V/STOL tunnel.

The predictions were made with the E l ? ) (Elementary Vortex Distribution) program developed by Lopez, Shen, and Wassun at McDonnell- Douglas. The tesls were made on a straight, rectangular wink and investigated two types of jet .flap concepts: a pare jet flap with high j e t deflectio~~ and a wing with blading at the knee of a plain trailing-edge flap. The tests inves- tigated full- and partial-span blowing fot wing aspect ratios of 8.0 and 5 . 5 znJ momentum coefficients from 0 to about 4 .

I b e total lift, drag, and pitching-moment coefficients predicted by the theory wece in excellent agreement with experimental values for the pure jet The pressure caefficients on the wing, flap, even with the high jet deflection.

and hence the circulation lift coefficients, were underpredicted, however, because of the linearizing assum>tions of the planar theory. The lift, drag, and pitching-moment coefficients, as well as pressure coefficients, were under- predicted for the wing with blowing over the flap because of the failure of the theory to account for the interaction effect of the high vt-locity jet passing over the flap.

Jet-flap theory is a relatively simple powered-lift theory developed by assuming that the jet exhaust that augments lift leave9 the wins trailing edge at small angles as a thin sheet. The theory was first developed in two dimen- I), then in three dimensions by Maskell and Spence and sions by Speuce (ref.

others. (See ref. 2.) More recently, lifting-surface programs patterned after those for conventional wings have been developed that can predict chordwise and spanwise loadings for complex wing p1an:orms and arbitrary distributions of momentum coefficient and jet rieflection.

These programs include the EVD (Ele- mentary Vortex Distribution) program (ref. 3) and the Vortex-Lattice Program for Jet-Flapped Wings (ref. 4 ) .

Although the basic a~suntptions somewhat restrict t h ~ theory, it could have important applications. LesipnerJ are examining the jet-flap concept, fcr example, in connection witti the two-dimensional nozzles being considered fcr 167- PAPE'INTENTIONALLY BUNK

advanced supersonic aircraft. These nozzles would be mounted at the trailing

edge of the wing and could be deflected to provide lift augmentation - either

to imprljve performance or, in the case of fighters, to improve maneuvering; since these nozzles spread the exhaust into a fairly thin sheet, jet-flap the- ory would apply in their design. For many STM, applications, flow conditions may be outside the strict limitations of the theory; nevertheless, there are indicaeions jet-flap theory could be used. Although the theory is based on small-disturbance concepts, the theory predictions have agreed for some cases with test data at high deflections.

Jet-flap theories have also predicted aero- dynamic characteristics of other configurations such as the augmentor wing and the externally blown flap (ref. 5 ) .

The different applications of the theory have not been examined in detail, however, nor have the theories themselves been verified to any greac extent because the necessary experimental data have not been availbble. For most of the powered-lift data available, the distributions of momentum coefficient and jet-deflection angle are not defined well enough to use in theory predictions.

The data that ha-~e these distributions defined are limited to just a few blowing spans and jet deflections. Detailed pressure distributions are not generally available for comparison with theo zt ical. predictions.

To provide some of the necessary data, Langley Research Center conducted a series of wind-tunnel tests that investigated a fairly wide range of jet-flap parameters. This paper compares predictions made with a representative jet- fiap theory, namely the EVD theory (ref. 3 ) with these experimental data. The test model had a straight, untapeted wing. It was tested with two powered-lift configurations which, while they do not quite agree with the assumptions of the theory, would be of interest in STOL applications. In one configuration the wing was equipped with a pure jet flap with high jet deflection, whereas in the other, the wing was equipped with blowing over a plain trailing-cdge flap. Partial- and full-span biowing and two wing aspect ratios (8.0 aild 5.5) were investigated.

The model was tested through an angle-of-attack range from about -4O to 20° at momentum coefficients from 0 to about 4.

aspect ratio influence coefficient relating vorticity at a point j to downwash at a point i span net drag coefficient, based on model drag minus component of model thrust in drag direction lift coefficient Cr, jet-reaction lift coefficient C~,jr

circulation lift coefficient pitching-moment coefficient (referred to the wing apex) jet-momentum coefficient section pressure coefficient difference between upper and lower surface pressures section jet-momentum coeffjcient axial force normal force constant in thickness correction factor thickness correction factor wing area, m2 (it2) blown area of wing (blown span times wing chord), m2 (ft2) win^ thickness-chord ratio induced downwash at a control point i chordwise distance spanwise distance angle of attack, deg flay deflection, deg jet turning anble, deg thrust efficiency factor; actual thrust divided by nominal calibrated thrust

MODEL AND APPARATUS The t e s t model is shcwn i n tile Langley V/STOL t u n n e l i n f i g u r e 1. The fuse- l a g e was designed s m a l l enough t o have n e g l i g i b l e e f f e c t on t h e wing aerodynamic c h a r a c t e r i s t i c s , y e t l a r g e enough t o c o n t a i n t h e model b a l a n c e , p r e s s u r e gages, and a s s o c i a t e d t u b i n g r e q u i r e d i n t h e t e s t s .

Figure 2 snows some o f t h e wing d e t a i l s . The wing had a 25.4-cm (10 i n . ) cho.J and an NHCA 0018-64 a i r f o i l s e c t i o n .

It was t e s t e d a s an a s p e c t - r a t i o - 8.0 wing (2.03-m (80 in.) span) w i t h blowing o v e r t h e f u l l span, two-thirds of t h e span, o r o n e - t h i r d o f t h e span; t h e n t h e outboard o n e - t h i r d of t h e wing was removed and i t was t e s t e d a s a n a s p e c t - r a t i o - 5 . 5 wing (1.02-rn (40 i n . ) span) w i t h f o l l - o r half-span blowing. The leading-edge s l a t shown was used t o pre- vent s e p a r a t i o n a t t h e l e a d i n g edge a t h i g h a n g l e s of a t t a c k and h i g h j e t d e f l e c - t i o n s . A i r f o r blowing was provided by t h e t u n n e l h i g h - p r e s s u r e a i r s u p p l y which was brought i n througn t h e s t i n g .

For t h e p u r e j e t f l a p ( f i g . 2). t h e wing remained i n its b a s i c a i r f o i l shape, w i t h t h e t r a i l i n g edge u n d e f l e c t e d , and a i r was e j e c t e d from a s l o t on t h e lower s u r f a c e at t h e t r a i l i n g edge. The a i r was e j e c t e d a t a n a n g l e af approximately 600 w i t h r e s p e c t t o t h e wing chord l i n e .

For blowing o v e r t h e f l a p ( f i g . 2 ) . t h e p a r t of t h e wing c o n t a i n i n g t h e jet f l a p was removed and r e p l a c e d w i t h a 25-percent-chord d e f l e c t a b l e p l r - . A f l a p .

A " s l o t " a t t h e knee o f t h e f l a p , c o n s i s t i n g of 300 h o l e s 0.159 crn (0.063 i n . ) i n d i a m e t e r and e q u a l l y spaced i110ng t h e span, provided t h e a i r f o r blowing.

The f l a p was d i v i d e d i n t o t h r e e spanwise segments f o r d i f f e r e n t amounts of p a r t i a l - s p a n blowing, and o n l y t h e f l a p segment a l o n g t h e blowing span was d e f l e c t e d . For example, v i t h 1/3-span blowing, o n l y t h e inboard 1/3-f l a p seg- ment was d e f l e c t e d and t h e remaining two outboard segments were u n d e f l e c t e d .

F l a p d e f l e c t i o n s of o0, 15O, 30°, 45O, and 6C0 were i n v e s t i g a t e d .

Model f o r c e s and m m e n t s were measured w i t h a six-component s t r a i n - g a g e b a l a n c e .

S t a t i c p r e s s u r e s were measured a t s i x spanwise s t a t i o n s on t h e r i g h t

wing p a n e l (& = 0.15, 0.30, 0.45, 0.60, 0.78, and 0.93 and, as a check on

)

symmetry, a t one s t a t i o n (& 0.30) on t h e l e f t p a n e l . There were 31 o r i f i c e s on t h e wing a t each s t a t i o n - 19 on t h e upper s u r f a c e and 12 on t h e lower s u r f a c t THEORETICAL CALCULATIONS The t h e o r e t i c a l c a l c u l a t l o n s were made w i t h t h e Elementary Vortex Distri- b u t i o n ( E m ) program d e s c r i b e d i n r e f e r e n c e 3. The E M program is R l i f t i n g - s u r f a c e program t h a t r e p r e s e n t s t h e wing and j e t wake w i t h a v o r t e x s h e e t o f v a r y i n g i n t e n s i t y . The v o r t e x s t r e n g t h on t h e wing is determined by s a t i s f y - i n g t h e tangent flow boundary c o n d i t i o n on t h e wing. T h i s is done by s e t t i n g t h e sum of t h e induced v e l o c i t i e s ( f i g . 3) e q u a l t o t h e comFonents of wi

. .: 7 free-:tre.am v e l o c i t y normal t o t h e wing s u r f a c e so t h e r e can b e no flow through t h e wirlg surface. The v o r t i c i t y along t h e jet wake is determined by using t h e basic S ~ e n c e r e l a t i o n s h i p t h a t expresses v o r t i c i t y i n terms of s e c t i o n momentum coeff5.c e n t and t h e change i n induced downwash with respect t o downstream d i s - tance r 4 follows: Tile program has been l i n e a r i z e d by assuming small p e r t u r b a t i o n s and t h a t ,111 * r n t i c i t y l i e s i n t h e plane of t h e wing. The boundary conditions have been projta.: ed back t o t h e plane of t h e wing. The camber, t w i s t , and j e t - d e f l e c t i o n an:Sic!s a r e assumed t o be small.

T?,e EVD program adds a degree of s o p h i s t i c a t i o n t o t h e b a s i c vortex scheme b:r assllming a continuously varying chordwise v o r t i c i t y constructed from d i f f e r - el.:: types of b a s i c vortex elements. The E V D uses overlapping t r i a n g u l a r ele- meuts t.o obtain a l i n e a r l y varying v o r t i c i t y between two p o i n t s on t h e wing, a of lown nut ream distance, is used t o represent t h e t r a i l i n g j e t sheet f a r dcwist ream.

T!le EV3 program accounts f o r wing camber and t w i s t and allows f o r a trailing-,edge f l a p , but t h e assumption is made t h a t t h e jet is emerging from t h e trai!ing edge of t h e f l a p , not from a point on t h e upper s u r f a c e cf t h e wing. Thle progrr assumes a t h i n wing; however, thickness e f f e c t s c m be accounted f o r by multiplying c i r c u l a t i o n l i f t and wing pressure c o e f f i c i e n t s by t h e Eollowing c o r r e c t i o n f a c t o r : 1.144 w i t ' 7 f . " ' l - s p a n blowing t o 1.048 with 113-span blowing.

1.n m a ~ i n g t h e E V D c a l c u l a t i o n s , EVD elements were placed a t 20 spanwise s t a t i o n p along t h e semispan f o r t h e aspect-ratio-8 wing and a t 16 spanwise s t a t i c ~ t l s f o r t h e aspect-ratio-5.5 wing.

There were s i x chordwise elements on thb- ding -and f i v e on t h e jet.

The aomentu? c o e f f i c i e n t and j e t - d e f l e c t i o n angles needed t o make the cal- were t 2asurc:d normal t o and along the wing cl;ord l i n e and resolved i n t o a

r e s u l t a n t f o r c e , An e f f i c i e n c y f a c t o r q, which is t h e r a t i o of t h e a c t u a l o r r e s u l t a n t t h r u s t t o t h e nominal t h r u s t , w 3 s t h e n determined, as was t h e jet d e f l e c t i o n o r t u r n i n g a n g l e tij, which is t h e a n g l e between t h e r e s u l t a n t t h r u s t and t h e wing c h o r d l i n e . The v i n g w i t h blowing o v e r t h e f l a p had e f f i c i e n c i e s varyitcg from about 0.85 at bf = 0' t o between 0.75 and 0.79 a t 6f = 600.

The t u r : t n g a n g l e w i t h blowing o v e r t h e f l a p a a about e q u a l t o t h e a n g l e o f t h e f l a p upper-surface d e f l e c t i o n (6f + 13.507. The p u r e j e t f l a p was c n l i - b r a t s d t h e same way a s t h e blown f l a p , e x c e p t t h a t it was c a l i b r a t e d w i t h t h e jet f l a p i n p l a c e s o t h a t tire nominal t h r u s t was e q u a l t o a c t u a l t h r u s t and its e f f i c i e n c y f a c t o r was 1.0. The s t a t i c r e s u l t s show t h a t t h e j e t - d r f l e c t i o n -; x, , , > . > : 4 a n g l e f o r t h e p u r e jet f l a p was between 61" and 63O. I n t h e c a l c u l a t i o n s , t h e > . +

:.+..;i $$j

momentum c o e f f i c i e n t and j e t - d e f l e c t i o n a n g l e s were assumed t o be uniformly .. . . -.&, ;,.. : d i s t r i b u t e d o v e r t h e blowing span.

I : . , ~. ' $-@

. .

, i ' . , - ; x RESULTS AND DISCUSSION Pure J e t Flap Compnrison of t o t a l l i f t , d r a ~ , and p i t c h i n g moment.- The l i f t , d r a g , and p i t c h ~ 8 - n r a m e n t comparisons shown i n f i g u r e s . 5, 6 , and 7 i n d i c a t e e x c e l l e n t agreement between t h e o i y and experiment f o r t h e pure-jet-f l a p wing. The r e s u l ts f o r f u l l - s p a n blowing w i t h tile a s p e c t - r a t i o - 8 wink ( s ~ = 61.4') a r e given i n f i p - u r e 5; t h e r e s u l t s w i t h 1/3- and 2/3-span blowing f o r t h i s wing a r e given i n f i g u r e s 6 and 7 , r e s p e c t i v e l y . Except a t CLl = 0 , t h e t h e o r e t i c a l p r e d i c t i o n s g e n e r a l l y a g t e c d c l o s e l y w i t h experiment through t h e range f o r t h e t h r e e Cp blowing spans.

The poor agreement oi: was caust-d iby t h e f a c t thrlt t h e r e was a good CL, = 0 d e a l of s c p ~ r a t e d flow on t h e wing witllout b1c)wiu~. The Iesding-cdgc sl:lt i n t e r - f e r e d wich t h e flow a t low a n ~ l e s of a t t a c k , nud t h e fltrw was s e p a r a t e d around t h e t r a i l i n g edge st a l l a n g l e s o f a t t a c k becnusc t h e a i r f o i l w r r s r e l a t i v e l y t h i c k ( t / c = 0.18). J u s t n rmxierotc amount of blowing cleaned up t h e stbparat,eu flow s o t h a t t h e t h e o r y was brought i n t o c l o s e :lgreement wit!^ t h e experimental d a t n .

The t h e o r y s l i g l i t l y u n d e r p r e d i c t e d t h e l i f t a t Cil = 3.9, but t h i s ctisagrec- ment between t h e o r y and tl\i!\erCtnent was not t v p i c a l : d s t n f o r o t h e r conf i p u r n t i o n s showed good agreement a t high The f a c t t h a t t h e agreement was a s good :IS i t C .

W R Y v a l i d a t e s t h e t h i c k n e s s c o r r e c t i o n ttlc~t was a p p l i e d t o t h e l i f t c o e f f i c i e n t s e s t i m a t e d by t h e thin-wing theory. Without t h l s c o r r e c t i o n , t h e p r e d i c t e d l i i t coef f i c t e n t s would be n~ i i c e a b l y l owcr t h a n e s p r r inicnt tllroilgllaut tl\c C), range.

The d r ~ ~ g c o c f f i c i e r r t s shuwn a r e based on t h e n e t f o r c e i n t l i r drng d i r e c t i o n measured by t h e model balance and, t l l e r e f o r e , incliicic. t h e mode? t h r u s t . Thc prn- f i l e d r a g c o r f f i c i e n t o f t l ~ e model ( t h e vnliie where t h e C,, = 0 curve i n t c r s r c t s i t h e CL = 0 axis) was about 0 . 0 7 , which was very s m : ~ l l compared w i t h t h e o v e r a l l

Most of the drag developed was induced drag, level of drag being measured.

and the results indicate the theory was able to predict the induced drag very accurately. The fact that the theory assumed 100-percent suctio.1 and gave good drag prediction indicates that the wing was experiencing full thrust recovery.

Results for the aspect-ratio-5.5 wing are not presented, but there was excel- lent agreement between theory and experiment for this wing also.

Comparison of force components.- The fact that the theory gave good pre- dicticfis of total force and moment coefficients at the high jet deflection, even though it has been linearized and assumes small jet-deflection angles, is consistent with some of the previous results for the jet flap, which also show (See, for example, good agreement with test data at high jet deflections.

Spenca's two-dimensional comparisons in reference 1 and also comparisons for an augmentor wing in reference 5 . ) This good agreement can be explained by exam- It can be shown that while the ining the components of the forces and moments.

predicted total forces and moments agreed with experiment, the components of the forces and moments did not agree. The following table compares the theoretical and experimental components of lift coefficient for the 213-span blowing case at a Cp of 3 . 9 (Nu = 3.6) m d a = 0 ' . The components shown (for no flap deflection and a = 0 ° ) are the jet-reaction lift coefficient CL,jr, which is the lift coefficient due to the thrust acting at the trailing edge, and the circulation lift coefficient C L ~ , which is the lift coefficient obtained by integrating the pressure distributions on the wing.

I Planar theory 1 2 . 3 2 1 3 . 8 8 1 6 . 2 0

I

Experiment 1 3.15 1 3.17 1 6.32 1

1 . ----- - -1

Whereas the total lift coefficients agree within about 2 percent, the small-angle theory overestimates the jet-reaction lift co~ifictent: the small-

angle value for CL,~: is r ~ C ~ - ~ b (6, in radians , whereas the true jet-reaction

)

lift coefficient is The planar theory, on the other hand, will nC,, sin b j .

underestimate the circulation iif t , and hence, the pressure distribution on the

wing.

Comparison of pressure distributions.- The pressure distributiniis in fig- ure 8 show the extent to which the iinear theory underestia?tes the pressures on the wing, and hence, the circulation lift developed. Pressure distributions in of 3 . 9 .

figure 8 are for the pure-jet-f lap wing with 213-span blowing at a CU The plots shown give the net pressure dif fercnce Ac Sctwecn the upper and P lower surfaces as n function of nondtmensiunal chordwise distance nLong the wing. (In the EVD calculations Acp = 2y, where y is the vorticity at a The theoretical pressure distribut-ions s\~own have been corrected given point .)

for wing thickness effects. The pressure distributions arc given for three spanwise stations; the two inboard stations have blowing, the one outboard

station has no blowing.

It is seen that the theory underpredicts the pressures at the two inboard stations for which there is blowing ( b % = 0.15 and 0.45) but Is in good agreement with experiment: at the olltboard station where there is no blowing.

The lower wing pressures predicted by the theory can be attributed to the high jet deflections involved and to the fact that the theory satisfies bound- ary conditions in the plane of the wing rather than on the jet wake (see fig. 3 ) .

The effect of these planar assumptions is demonstrated by the results of a two-dimensional study in figure 9. Figure 9 shows pressure distributions calculated with a program developed by Clever (ref. 6) for a flat-plate two- dimensicnal wing with C , , = 3.5 and a = 0 ' . The nonplanar theory was devel- oped without making linearizing assumptions; the planar theory assumes small angles and that vorticity lies in the plane of the wing.

The comparisons show that at 6j = lo0, there is no difference between the nonplar.ar and planar theo- ries; at 6j = 30°, which is about the limit of the smali-angle assumptions, small diffe,rences start to appear. At 61 = 60°, the linearized planar theory gives lower pressures than the nonplanar theory, particularly close to the wing trailing edge.

At both 6j = 30°, the total as well as the components of 6j = lo0 and lift coefficient were about the same for the nonplanar and planar theories. The table in figure 9 compares these lift coefficients at b j = 60° and shows that, as was the case for the three-dimensional planar theory and experiment, total lift coefficient was about the same for the planar and nonplanar the.-ies, even though the components did not agree. The planar theory overestimated the jet- reaction lift, but compensated for it by underestimating the circulation lift by about the same amount.

Wing With Blowing Over the Flap Comparison of total lift, drag, and pitching moment.- Lift, drag, and pitching moments for the wing with blowing Gver the trailing-edge flap are shown in figure 10. The results are for the aspect-ratio-8 wing, full-span

blowing, and af = 30". They are typical of those for all blown-flap configu-

rations in that they show the theory consistently underestimated the lift and pitching moments for this wing throughout the C , , range. The predicted lift- drag curves were in good agreement with experiment, but the drag at a given angle of attack was substantially lower for the theory than for experiment.

These results indicate there is a substantial interaction effect due to the jet exhaust passing over the flap that was not accounted for in the theory, which assumes the jet exhaust emerges from the trailing edge of the wing. This is substantiated by the pressure distributions for the blown-flap wing as described in the following section.

Comparison o f p r e s s u r e d i s t r i b u t i o n s . - Figure 11 snows t h e p r e s s u r e d i s - These t r i b u t i o n s f o r <he blown-flap wing f o r two f l a p d e f l e c t i o n s , 30° and O".

d i s t r i b u t i o n s , given a t t h e 15-percent-semispan s t a t i o n , show b o t h upper and These were o b t a i n e d by s o l v i n g t h e t h i c k n e s s problem lower s u r f a c e p r e s s u r e s .

f o r upper and lower s u r f a c e v e l o c i t i e s w i t h no blowing, t h e n adding t h e s e t o t h e v e l o c i t i e s determined f o r t h e t h i n wing w i t h blowing from t h e EVD program.

The v e l o c i t i e s were t h e n converted i n t o p r e s s u r e c o e f f i c i e n t s by u s i n g t h e I incompressible Bernouli equation.

The p r e s s u r e d i s t r i b u t i o n f o r 6f = 30° i n d i c a t e s a high n e g a t i v e p r e s s u r e peak around t h e f l a p hinge l i n e and j e t s l o t l o c a t i o n ( a t 0 . 7 5 ~ ) ; t h e t h e o r e t i - c a l p r e s s u r e d i s t r i b u t i o n a l s o h a s a n e g a t i v e p r e s s u r e peak a t t h i s l o c a t i o n because t h e f l a p is d e f l e c t e d . The p r e s s u r e s given by t h e t h e o r y around t h e hinge l i n e , however, seem t o be much lower than t h e experimental v a l u e s , i n d i - c a t i n g t h a t t h e high-velocity jet emerging from t h e s l o t a l s o has a s u b s t a n t i a l e f f e c t . The e f f e c t of t h e jet is even more apparent a t !if = oO. The e x p e r i - mental d a t a f o r 6f = 0' a g a i n i n d i c a t e s a very high n e g a t i v e p r e s s u r e around t h e j e t s l o t and hinge l o c a t i o n ; however, t h e t h e c l r e t i c a l d i s t r i b u t i o n i n d i c a t e s no such peak because t h e f l a p i s u n d e f l e c t e d . The t h e o r y , i n t h i s c a s e , t r e a t s t h e wing as though it were a pure j e t f l a p w i t h a j e t d e f l e c t i o n e q u a l tr t h e a n g l e of t h e upper s u r f a c e of t h e f l o p .

t h e j e t appears a s a s i n g u l a r i t y t h a t is s i m i - The f a c t t h a t t h e e f f e c t of l a r t o l o g a r i t h m i c s i n g u l a r i t y caused by f l a p d e f l e c t i o n i n d i c a t e s t h a t it might be p o s s i b l e t o modify t h i s s i n g u l a r i t y t o account f o r Cv e f f e c t s as w e l l a s f l a p e f f e c t s . I f t h i s cannot be done, then a more g e n e r a l wing-jet i n t e r a c t i o n program ( s i m i l a r t o r e f . 7 ) would be needed t o account f o r t h e j e t flow over t h e f l a p .

CONCLUSIONS Comparisons made between t h e o r y and ex;>riment f o r a s t r a i g h t , untapered wing w i t h two t y p e s of powered l i f t (a pure j e t f'op and blowing over t h e f l a p ) i n d i c a t e d t h e f c l l o w i n g conclusions: 1. The l i f t , drag, and pitching-moment c o e f f i c i e n t s p r e d i c t e d by t h e l i n - e a r i z e d p l a n a r theory were i n e x c e l l e n t agreement w i t h t h e experimental v a l u e s f o r t h e p u r e jet f l a p , even though t , ~ e j e t d e f l e c t i o n was l a r g e (61° t o 63').

2. The p l a n a r theory underpredicted t h e p r e s s u r e c o e f f i c i e n t s and hence t h e wing c i r c u l a t i o n lift. The lower c i r c u l a t i o n l i f t was compensated f o r by a h i g h e r j e t - r e a c t i a a l i f t , under t h e small-angle assumptions, s o t h a t t o t a l li,t and p i t c h i n g moments were c l o s e t o t h e c o r r e c t v a l u e s .

3. The l i f t , d r a g , and pit.ctring-moment c o e f f i c i e n t s a s w e l l a s p r e s s u r c c o e f f i c i e n t s a t a given a n g l e a f a t t a c k were underpredicted f o r t h e wing w i t h blowing over t h e f l a p because of t h e f a i l u r e of t h e t h e o r y t o account f o r t h e i n t e r a c t i o n e f f e c t of t h e h i g h - v e l o c i t y J e t p a s s i n g over t h e f l a p .

REFERENCES 1. Spence, D . A . : The Lift Coefficient of a Thin, Jet-Flapped Wing. Proc. Roy.

Soc. (London), Ber. A, vol. 238, no. 1212, Dec. 4, 1956, pp. 46-68- 2 . Margason, Richard J.; Yip, Long P.; and Gainer, Thomas G . : Recent Develop- ments in Propulsive-Lift Aerodynamic Theory.

Aerodynamic Analyses Requir- ing Advanced Computers, Part 11, NASA SP-347, 1975, pp. 871-895.

3 . Lopez, Michael L. ; Shen, Cheng-Chung; and Wasson, Norman F. : A Theoretical Method for Calculating the Aerodynamic Characteristics of Arbitrary Jet- Flapped Wings. Rep. No. M C J 55519 (Contract N00014-71-C-0250), McDonnell Douglas Corp., May 1973.

Volume I - The Elementary Vortex Distribution J2t-Wing Liftinp Surface

Theory.

Volume I1 - EVD Jet-Wing Computer Program User's Manual.

4. Clever, W . C . : A Vortex Lattice Program for Jet Flapped Airfoils.

TFD-73-70, Los Angeles Div., North American Rockwell Corp., Jan. 26, 1973.

5. Lopez, M. L.; and Shen, C . C . : Recent Developments in Jet Flap Theory and Its Application to STOL Aerodynamic Analysis. AIAA Paper No. 71-578, June 1971.

6. Clever, W.: Linear and Nonlinear Two Dimensional Jet Flap Analysis.

TFD-72-698, Los Angeles Div., North American Rockwell Corp., Nov. 22, 1972.

7. Lan, C . Edward; and Campbell, James F.: Theoretical Aerodynamics of Upper- Surface-Blowing Jet-Wing Interaction. NASA TN D-7936, 1975.

Figure I.- Test node1 in the V / S T O L tunnel.

PURE JET FLAP NACA 0018-64 A l RFO l L BLOWING OVER FLAP Figure 2.- Model d e t a i l s .

WING JET SHEET

~ e c 4 4 4 e c

B A S l C v VORTEX MODEL \\ \\ SQUARE-ROOT S INGULARITY

/ r EVD ELEMENTS

S I N G U L A R I T I E S EVD I N 3 D I M E N S I O N S CONTINLiOlJS EVD D l STR I B U T I O N Figure 3.- Jet-flap lifting-surface theory, BLOWN S P A N 0 FULL U TCYO-THIRD A 6 FlAGGED SYMBOLS: PURE JET FLAP

-

0 20 40 60 80 0 20 40 60 80 FLAP DEFLECI'ION, deg F V I P DEFLECTION, deg Figure 4.- Static turning characteristics.

EVD THEORY Figure 5 . - Theory and e x p e r i m e n t comparison for purr J c t f l a p ; A - 8; f u l l - s p a n b l o v i ~ g ; b j = 6 1 . 4 ' .

C U I? 0.1 ; c ' 0.4 3 0.3 b 1.4

n 1.9

-

EVD THEORY Figure 6 . - Theory and experiment comparisl)n f o r pure j e t f l a p ; A = 8; 113-spsn blowing; 6, = 6 3 . 4 ' .

C , 0 0.1 0 0 . 5 b 1.9 15 3.9

-

EVD THEORY F i g u r e 7 . - T h e o r y a n d expt3riment c o m p a r i s o n f a r p u r e J e t i1;lp; A = 8; ? / : - s p a n b l o w i n g ; 5j = 61.W.

0 I I I I 1 EX PER I MENT EV3 THEORY /

- - _ - - -

F i g u r e 8.- T h e o r e t i c a l a n d e x p e r i m e n t a l prt1esurs3 d i s t r i b n ~ i o n s f o r p u r e j e t f l a p . 313-span blowing; CL, = 3.3.

18 '< & pOO$u

- NONPLANAR THEORY

----- PLANAR THEORY

L l f l COEFFICIENTS FOR 6. = 60° . % , -*., i .."

* E ! .: . - ' -, A c ~ J I .

xic Figure 9.- E f f e c t of planar assumptions i n t w o dimensions ( r e f . 6 ) . C,, = 3 . 5 ; a = OO.

i

, - I : i 0 0.46

, . - -

0 0.92 I D 1.85

o 3.72

- EVD THEU?Y

-

I J A I I 1 0 4 8 12 16 20 -2 -1 0 1 2 a, de9 Figure 10.- Theory and exper:.ment comparison f o r wing w i ~ h A = 8 . 0 ; full-span blowing; bf = 3 0 ~ .

blowing o l e r f l a p

+ EXPERIMENT

---

EVD THEORY

-16 -

Q 3 -1 2 -12 -

- a , = NO

-8

\ -8

C P C P -4 I .

b i 4~ .2 .4 . 6 .8 1.0 ! i Ffgure 11.- Pressure distributions for wing with blowing over flap.

15-percent semisyan statio-.; CU = 3 . 9 .

EXTERNALLY BLOWN FLAP IMPINGENENT PARAMETER Danny R. Hoad Langley Directorate, U.S. Army A i r Mobility R&D Laboratory

-q SUMMARY

This paper presents a comparison of t h e performance of two e x t e r n a l l y

. : I

blown f l a p (EBF) wind-tunnel models with an engine-exhaust f l a p impinge- ment c o r r e l a t i o n parameter. One model was a four-engine EBF t r i p l e - s l o t t e d f l a p transport. I s o l a t e d engine wake surveys were conducted t o d e f i n e t h e wake properties of give s e p a r a t e engine configurations f o r which performance d a t a were available. The other model was a two-engine EBF trans- port f o r which t h e engine wake p r o p e r t i e s were estimated. The c c r r e l a t i o n parameter was a function of engine-exhaust dynamic pressure a t t h e f l a p location, a r e a of engine-exhaust f l a p impingement, t o t a l exhaust a r e a a t t h e f l a p location, and engice t h r u s t . The d i s t r i b u t i o n of dynamic pressure f o r t h e f i r s t model was measured ; however, t h e d i s t r i b u t i o n f o r t h e second model was assumed t o be uniform.

; : . J . , -, w . -.

Numerous concepts have been developed f o r achieving short-take-off-and- --: .,:.:i landing (STOL) performance. One approach which was s e l e c t e d f o r an ~ d v a n c e d . i medium STOL transport (AMST) configuration, t h e YC--15, i s t h e e x t e r n a l l y > ' * .- I blown f l a p (EBF). Most EBF concept development h~.s been achieved with experimental i n v e s t i g a t i o n s ( r e f s . 1 t o 8) of various engine and airframe conf iguratians. While very limited a c d y s e s ( r e f s . 9 t o 11) of these config- u r a t i o n s have been attempted, some work has been done with an empirical a n a l y s i s using a c o r r e l a t i o n parameter (impingement parameter) based on t h e v e r t i c a l d i s t a n c e t h a t t h e f l s ~ t r a i l i n g edge extends i n t o t h e j e t exhaust from t h e engine center l i n e and t h e r a d i u s of t h e j e t exhaust a t t h e f l a p t r a i l i n g edge. (See r e f . 12.)

The present paper describes t h e r e s u l t s of a r e l a t i v e l y simple a n a l y s i s based on a n engine-exhaust f l a p impingement p e r m e t e r , which is a function of t h e engine-exhaust dynamic pressure a t t h e f l a p l o c a t i c n , t h e a r e a of engine-exhaust f l a p impingement, t h e t o t a l exex'~aust area a t t h e f l a p l o c a t i o n , and t h e t h r u s t . Isole.ted engine wake surveys were conducted t o d e f i n e t h i s parameter f o r one of t h e EBF models f o r which aerodynamic performance d e t a were a v a i l a b l e ( r e f . 2). A unif orm dynamic pr eaerae p r o f i l e was assumed t o deter- mine t h i s parameter f o r t h e other EBF model. (See r e f . 13.)

SYMBOLS t o t a l a r e a of engine exhaust which impinges on f l a p , m

-

incremental a r e a of engine exhaust which impinges on f l a p , m t o t a l area 3f edgine exhaust a t f l a p l o c a t i o n , m L i f t

l i f t c o e f f i c i e n t , -

qs thrust-removed l i f t c o e f f i c i e n t Thrust t h r u s t c o e f f i c i e n t ,

q s

l o c a l wlng chord, m engine-exhaus t f l a p impingement free-stream dynamic pressure, N/m incremental dynamic pressure of engine exhaust which impinges on f l a p , ~ / m ~ engine-exhaust dynemic pressure which impinges on f l a p , ~ / m wing a r e a , m s t a t i c t h r u s t , N v e r t i c a l distance, m angle of a t t a c k , deg nominal f l a p d e f l e c t i o n angle, deg engine-exhaust d e f l e c t i o n (measured from body a x i s ) , -1 Normal f o r c e tan s deg Axial f o r c e

J ( N ~ - I force)' + (Axial f o r c e )

~ t a t i c - t h r u s t recovery e f f i c i e n c y , T

Abbr evia t iona : BPR bypass r a t i o EBF e x t e r n a l l y blown f l a p MODELS Two wind-tunnel i n v e s t i g a t i o n s were conducted t o determine t h e e f f e c t of d i f f e r e n t engine-exhaust c h a r a c t e r i s t i c s on t h e performance of two s e p a r a t e EBF t r a n s p o r t con£ i g u r a t i o n s ( f i g s . 1 and 2).

A four-engine EBF trangport (model 1, f i g . 1 ) was t e s t e d i n t h e Langley VIiTOL tunnel. It had a 25 quarter-chord sweep, leading-edge s l a t s d e i l e c t e 50 . a$ t r i p l e - s l o t t e d full-span f l a p s whose elements were d ~ f l e c t e d 0 . 20 , and 40 , respectively, f o r the take-off configuration, and 15 , 35 , and 55O, respectively, f o r the landing configuration. The engines were simulated by a two-part e j e c t o r s i m i l a r t o that i n f i g u r e 3. Each engine simulator was f i t t e d with f i v e separate cowl assemblies intended t o represent f i v e d i f f e r e n t engine configurations ( f i g . 4) : (1) TF34 engine (BPR 6.2). (2) TF34 with n o i s e suppressor nozzle (daisy nozzle), (3) ctretched version of the TF34 (4) JT15D engine (BPR 3 . 2 ) , and (5) hi-gh-bypass-ratio (modified BPR 6.2).

engine (BPR 10). The modified BPR 6.2 engine was b u i l t so t h a t t h e engine e x i t would be a t t h e same chordwise l o c a t i o n a s t h e f a n e x i t on the d a i s y nozzle. The B P R of these engine simulators does not describe i n any way t h e s i z e , horizontal p o s i t i o n , o r v e r t i c a l p o s i t i o n of t h e simulator, but is only intended t o be a means of nomenclature. The important a s p e c t s of t h e simula- t o r a r e not the c h a r a c t e r i s t i c s a t t h e e x i t , but t h e wake c h a r a c t e r i s t i c s a t Since t h e t h e l o c a t i o n of t h e f l a p a s is evident subsequently i n t h i s paper.

exhaust and wake c h a r a c t e r i s t i c s of t h e s e v e r a l f u l l - s c a l e engines represented ere unknown, it is not possible t o r e l a t e t h e present r e s u l t s t o t h e per- formance of t h e f u l l - s c a l e engines. For f u r t h e r d e t a i l s of t h i s model s e e reference 2.

A two-engine straight-wing EBF t r a n s p o r t (model 2, f i g . 2) was t e s t e d i n It had t h e 5.18- t e s t s e c t i o n of t h e L8ngley 330-MPH 7- by 10-foot tunnel.

a leading-edge s l a t deflected 40 , n double-slotted f l a p d e f l e ~ t e d ~ 4 0 ~ f o r t h e take-off configuration, and a t r i p l e - s l o t t e d f l a p deflected 60 f o r t h e landing configuration. The engines were simulated by a two-part e j e c t o r a s presented i n f i g u r e 5. The engine v e r t i c a l p o s i t i o n on t h i s configuration was varied ( t o t h r e e positions) t o determine its e f f e c t on t h e performance of t h e con£ i g u r s t i o n ( f i g . 6).

TEST Both models were mounted on a sting-supported sLc-component strain-gage balance f o r measurements of t h e t o t a l f o r c e s and moments. I s o l a t e d engine wake surveys w r e conducted f o r each engine configuration on model 1 so t h a t t h e engine-exhaust f l a p impingement parameter could be determined. Dynamic pressure measurements were made with a preesure rake positioned s o that t h e probes were alined along a r a d i a l l i n e from t h e geometric c e n t e r l i n e . Four r a d i a l p o s i t i o n s were chosen f o r t h e d a i s y nozzle and tw r a d i a l p o s i t i o n s were chosen f o r t h e o t h e r four engine airnulators. These measurements were repeated at various downstream l o c a t i o n s t o o b t a i n dynamic pressure p r o f i l e c h a r a c t e r i s t i c s . I s o l a t e d engine wake surveys were not a v a i l a b l e f o r t h e engine simulators on model 2. Since t h e game engine was used i n a l l t h r e e positions, it was f e l t t h a t assuming a 1 0 spread angle would be s u f f i c i z n t t o d e t e r n i n e t h e r e l a t i v e influence of t h e exhaust f l a p Impingement parameter.

Jet d e f l e c t i o n angles b j and s t a t i c - t h r u s t recovery e f f i c i e n c y q f o r both models were determined from measurements of t h e nonnel and a x i a l f o r c e s made i n t h e s t a t i c - t h r u s t condition with f l a p s deflected and leading-edge slat deployed.

CALCULATIONS Isolated engine wake surveys were a v a i l a b l e f o r each engine configuration used on model 1. The f l a p impingement parameter was computed using the d i s - t r i b u t i o n of dynamic pressure a t t h e f l a p l o c a t i o n i n t h e foll.owing manner: A schematic of exhaust impingement on the f l a p is presented i n f i g u r e 7 .

The term C q caa be seen a s a summation of a l l t h e dynamic pressure i f # i %,i measurements multiplied by t h e associated f l a p a r e a on which they impinge In equation ( I ) , is the t o t a l a r e a of engine exhaust at thz f l a p impinge- ment plane and I)fAjis the nouilnal f l a p d e f l e c t i o n angle. Since the t e n is a t h r u s t o r f o r c e term, t h e parameter was divided by t h r u s t T

f %,I 4,i

and nondimensionalized by an a r b i t r a r y constant S (wing area!.

Since i s o l s t e d engine wake surveys were not a v a i l a b l e f o r t h e engine configuration used on model 2 , the dynamic pressure was assumed t o be uniform a t the f l a p 1oca:ion.

The exhaust was aesumed t o spread a t an angle of 10' t o determine the a r e a of t h e exhaust a t t h e f l a p location.

The impingement

-

parameter i n t h i s case is s l i g h t l y simplified i n t h a t t h e dynamic pressure is assumed constant over t h e f l a p ; t h a t is, S PA = qj Af sin 6f (2) j The l i f t developed by a powered-lift system can be separated i n t o t h r e e p a r t s according t o source: (1) t h e l i f t t h a t would have been produced by t h e unpowered wing, (2) t h e l i f t due t o t h e component of t h e jet which has been (3) t h e l i f t due t o c i r c u l a t i o n i n d u ~ e d redirected by t h e f l a p system, and by t h e blowing. I f the portion of t h e l i f t due t o t h e j e t is removed from t h e t o t a l l i f t , a thrust-removed l i f t c o e f f i c i e n t given by

-

= CL - s i n (6 + a)

'L, tr

v j

rc remains which can be r e l a t e d t o t h e f l a p impingement parameter.

RESULTS AND DISCUSSION AS a t e The f l a p s t a t i c turning e f f e c t i v e n e s s parameters f o r both mode1 These parameters f o r t h e f i v e presented i n f i g u r e 8 i n polar coordinate form.

engine simulators on model 1 and t h e t h r e e p o s i t i o n s of t h e engine simulator on model 2 i n t h e take-off and landing configurations a r e presented a t a par- t i c u l a r l e v e l of t h r u s t . This was the t h r u s t l e v e l used t o o b t a i n a t h r u s t c o e f f i c i e n t of 2 i n the wind-tunnel t e s t f o r each configuration.

The perpendicular d i s t a n c e from a d a t a point i n f i g u r e 8 t o t h e horizontal axisowould represent t h e l i f t component due t o t h r u s t a t a n angle of a t t a c k of 0 .

If it were assumed t h a t with zero power a l l engine simulator configu- r a t i o n s had i d e n t i c a l c h a r a c t e r i s t i c s and that t h e only a d d i t i o n s t 9 t h e aerodynamic c h a r a c t e r i s t i c s a t an angle of a t t a c k of O0 were those components i n f i g u r e 8, an assessment could be made ss t o t h e r e l a t i v e merit of t>e configurations. A s discussed i n reference 2, t h i s assesument of t h e aero- dynamic c h a r a c t e r i s t i c s of an engine model configuration is not n e c e s s a r i l y true. A more pertinent comparison based on t h e engine-exhaust f l a p impinge- ment parameter is presented i n t h i s paper.

' i .

The f l a p impingement parameter is a c o r r e l a t i o n parameter which r e l a t e s I , I A: 1 ; the engine-exhaust p r o p e r t i e s t o t h e performance of t h e engine .sodel configu- I . # . .

r a t i o n . The l i f t c o e f f i c i e n t performance and impingement parameter a r e I . : ! . , presented f o r several model f l a p combinations a t an angle of a t t a c k of 0 ' a s I .

f 0 llow3 : I . '.

I . .*

Figure Model Flap configuration 9 1 Landing 1 0 1 Take-of f 11 2 Landing 1 2 2 Take-o f f The e f f e c t of engine configuration f o r landing f l a p d e f l e c t i o n on model 1 (fig. 9) i n d i c a t e s t h a t t h e engine simulators which produce t h e l a r g e s t im- pingement parameter (BPR 6.2 and modified BPR 6.2; provide t h e l a r g e s t l i f t c o e f f i c i e n t . The simulator which produces t h e smallest impingement parameter (BPR 3.2) provides t h e smallest lift c o e f f i c i e n t . This impingement parameter is r e a l l y a measure of t h e proportion of o r momentum, which impinges on q A J 1' t h e f l a p and, i n t u r n , is d e f l e c t e d and induces c i r c u l a t i o n . This would i n d i c a t e that t h e more momentum captured by t h e f l a p system, t h e b e t t e r t h e combination w i l l perform. The e f f e c t of engine configuration f o r take-off f l a p d e f l e c t i o n on model 1 (fig. 10) is s i m i l a r except f o r t h e r e l a t j v e per- formsnce and magnitude of impingement parameter f o r t h e d a i s y nozzle. The d a i s y nozzle has e i g h t f a n lobes and nine gas generator lobes. Since t h e wake survey f o r t h i s engine included only two f a n lobes and two g a s generator lobes, it was not comprehensive enough t o adequately d e f i n e t h e p r o f i l e .

The comparisons of performance and impingement parameter f o r model 2 with landing and take-off f l a p d e f l e c t i o n s are presensed i n f i g u r e s 11 and 12, The r e s u l t s f o r t h e landing configuratiqn i n d i c a t e that moving t h e engine exhaust v e r t i c a l l y toward t h e wing lower s u r f a c e increases t h e proportion of qjAjr o r momentum, which is captured by t h e f l a p system and, i n t u r n , generates increased l i f t . Figure 1 2 f o r t h e take-off configuration i n d i c a t e s t h e same trend, although a t a somewhat lower v a l u e because t h e r e is l e s s f l a p pto- j e c t i o n t o capture t h e exhaust.

To f u r t h e r r e l a t e t h i s f l a p impingement parameter t o t h e l i f t performance of t h e models, t h e thrust-removed l i f t coef f ' l e n t a t an a n g l e of a t t a c k of 0' was computed f o r each configuration and is presented i n f i g u r e 1 3 a s a function of t h e impingement parameter, It is evident that t h e thrust-removed l i f t c o e f f i c i e n t is t h e prime aerodynamic f a c t o r which can be r e l a t e d t o t h e im- pingement parameter, because t h e d a t a f i t a s t r a i g h t l i n e which i n t e r c e p t s PA = 0 a t t h e value of ( + , which corresponds t o t h e power-off condition.

( I n t h i s case, PA should be zero.) I f t h e blowing d i d not induce super- c i r c u l a t i o n l i f t , t h e s e d a t a poi,its would be on h o r i z o n t a l l i n e s . I n each model, it can be seen t h a t t h e landing configuration d a t a d e s c r i b e a Line with a l a r g e r s l o p e than that f o r t h e take-off configuration. This is more evidence t h a t s i n c e t h e landing f l a p s capture more of t h e e x h a u ~ t flow, more c i r c u l a t i o n l i f t is induced. This again emphasizes t h e f a c t t h a t t h e engine- exhaust f l a p impingsment parameter 's a measure of t h e proportion of momentum captured by t h e f l a p and provides a method t o a s s e s s t h e r e l a t i v e performance of engine-wing combinations.

,=- This paper provides a technique t o a s s e s s t h e r e l a t i v e performance of e x t e r n a l l y blown f l a p (EBF) c o n f i g u r a t i o n s by means of an engine-exhaust f l a p impingement parameter. This parameter was determined t o be a function of t h e proportion of momentum which is captured by t h e f l a p system.

The l i f t produced by an EBF configuretion can be r e l a t e d t o t h e propor- t i o n of momentum captured by t h e f l a p system. Furthermore, It has been shown - t h a t t h e thrust-removed-lift coeff i c f e n t can be d i r e c t l y r e l a t e d t o t h i s captured momentum, defined by t h e engine-exhaust f l a y impingement parameter.

REFERENCES 1. STOL Technology. NASA SP 320, 1972.

2. Hoad, Danny R. : L o n g i t u d i n a l Aerodynamic C h a r a c t e r i s t i c s of a n E x t e r n a l l y Blown F l a p Powered-Lift Model With S e v e r a l P r o p u l s i v e System Simulators.

NASA TN D-7670, 1974.

3. P a r l e t t , L y s l e P.; Freeman, Delma C., Jr.; and Smith, C a a r l e s C., Jr.: Wind-Tunnel Investigation of a J e t T r a n s p o r t A i r p l a n e C o n f i g u r a t i o n With High T'nrust-Weight R a t i ~ and a n External-Flow Jet Flap. NASA TX D-6058, 4. Smith, C h a r l e s C., Jr. : E f f e c t o f Wing Aspect. R a t i o and F l a p Span on Aerodynamic C h a r a c t e r i s t i c s of an E x t e r n a l l y Blown J e t - F l a p SYUL Model.

NASA TN D-7205, 1973.

5. P a r l e t t , L y s l e P.; Smith, C h a r l e s C., Jr.; and Megrail, James L.: Wind- Tunnel I n v e s t i g a t i o n of E f f e c t s of V a r i a t i o n s i n Reynolds Number and Leading-Edge Treatment on t h e Aerodynamic C h a r a c t e r i s t i c s of a n E x t e r n a l l y Blown J e t - F l a p Configuration. N A S A TN 0-7194, 1973.

6. Aoyagi, Kiyoshi; F a l a r s k i , Michael D. ; and Koenig , David G. : Wind-Tunnel

I n v e s t i g a t i o n of a Large-Scale 250 Swept-Wing Jet T r a n s p o r t Model With a n E x t e r n a l Blowing T r i p l e - S l o t t e d Flap. NASA TM X-62197, 1973.

7. Reeman, Delma C., Jr.; P a r l e t t , L y s l e P.; and Henderson, Robert L.: Wind- Tunnel I n v e s t f g a t i o n of a Jet T r a n s p o r t A i r p l a n e C o n f i g u r a t i o n With a n External-Flow J e t F l a p and Inboard Pod-Mounted Engines. N A S A TN D-7004, 1970.

8. Vogler, Raymond D.: Wind-Tunnel I n v e s t i g a t i o n of a F o u r - E ~ g i n e E x t e r n a l l y Blowing J e t - F l a p STOL A i r p l a n e Model. N A S A TN D-7034, 1970, 9. Johnson, William G . , Jr.; an.d Kardas, Gerald E.: A Wind-Tunnel I..vestiga- t i o n of t h e Wake Near t h e T r a i l i n g Edge of a D e f l e c t e d E x t e r n a l l y Blown Flap. NASA T M X-3079, 1974.

10. Coldhammer, M. I.; Lopez, M. L.; and Shen, C. C.: Methods f o r P r e d i c t i n g t h e Aerodynamic and S t a b i l i t y and C o n t r o l C h a r a c t e r i s t i c s of STOL A i r - c r a f t . Volume I - Basic T h e o r e t i c a l Methods. AFFDL-TR-73-146-Vol. I, U.S. A i r Force, Dec. 1973.

11. May, Fred; and Widdicon, C o l i n A,: STOL High-Lift Design Study. Volume 1.

State-of-the-Art Review of STOL Aerodynamic Technology.

AFFDL-TR-71-26-Vol. I, W.S. A i r Force, Apr. 1971.

-

a a

12. Roe, M . He; Renselrer, D . J.; Quam, R. A . ; et al.: STOL Tactical Alrcraft Investipation - Externally Blown Flap. Volume I1 - Design Compendium.

AFFDL-TR-73-20, Vol. 11, U.S. Air Force, Apr. 1973.

(Available from DDC a8 AD 770 110.)

13. Johnson, William G., Jr.: Longitudinal Aerodynamic Characteristics of a Wing-Body Combination Having a Rectangular, Aspect-Ratio-6, Slotted Supercritical Wing With Externally Blown Flaps. NASA 'I'M X-2319, 1971.

I i

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-: A d B-0 C-C CI- D H w ~ @ 3s- Two-part engine 8imulhtor for model 1.

M O D I F I E D BPR 6.2 WIM T I K E - O F F FLAPS BPR 62 WITH ZANDlNG FLAPS DA I S Y NOZZLE W I T d kw' BPR 1 0 WITH I A N D I N G SLAPS

LANDING ;LAPS - A

BPI? 32 W I T H TAKE-OFF FLAPS figure 4 . - Engine confiprationa w e d on model 1 .

LPERFORATED PLATE Flmrc 5.- Two-part engine s m l a t o r for mdel 2.

d TAKE-OR FLAP 60° LANDING FLAP Figure 6.- Engfae poeitions u a d on mdel 2.

TAKE-OFF LANDING FLAP FLAP

a

+

A L 1 and 2 .

2 BFR 6.: 10

c MODIFIED BPR 6.2

+ D A I S Y !;O;'LE

BPK 10

e

.y "

i 2 4 figure 9 . - Performanre PIIA * - - 1 - -

1 11

0 BPR 6.2

o MODIFIED BPR 62

0 D A I S Y NOZZLE A BPR 10.0 b BPR 3.2 4 4 P~ 2 2 0 2 4 c,.i Figure 1 0 . - Performance and impingement Farameter for model 1 with take-off flap deflection. a = 0".

0 1 ?

C IJ Figure 11.- Performance and impingement parameter for model 2 with a = 0 ' .

landing flap deflection.

Figure 12.- Performance and impingement parameter for model 2 with take-off flap deflection.

a = 0 ' .

MODEL 1 MODEL 2 ENG l NE TAKE-OFF LAND1 NG Z/C TAKE-OFF LAND1 NG F i A P FLAP FLAP FLAP BPR 6.2 CI M O D I F I E D BPR 6.2 O B

D A I S Y NOZZLE 3 •

GPR 10 A A BPR 3.2 b L L A N D I N G FLAP7 LANDIWG FLAP TAKE-OFF FLAP PONER OFF CL Figure 13.- Thrust-removed lift coefficient as a function of impinge- ment parameter. a = 0 ' .

m 'pxoqa Suyn m 'ueds Suyn .poqaam ~ e a y a d ~ e u e delj-aaf auo oa szaaamczed anduy so pasn sa-[nsaz aq3 qaym aqna meazas asneqxa auysua aqa 30 uoyaeao~ PUB qa%uaias mnauamou aqa xoj suoyanqpasyp asynueds dolahap oa pasn uaqa aian sauamaxnseam *u~yaaxn%y3uoa ajy~-paxanod delj m o l q dl~euxaaxa ue 30 ayen zeau aqa u~ asaqL pauyeJqo axan sJuamaznseam 1aa3aa day30pfi q ~ y q n uy (T *;ax) uoyae8yasaauy Tauuna -puyn e ~ 0 x 3 s31nsax 30 srsdleue aqa 30 dxennuns e s s n ~ s y p ~ l y n zaded s y q ~ *szaaamezed anduy se pasn axe qayqn uoyaeao~ pue qa%uaxas mnauatnom ayan aqa 30 suoyadynsap aqa uo d ~ a u e a y ~ y u S y s puadap s p o q ~ a m asaqa ae1!3 m o q s seq aauayxadxa .sayen a3y~-pazanod dq3-aaf yDyqa j c . zoraeqaq a43 a q p a z d 02 pasyaap uaaq aaeq spoqaatn ~ e a y a d ~ e u e snoyzea 'ajex3xre asaql I I 30 u%ysap aqa uy asysso OJ, * a j e x ~ x y e S u y p u e ~ pue 330-aye3 p a ~ n p a x xo azoqs 103 se Tlan se a a u m x o 3 ~ a d 8uypuel pue 330-aye3 xyaqa a ~ u e q u a 02 zapxo uy sazodsuexa lneq-%uo~ ~euoyauaauoa oa uoy3u3y~dde 303 pauymexa Suyaq nou sy agyl palanod 'salnsax leayad~eue pue ojep p ~ u a m y x a d x a aqa uaanlaq auamaaz8e alqe -uoseaz noqs sllnsaz aqa 'poqaam delj-laF ~e3yabTeue auo oa sxaaamexed anduy se pasn aze uoyaaazyp pue y?%uszas mnauamom 30 suoyanqyxasyp daazns ayen uaqn 'os~v .aspa 8uylyeza d e ~ j aqa ae qa%ua~as mnauamom aqa 8uya3ypaxd 30 sueam e so x o a ~ e 3 dxaaoaax asnxqa de13 aqa 30 asn aqa sarjyxah q ~ y q m 'uoyaepxzoa poo8 e uy salnsax eaep a3103 araaas moz3 sanlen paaarpazd aql q ~ y n s d a u n s aqen paads p z ~ m r o 3 m o q pauyeaqo s a u a y 3 ~ ~ ~ a o a mnauamom aqa 30 uosy~edtnoa a aeqa pamoqs sysdleue syq~, .poqaam le3yadyeue de~3-aaC auo oa sxaaa -mezed anduy sa pasn salnsax aqa qayn aqna meaxas asneqxa auySua aqa 30 uoyae~ol pue q ~ % u a z a s mnauamom aqa 203 suoylnqyzasyp ~ S ~ M U E ~ S d o ~ a a a p oa pasn azam sauam -axnseam asaqL *uoyaexn%yjuoa 2 3 ~ ~ - p a x a n o d do13 umolq d ~ ~ e u x a a x a ue 30 ayen zeau aq3 uy pauyeaqo axan sauamaxnseam xoaaaa X a y ~ o l a ~ q q q ~ uy uoyaoSyasa~uy lauuna-puyn e moxj salnsaz 30 sysd~sue aqa 30 dxemmns e sassnasyp xaded syq~,

Lift

-

lift coefficient, LS Gross static thrust

thrust coefficient, -

&ps section momentum coefficient from wake surveys total momentum coefficient from wake surveys number of velocity vectors in jet thickness free-stream dynamic pressure, ?a wing area, m 2 magnitude of velocity vector, mlsec magnitbde of specific velocity vector associated uith 6j,l, mlsec free-stream velocitJ7, nlsec distance along span of wing, m angle of attack, deg flap deflection, deg Normal force jet deflection angle,

arctan - Axial force ' deg

local downwash angle minus wing angle of attack, deg mean turning angle at any given survey station, deg change in turning angle due to power, Ceg Resultant force flap thrust recovery, Static thrust jet thickness DISCUSSION The investigation was conducted in the Langley V/STOL tunnel on an exist- ing externally blown flap model (ref. 2 ) . As shown in figure 1, measurements were made at 11 spanwise $tations between the edge of the fuselage and the

" v , - , . - - ' - - - .... , . q r- ---?..- ,- -

midsemispan station on the wing. The measurements were made with a split-film total vector anemometer which was capable of sensing the three velocity vector ; components (ref. 3 ) . At each spanwise station, the probe was moved through a vertical arc through the wake and measured more than iOO velocity vectors at

:: i

- -;!

each survey station for each engine-power-off and engine-power-on condition.

d

: : 1

i Figure 2 shows a sample of the power-on velocity profiles taken at zero .-.

angle of attack and for thrust coefficients of 2.0 and 4.0. The profiles are

..;:I

lrq . - .1

shown at the two spanwise stations, -$- = 0 . 2 2 7 and Y- = 0.376, which are

, .

i .

. ' , , ' !

h, 2 b / 2 just inboard of the engine center lines. The outline of the wing-flap system is shown to locate the profiles relative to the 35O deflected flap and engine thrust center line. Units of free-stream velocity are indicated by dashed arcs within the profile. The profiles show higher than free-stream velocity flow above the wing and lower than free-stream velocity flow below the engine exhaust stream tube. The planform sketch in figure 1 shows that at the inboard station, the flap trailing edge is at a greater distance from the engine exhaust plane than at the outboard station. This greater distance results in the inboard exhaust stream tube being more diffuse than the jet sheet wake at the outboard station. Figure 2 shows that this wake characteristic hoids for both thrust levels.

i To establish the vertical bounds of the engine ex.. ..ust stream tube, a jet i

' I

thickness T was defined as the wake region in which the change in the local I velocity vectors exceeded 0.3 m/sec. The magnitudes of the velocity vectors 1 1 I .

within the jet thickness were integrated to obtain a section momentum coeffi- / , I

cient c ~ " at each spanwise station by the following equation: I I

I

, , This equation provides for the subtraction of the power-off mrlentum from the power-on momentum in coefficient form per unit area. The results of this inte- 3 as a function of the spanwise location gration are shown in figure

$jZ. The

data show that t:~c: peaks of the distributions have shifted significantly inboard of the ensine center lines. This shift is contrary to the accepted thought that the engine exhaust spreads significantly outboard on a swept-wing configurat~on.

In addition to obtaining values for the distribution of the momentum strength and location along the flap trailing edge for analytical purposes, a comparison of a total momentum coefficient due to power from the forward speed- wake surveys and the predicted momentum coefficient obtained by use of static force data would be useful in verifying the technique generally used for defin- ing the strength of engine exhaust momentum at the flap trailing edge.

Tn obtain t h z wake survey value, the section momentum coefficient distr.butions were extrap- olated to the edge of the fuselage and to the most outboard extent

of the wing felt to be impacted by the spread engine exhaust flow

(ih = 0*6)*

The distributions were then integrated spanwise to obtain the total momentum coefficient C , , due to the engine exhaust at the flap trailing edge. The predicted value from static force data is obtained from the product of the

static parameter q and the engine thrust coefficient C y , where n is

defined as the efficiency of turning the engine exhaust flow through some deflec- tion angle. When the static force data are plotted as shown in figure 4, the value of q can be obtained as the radial dista~ice from the origin to the data point. A comparisori of the integrated values of engine momentum coefficient from the forward speed wake surveys with the product of the static values of 0 and CT is shown in figure 5, and the result is a very good agreement for all of the thrust levels and angles of attack tested.

' i A mean turning angle was determined for each spanwise survey station . I 6j* i by the following equation: , 1 : I * where tij is the local downwash angle minus the wing angle of attack, Vj is the magnitude of the corresponding streamwise velocity vector, and n is the number of measured velocit) vectors between the chosen limits. This equation weights 6 , by the velocity magnitude since it was felt that the higher velocity vectors had a stronger influence on the overall turning angle of the jet. The procedure was applied to the power-off and power-on data, with the resulting spanwise distribctions shown in figure 6 by the circles and squares, respectively. These distributions show a change in the peak locations from

those in the momentum distributions - the peaks are now shown on the engine

8 1 .

center lines. This characteristic indicates that for this configuration

9 1

(6f = 3 5 0 ) , most of the engine exhaust is passing under the flap trailing edge a . * .

with little turning. Since most jet-flap theories use the description of the

'1

jet-wake effects as an addition to basic wing theory, the l o c d flow angle due 1 ..I 1 ~ to power was obtained by subtracting the power-off distributions from the power- ., ' a on distributions. The results are shown in figure 7 as a function of spanwise - I _ . .

, .

station. 1 .

The distributions which were obtained for the momentum coefficients (fig. 3 ) and jet-deflection angles (fig. 7 ) were then used as input parameters to a jet- flap analytical method developed by Lissaman (ref. 4 ) . The results of the com- puLation are compared in figure 8 with the experimental lift coefficients as a function of wing angle of attack. The comparison indicates analytical results similar to the corresponding experimental data.

i i CONCLUSIONS

I

I The analysis of 'esults from a survey of the near wake of an externally I blown fiap configuration has resulted in the following conclusions:

.i

1. A comparison of the momentum coefficients obtnlnad from forward speed + wake surveys with the predicted values from static force data results in a good

1 . 1 I

! correlation, which veiifies the use of the flap thrust recovery factor as a Reens of predicting the momentum strength at the flap trailing edge.

I , i

- I

2. When wake survey distributions of momentum strength and direction are

. I

used as input parameters to one analytical jet-flap method, the results show

I

I ' rc?asonable agreement between the experimental data and analytical results.

9 /

1. Johnson, William G., Jr.; and Knrdas, Gerald E.: A Wind-Tunnel Investigation of the Wake Near the Trailing Edge of a qeflccted Externally Blown Flap.

NASA TM X-3079, 1974.

2 . J,:,,,son, William G., Jr. : Aerodynamic Chai acteristics of a Pow;?red, Exter- nally Blown Flap STOL Transport Model Wi: h Twa Engine Simulator Sizes.

NASA TN D-8057, 1975.

3. Olin, J. C.; and Kiland, R. S.: Split--Film Anemometer Sensors for Three- Aircraft Wake Turbulence and Directional Velocity-Vector Measurements.

Its Detection, John H. Olsen, Arnold Goldburg, and Milton Rogers, eds., Plenum Press, Inc., 1971, pp. 57-79.

LL Figure 1 . - Spanwise survey s t a t i o n s .

rigure 2.- Streamwise v e l o c i t y p r o f i l e s . a = OO; Vm = 2 5 . 4 5 m/sec.

. 7

- INBOARD

ENGINE

6 - ENGINE

I 3 - 1 i 2 - 1 I I I Figure 3.- Experimental momentum coefficient distribution.

(1 = O ; CT 3 . 0 .

6. = 90 THRUST I E \ ' E L S 90Rh~AL~ORC~ S T A T I C THRUST .4

NN !?I_U_D_1NAL FPKE

S T A T I C THRUST Figure 4.- Static tr~rning <ti;gles and thrust recovery factor.

Figure 5. - Momentum coefficient from forward

speed and static data.

I - - - - INSCARD ENGINE ENG! NE I 0 POWER OFF 0 POWER ON I I Figure 6.- Experimental distribution of jet-deflection angle. c t = 0 ' ; CT 7.0.

-

INBOARD - ~ 6 : g - J

- I I 1

I I 1

I I

-

I

I I 1 1

I ! ! I 11 I I I I -4 0 . 2 .4 .6 .8 1.0

&

Figure 7 . - Experimental jet-deflection angle due to power.

a = 0 ' ; cT z 2.0.

Figure 8.- ComparLson of theory with experimental data for lift coefficient.

AERODYNAMIC CHARACTERISTICS IN GROUND PROXIMITY James L. Thomas, James L. Bassell, Jr., and Luac T. Nguyen NASA Langley Research Center Results from recent investigations in the Langley V/STOL tunnel of an exter- nally blown t- n and an upper-surface blown flap configuration in ground proxim- ity are presented. Comparisons of longitudinal aerodynamic characteristics indicate that in ground proximity, drag is reduced for both configurations, but changes in lift are configuration dependent. Steady-state analyses ~f the land- ing approach indicate an increase in flight-path angle for both configuratiocs in ground proximity because of the drag reduction. Dynamic analyses with a fixed-base simulator indicate that the resultant flight path during landing approach is dependent on the initial flight-path angle and the control t2ch- nique us2d.

/

Effects of asymmetries, such as siacslip or roll and engine-out character- I istics, in ground proximity were also available £ram the wind-tunnel tests.

Sideslip characteristics were generally unaffected by ground proximity. Roll attitudes were unstable at heights near gear touchdown height, and no signifi- cant yaw-roll coupling was noted. Engine-out characteristics were unaffected

I

by ground proximity.

i I I INTRODUCTION

i

In 1969, an investigation (ref. 1) was conducted in the 17-foot test section

I

I of the Langley 300-MPH 7- by 10-foot tunnel to determine the aerodynamic char- !

acteristics in ground proximity of the four-engine externally blown flap (EBF) i configuration shown in figure 1. Various combinations of the segmented full- 1 span double-slotted flaps were tested. Typical flap deflections were 30° in a j take-off conf iguration and 60° in a landing configuration. Both high and low i ositions of the wing were tested. Changes in lift, drag, and pitching moment I in ground proximity were measured over a moving ground belt.

I The results from that investigation were used as the basis for a study pre- 2 ) of ground proximity sented at the STOL Technology Conference in 1972 (ref.

The conclusions of that study were effects on powered-lift landing performance.

that the lift loss in ground proximity for most powered-lift configurations could be correlated with the height of the flap trailing edge and the level of devel- I The lift loss increased as the trailing edge of the flap approached oped lift.

I

the ground and increased with increasing lift coefficient . The "adverse" ground

effect was, therefore, greater for low-wing configurations. Steady-state and in-flight simulator analyses icdicated that acceptable landings could be mads with conventional applications of power and elevator although the landing task was more difficult for a low-wing as opposed to a high-wing configuration.

Since aerodynamic characteristics are a function of the height above the . .

- . / ground, there is a need to assess possible adverse effects of airplane position asymmetries, such as sideslip or bank angle, in ground proximity which might be * , : i .. " critical during the landing approach. For the example of an airplane banked in - - : I '* - . 2 2 ground proximity, the lift loss might increase on the wing closer to the ground

. - I

and be reduced on the higher wing, thereby causing a rolling moment into the i *I ground beyond the available control power. This consideration led to tests in

i ' i

the Langley V/STOL tunnel of the EBF model shown in figure 2. The model is a i ' 1 four-engi-ne configuration with full-span triple-slotted flaps very similar in I . . -; planform to the earlier EBF model tested. Flap trailing-edge deflection angles

i - 1

Forces were 400 in a take-off configuration and 55O in a landing configuration.

.,I

and moments were measured over a moving ground belt with a boundary-layer removal system in the front of the test section over a range of test conditions. The tests allowed an assessment of the effect of airplane position asymmetries, including roll angle, sideslip angle, and combined roll and sideslip angles, in

\ .i

ground proximity as well as a comparison of longi-udinal characteristics in { / I ground proximity with those for the earlier EBF configuration.

. .I

The upper-surface blown (USB) concept is a rather different type of powered- I.

lift concept for which little data in ground proximity are available and which might have unexpected changes in aerodynamic characteristics near the ground,

. ;I

This consideration led to tests in particularly with one engine inoperative.

_ i the Langley V/STOL tunnel over a moving ground belt of the USB model shown in figure 3 . The model is the twin-engine configuration discussed by Phelps,

: 1

Johnson, and Margason in reference 3. Trailing-edge deflection angles of the

; 1

Coacda flap behind the engines were 20° in a take-off configuration and 60° in a landing configuration. Outboard of the Co~nda flaps were double-slotted flaps

! .I

i - and a blown drooped aileron. The wind-tunnel results allowed an assessment of

:j

the longitudinal and engine-out characteristics of the USB configuration.

I This paper thus updates the previous study on powered-lift aerodynamics in ground proximity with recent research results in the V/STOL tunnel. Comparisons of longitudinal aerodynamic characteristics for the EBF and for the USB configu- rations in ground proximity are possible. Steady-state and dynamic analyses of the landing approach for a typical STOL airplane are made to indicate the con- sequences of the aerodynamic changes in ground proximity.

SYMBOLS Measurements and calculations were made in U.S. Customary Units and are pre- sented in both the International System of Units (SI) and U.S. Customary Units.

A aspect ratio b wing span, m (ft) drag coefficient

CD

4 incremental drag coefficient, CD - CD,, lift coefficient incremental lift coefficient, CL - CL,, ...

, - C~ rolling-moment coefficient effective dihedral parameter C \ ~ Cm pitching-moment coefficient

% I yawing-moment coefficient

directional stability parameter

h s

static thrust coefficient

%

h height of wing quarter-chord above ground, m (ft) mordent of inertia about pitch axis, kg-m2 (slug-ft ) =Y 12 mass, kg, (slugs!

V velocity, mlsec (ftlsec) a angle of attack, deg 8 angle of sideslip, deg Y flight-path angle, deg 6f flap deflection angle, deg sweep angle at wing quarter-chord, deg hc/4 X taper ratio @ bank angle, deg Subscripts: o initial value w free-air condition - 1 , . . ,

Abbreviations : BLC boundary-layer control EBF externally blown flap L.E. leading edge ! ' S B upper-surface blown LONGITUDINAL AERODYNAMIC CHARACTERISTICS I N GROUND PROXIMITY

I

Both the EBF and the USB models tested in the Langley V/STOL tunnel were sting supported over a moving ground belt with a boundary-layer removal system ahead of the belt. However, the ground belt was not available during most of the EBF tests, although the boundary-layer removal system was always available.

The variation of lift coefficient with the height-span ratio h/b is presented with the boundary-layer removal system operating and with the ground belt on and off. hestilts dre presented for the EBF and USB models in the take-off configura- tion at constmt angle of attack through a range of thrust coefficient. The shaded area represents the conditions given by Turner (ref. 4) for which a mov- ing ground belt is required to simulate ground proximity correctly. The results indicate a slightly lower level of lift without the belt operating at free-stream velocity, although the trends are predicted very well. For the range of lift coefficient and height-span ratio, the ground proximity can be properly simulated with only a boundary-layer renovsl system in the test section.

The longitudinal characteristics in ground proximity d f the recently tested EBF (6f = 55O) and USB (6f = 6 0 ° ) models and the previously tested EBF (6f = 600) model are presented in figure 5. The longitudinal forces and moments are pre- sented as a function of h/b at constant angle of attack and at thrust coeffi- cients appropriate for a free-air lift coetficient of about 4.25. Both EBF configurations show similar lift losses in ground proximity. The USB configu- ration shows a slight lift increase in ground proximity before losi~g lift at the lower heights. The lift is concentrated at the inboard sections of the wing for the USB configuration, whereas the lift is spread more outboard on thp span for the EBF configuratioris. The differences in lift distribution may acco .t for some of the differences in lift in ground proximity, although there are ale.

differences in sweep between the configuraticns. Both EBF mzdels are swept back 25O at the quarter-chord and the US3 coiifiguration is dnswept. ' 1 1 1 three configurations, however, show a decrease in drag associated with the reduction in jet deflection angle as the ground is approached.

The pitching-moment data of figure 5 are untrimmed at different settings of tail incidence and are presented only to show the trends in ground proxi-mity.

The EBF models show nose-down moment increments in ground proximity and the pilot will have to exert trim control during landing. The trim control is usual.1~ o b ' a i n e d from a download at the tail, so that the trimmed lift loss in

The USB c o n i i g u r a t i o n i n d i c a t e s o n l y a s l i g h t ground p r o x i m i t y is i n c r e a s e d .

nose-down moment i n ground p r o x i m i t y . The sweep d i f f e r e n c e s between t h e con- f i g u r a t i o n s a r e p r o b a b l y t h e c a u s e o f t h e d i f f e r e n c e s i n p i t c h i n g moment.

P r e s e n t e d i n f i g u r e 6 a r e l i f t aird d r a g a s a f u n c t i o n o f h e i g h t - s p a n r a t i o f o r t h e EBF and USB c o n f i g u r a t i o n s t e s t e d i n t h e V/STOL t u n n e l .

R e s u l t s a r e p r e s e n t e d w i t h t h e f l a p s a t r e d u c e d d e f l e c t i o n s c o r r e s p o n d i n & t o . ike-off con- d i t i o n s a t two v a l u e s o f t h r u s t c o e f f i c i e n t . ht t h e reduced f l a p s e t t i n g s t h e l i f t a n 4 d r a g i n ground p r o x i m i t y change s i g n i f i c a n t l y less even though t h e l i f t l e v e l s a r e comparable t o t h o s e g i v e n i n f i g u r e 5 f o r *.he l a n d i n g c o n f i g u - r a t i a n . The pitching-moment c h a n g e s , a l t h o u g h n o t p r e s e n t e d , were ~ l s o r e d ~ c e d ! w i t h t h e lower f l a p d e f l e c t i o n s .

AVALYSIS OF LANDING APPROACH The c h a n g e s i n aerodyn:tmic c h i i r a c t e r ist i c s i n ground p r o s ! r . ~ i t y a r e most c r i t i c a l d u r i n g t h e l a n d i n g a p p r o a c h and t h e e f f e c t s o f t h e s e c h a n g e s a r e con- s i d e r e d i n b o t h s t c a d v - s t a t e and dvnamic :in;llvses. The a r r o d y n a m i c i n p u t s were t h o s e f o r c e s and n i ~ m e u t s measured i n t h e L a n g l r y V/STOL t u n n e l f o r t h e EBF and llSB c o n f i g u r a t i o n s i!\ ground p r o x i m i t y . 'Trimmed l i f t and dr,lg p o l a r s i n ground p r o x i m i t y were c o n s t r u c t c . d and n t y p i c a l v : l r i a t i o n of trimmed l i f t and d r a g i n ground prcjximity, riondimensionalized by tlie f r e e - a i r 1 i f t coef f i c i c u t , a r e sho..m i n t h e l e f t s i d r of i i g u r e 7. F l i g h t - p a t h a n g l e ;lnd a n g l e o f a t t a c k i n ground p r o x l m i t y f o r a c o n s t a n t - t h r b s t , c o n s t a n t - s p e e d iipproach c o r r e s p o n d i n g t o n trimmed l i f t c o e f f i c i r l n t of 4.0 and i n i t i a l f l i g h t - p a t h a n g l e of -6" a r e shown O i l t h e r i g h t s i d e o f i i g u r e 7 . The r e s u l t s i n d i c a t e ;hat. t h e r e d u c t i o n s i n d r a g more t h a n o f f - e t 3r.y o f tlre l i f t c h a n g e s t o i n c r e a s e t h e f l i g h t - p a t h a n g l e i n grvund p r o x i m i t y . The o n g i e of a t t a c k must b e reduced s l i g h t l y f o r t h e USB c o n i i g u r a t i o n and i n c r c n s e d s l i g h t l y f o r t h e EBF c o n f i g u r a t i o r ~ t o main- t a i n const:jnt trimmed l i f t . A s i m i l a r i ~ r c r c n s e i n ; l i g h t - p a t h i i n s l c i n ground p r o x i m i t y w i l s n o t e d ;lt t h e 1972 STOL T~clinc\logy Confcrtmce ( w f . 2 ) . Reduct i o n s of t h e f l i g h t - p a t h : ~ n ~ l c . t o z e r o i n a f l a r i n g m,i~~cuvt.r wcrc ~ ) l ~ s s i b l c wit11 .ippli- c ~ l t i o r l of e l e v a t o r ilnd power.

The s t e a d y - s t a t e n u ; l l y s i s o f t he l a u d i n g . ~ p p r o ~ i ~ - I r asstunics tlliit f o r c e ; I I I ~ moment c11;lngcs i n ground p r o s i n ~ i t y t r:inslatt. J i r e < - t l v i n t C) f 1 i g l ~ t - p i ~ t l l clri~npcs.

However, t h e mass and i n e r t i a l c l r a r a c t c r i s t i c s must b e c o n s i d c r ~ > d i n n dynamic Tlre mass a n a l y s i s t o p r o p e r l y s iniul:\tc t h e :1ctua1 a i rplitne 1ar1~iing . ~ p p r o a c h .

,and i n e r t i a l c l ~ n r c l c t c r i s t i c s of i t y p i c a l STOL a i r c r a f t ! m = 24 993 kg ( 1 7 1 1 sl(:gs) and Iy 334 642 kg-ml ( 2 4 6 819 s l i ~ g - f t l ) ) were u s e d a s i n p u t t o .4pproaclres wcre simu- t h e f ixed-base dynamic s i ~ n u l u t i o n progrlim d f r e f e r t ~ c e 5.

l a t e d o v e r a range of i n i t i n 1 f l i g h t - p a t h :.ngle, f r t . e - a i r l i f t c o e f f i c i e n t , rind The f ixcd-hasc s i m i r l a t o r r e s u l t s f o r n c o n s t a n t - t h r u s t c o n t r o l t e c h n i q u e .

'opproclch u s i n g o feedhnck c o n t r o l from t h e clc?vutor LC, m a i n t a i n s p c c ~ l a r e shown Free-<$ i r t t imrnel! l i f t coef f i- i n f i g u r e 6 f o r tht. EBF and USB couf i g u r ; i t ii.1.

r i e n t ts 4.27 and r e s u l t s a r c presented n s light-path t r a j e c t n r i c s f o r i n i t i a l A t t h e h i g h e r r a t e nf d r s t . c n t , t h e stt>odv- f l i g h t - p a t h a n g l e s of -b(' ;and -1.5".

s ; a t c r c s i r l t s da n o t have t inlc ti1 irlf lirt3ucC tllc f l i g h t ~ ~ t h , iiud n c l t l w r t h e from tlre i n i t i:t 1 f 1 i g h t p a t h . A s USB n o r t h c ICBF canf i g u r a t itjn dcv i:it~bs nrucll

I I a check on the results, the analysis was continued below gear touchdown height and eventually the flight-path angle was increased corresponding to the steady- state results. At the lower rate of descent, the changes in forces and moments in ground proximity have a chance to develop and both configurations perform self-flaring maneuvers at ground heights near 6.1 m (20 f t ) .

i '."

t .'

The flight-path changes in grou~d proximity are dependent on the particular type of feedback control system used. The results for a constant-thrust approach . . , using a feedback cont~ol from the elevator to maintain attitude are show. in fig- ,%: $ -, ure 9.

Neither configaration deviates from the steep flight path in ground prox- .,.: !

..

imity. At the lower iniuial rate of descent, the USB configuration performs a

, * ..-I - $<

self-flaring maneuver near 4.9 m (16 f t ) . The EBF configuration does not flare, , * ' . J although it never falls below the initial flight-path trajectory. ( ..:.I I , ! '

i -I

The drag reductions in ground proximity common to both the EBF and the ; ' j USB configurations are most important in determining steady-state flight-path increases in ground proximity. The lift changes are configuration dependent 1 . ; and the extent to which the steady-state results are experienced on the actual 8 . j airplane depends on the initial flight-paih angle and the particular type c f 1 1 7 i f jback control system used. None of the above analyses consider the flare n : . . ,uver which can be effected by application of either power or elevator.

S I D E S L I P AND ROLL I & GROUND PROXIMITY i - 1 I '

? ; .-,k

: . 1 The EBF rnodel in figure 2 was tested in the Landley V/STOL tunnel to deter- mine the effect of sideslip angle, bank angle, and combined sidesifn and bank angles jn ground pr~ximity. Results in figure 10 for the pure sides1;p condi- - i

tion in ground proximity are presented as effective dihedral (-clg) and direc-

t ional stability (C+,@) as functions of height-span ratio for several thrust -1.

coefffcients at constant angle of attack. Through sideslip angles of tlCo, the '1 , EBF model indicated strong stability with little change due to ground proximity.

The direcEions1 stability shows the expected increase with thrust because of the

I '

increased dynamic pressure at the tail. The effective dihedral is increased

slightly at the lower hcight-span ratios. I

I The effctct of roll in ground proximity is presented in figure 11 at a con- , , I stant angle of attack and constant thrust corresponding to a free-air l i f t coef- : ,I

f \ 1

f icient near 4.0. Rolling z.oment as a function of height-span ratio is presented fer various roll attitudes. The pure roll case is shown on the left side of the 1 , . ; , '-: figure, and positive bank angles corresponding t . o right wing down give laige ! 1 .

, :?

unstable rolling monents at height-span ratios near gear touchdown height. The !

co,,.bined roll and sideslip condition is given i . n the right side of the figure. 1 . 1 .

The increment in ro!.ling moment due to sideslip at 4 = O0 arises from the , ,:I

strong positive effective dihedral. Positive a r t d negative roll attitudes give unstable rolling moments at height-span ratios near gear touchdown height. The unstable rolling-moment increment due to roll. attitude is about the same at i.

6 = O0 and 6 = -10". indicating no L : i g l r ~ f icant yaw-roll coupling.

-.

ENGINE-OUT CHARACTERISTICS IN GROUND PROXIMITY The lateral-directional characteristics of the twin-engine USB concept in the event of engine failure during a powered-lift approach have been a matter of some concern.

A concerted effort has been made to develop lateral control sys- tems for this concept sufficiently powerful to trim out lateral as:mmetries due to engine failure. Figure 12 illustrates the effect of engine failure in ground proximity. Results are presented in terms of yawing moment and rolling moment as functions of angle of attack for both the free-air condition (h/b = m) and in ground proximity (h/b = 0 . 2 ) .

The results presented are for the landing con- figuration ( d f = 600) with the left engine inoperative and with the right engine

a t full thrust ( c ~ = 1.80). @?he left side of figure 12 illustrates engine-out

characteristics with all controls neutral. Surprisingly, the rolltng-moment asymmetry was unaffected by ground proximity; however, the adverse yawing moment to engine failure was reduced at the higher values of angle of attack.

due The right side of figure 12 illustrates 3 possible solution to the lateral asymmetry problem due to engine failure. The aileron on the engine-out side has been drooped 60° and is augmented with blowing boundary-layer crz~trol (BLC).

Also, the entire leading edge of the wing on the engine-out side is augmented with BLC to prevent flow separation at higher angles of attack. The results indicate that most of the rolling-moment asymmetry due to erzgine failure can be trimmed out with this lateral control system and that ground proximity had Adverse yawing essentially no effect on the rolling-moment trim capability.

moment due to this lateral control is, in general, very slight (compare left- and right-hand yawing-moment data), and ground proximity causes the same reduc- tion in yawing-moment asymmetry at higher v.,iurs of angle of attack as was observed with lateral controls neutral. It shouid be mentioned that the twin- engine USB concept requires a double-hinged rudder capable of handling the yaw- ing moments due to engine failure during take-off, and such a rudder control would be more than adequate to trim out the yawing moments shown in figure 12 for the landing configuration.

CONCLUSIONS The following conclusions are drawn from the recent wind-tunnel investiga- , t ions of powered-lif t configurations : f :; .; . .

; . !

1. Drag reductions in grounJ proximity are common to both EBF and USB con- figurations, whereas changes in lift are configuration dependent.

i ]

2. The extent to which the predicted steady-state flight-path increases in ground proximity are experienced on the actual airplane depends on the initial flight-path angle and the control technique used.

3 . . Lateral-directional charcct2ristics due to sideslip are unaffected by ground pro-~imity, whereas roll nttitudes give unstable rolling moments near gear touctldown height.

4. Engine-out chzracteristics both with and without corrective control are unaffected by groucd proximity.

REFERENCES 1. Vogler, Raymond D.: Wind-Tunnel Investigation of a Four-Engine Externally Blowing Jet-Flap STOL Airplane Model. NASA TK D-7034, 1970.

2. Hassell, James L., Jr.; and Judd, Joseph H.: Study of Ground Proximity Effects on Powered-Lift STOL Landing Performance.

STOL Technology, NASA SP-320, 1972, pp. 193-213.

3. Phelps, Arthur E., 111; Johnson, Joseph L., Jr.; and Margason, Richard J.: Summary of Low-Speed Aerodynamic Performance of the Upper-Surface-Blown Jet-Flap Concept. Powered-Lift Aerodynamics and Acoustics, NASA SP-406, 1976. (Paper no. 4 of this compilation.)

4. Turner, Thomas R.: Endless-Belt Technique for Ground Simulation. Con- ference on V/STOL and STOL Aircraft, NASA SP-116, 1966, pp. 435-446.

5. Grantham, William D.; Nguyen, Luat T.; Patton, James M., Jr.; Deal, Perry L.; Champine, Robert A.; and Carter, C. Robert: Fixed-Base Simulator Study of an Externally Blown Flap STOL Transport Airplane During Approach and Landing. NASA TN D-6898, 1972.

Figure 1.- EBF configuration t e s t e d in 1 7 - f o o t test section of LangPey 300-MPH 7- by 10-fooc tunnel. A 7.0; h = 0.3.

h , / 4 = 2 5 O ; Figure 2.- E B F uonfigurntioa t e s t e d in Lanbley V/STOL tunnel.

A = 7 . 3 ; A e i 4 = 2 5 ' ; X = 0.4.

Figure 3 . - US3 configuration tested l n tangley V/STOL tunnel.

> = 0 . 3 .

A = 8.2; A 4 4 = 0 ' ; Q v ,

EBF MODEL + 0 US8 MODEL

- - hlb Figure 4.- Effect of moving ground belt on l i f t In ground proximity with bo~ndary-layer removal system operating.

a - 5O.

USB MODEL #

------

-a 0 1 - --

gw"x-Doueu-sLomo F L I P

-1 L (REF. 1) -1 I

3.6 hlb Figure 5.- Ground effect on longitudinal aerodynamics of landing ccnfigurations. a = 5O; C,, = Constant; C L , , = 4 . 2 5 .

1 EBFMOQEL + = n o 0

U S B MODEL 4 = 20'

bib hlb Figure 6.- Ground effect on longitudinal aerodynamics of take-off configurations. a = 5O.

Rl~FR-Oi>lJCI~l?,I'TY OF THE

OHItiL\ttiL i1A(;li: IS POOR

0 .2 .4 .6 hlb hlb Figllre 7.- R e s u l t s from s t e a d y - s t a t e a n a l y s i s of c o c s t a n t - t h r u s t , c o a s t a n t - s p e e d l a n d i n g approach. C L , t r i m = 4.0.

LANDING GEAR 20 ABOVE GROUND, rn H E I G H T O F L A N D I N G GEAR ABOVE GROUND, 11) U S B

-

-

m

.I----

-1. ~ O G L I DE SLOPE

- - _

- -

I L L Y r - - 1 0 100 200 300 400 500 600 700

?:ti

LATERAL DISTANCE FROM I N I T I A T I O N , m F i g u r e 8.- R e s u l t s from f i x e d - b a s e s i m u l a t o r a n a l y s i s o f r - o n s t a n t - t h r u s t , c o n s t a n t - s p n e d l a n d i n g approach u s i ~ g e l e v a t o r c o n t r o l .

(cL, t r i m ) m = 4 - 2 7 .

t - 2 - 7 . 4 ' r i . . . I 1 . ' 1 , 1 j ; .

RFn"Ilr!i-' l .\.'I!$ i : 1 . 1 1- OF (fa*

HEIGHT OT C R r c ; r ; ~ i.!Lil;~ p m

LANDING GEAR 20 -

ABOVE GROUND, m

-

r = -1.5'

HEIGHTOF 15 LANDING GEAR ABOVE GROUND, 10 EBF m -1.5' GLIDE SLOPE 0 100 200 300 4 M 500 600 700 LATERAL D l STANCE FROM I N IT IATION, m Figure 9.- Results from fixed-base simulator analysis of constant-thrusr, constant-attitude landing approach using elevator control. ( c ~ , ~ ~ ~ ~ ) , = 4.27.

-.01 L

L A * . . -. -

Q

0 . 2 .O .6 . J 1.0

hlb Figure 10.- Static lateral-directional characteristtcs in ground proximity. P r i ; ' --slotted EBF model; 6f = 4 0 ' ; a = 5 '

Figure 11.- Effect of bank angle in ground proximity. Triple-slotted EBF model; 6f = 40°; a = so; C,, = 2; CL,- = 4 . 0 . I LEFT ENGI b!E CUI LEFT ENGINE OUT LEFT AILERON DR'JOPED 60" + ELC -2 0 2 4 6 8 1 0 - 2 0 2 4 6 8 1 0

~ 7 8 - 2 4 0 5 6 -

DISTRIBUTED UPPEX-SURFACE BLOWING CONCEPT Paul G. Fournier and Paul L . Coe, Jr.

-- NASA Langley Research Center

SUMMARY -

A low-speed investigation was conducted in the Langley V/STOL tunnel to determine the powered-lift aerodynamic performance of a distributed upper- - ,- ,.I) .- 1 surface-blown propulsive-lift transport model. The model used blowing slots ; : < ' . a I .-,fi . ...: across the span of the wing to produce a thin jet efflux near the leading edge , ! *:$.

and at the knee of the trailing-edge flap (internally blown jet flap). These , . . ; % concepts have both good propulsive-related lift and low drag-due-to-lift char- j : .:$* .- / The leading-edgt acteristics because of uniform spanwise propulsive thrust.

blowing concept provides low-speed lift characteristics which are competitive with the flap-hinge-line blowing coacept and does not require additional leading-edge treatment for prevention of abrupt stall.

INTRODUCTION i ! '. ..

; : : : t 1 . .

Several propulsive-lift concepts have been investigtted recently in efforts to develop a quiet short-take-off-and-landing (STOL) aircraft. The upper-surface- blown (USB) jet-flap concept appears to offer an attractivc! soiution for a quiet STOL aircraft. Most of the investigations to date have used configurations that

-. . .., ., ... . - c

direct the efflux from discrete engine nozzles over the wing upper surface and high-lift system to provide increases'in lift by means of Coanda turning of the

I 1 : s

jet (ref. 1 ) . These results indicated that the propulsive lift capabilities . : : : -*.ion to the high- were greatly dependent upon the nozzle geometry and their * turning.

lift system, with thin well-spread jets giving the best Co; Another version of the USB concept utilizes full-span slot nozzles near the These slot nozzies are beneficial in several ways.

leading edge and flap hinge.

First, they improve the aerodynamic performance by distributing the ~ropulsive efflux in the spanwise direction, which improves the induced lift and reduces the induced drag relative to that obtained with discrete USB nacelles. Second,

1 *.I

this arrangement reduces the propulsive noise by the use of very high aspect- ,. 2-1 ratio nozzles which produce high-frequency noise that damps out quickly and by shielding the ground with the wing.

Figure 1(~) shows a photograph and a sketch of this distributed blowing concept applied tc a subsonic transport configuration with an aspect-ratio-6.8 swept wing tested in the Langley V/STOL tunntl. The wing had internal plenums which supplied air to full-span slct nozzles located along the wing leading edge ...

and along the flap hinge line. It is anticipated that this arrangement would , .. .-: ..

. . .;.I

. .

I ,I.

A

approach the upper limit for propulsive induced lift. However, such an arrangement would introduce weight and volume penalties and would require sys- tems studies to determine if it is a practical coxept for subsonic traasport applications.

SYMBOLS

- 4

A aspect ratio

C~ drag coefficient , Drag/qS

c~ lift coefficient, Lift/qS

CL, js- jet-reaction lift coefficient circulation lift coefficient due ro power - i

I ! ...I

clJ thrust coefficient , Thrust/qS c wing chord 9 free-stream dynamic pressure S wing area a angle of attack, deg flap deflection measured streamwise, deg (Dual notation indicates 6f deflection of forward element with respect to the basic airfoil chord line, followed by the deflection of the rear element with respect to the chord line of the forward elcment. See fig. l ( b ) . ) DISCUSSION The basic data for distributed blowing over the flap and at the leading edge for several flap deflections are pres :nted in figure 2 for a noainal value of C , , of 2.0. These data show that with blowing over the deflected trailing- edge flaps and with no leading-edge high-lift device, there was an abrupt stall near a = 12O. However, for the leading-edge blowing configllration, there was no stall through the angle-of-attack range tested.

The theoretical minimum drag-due-to-lift curve (ref. 2 ) is plotted along with the basic data for both blowing concr:pts in figure 2 and shows that 'ihe basic data approach this curve quite well.

T variations of lift coefficient with angle of attack fcr the two blowLllg concepts are presented in figure 3 . All configurations had two element flaps

(fig. l ( b ) ) with the forward and aft elements deflected 450 and lSO, respec- t~vely. A comparison of the leading-edge and flap blowing indicates that at [ low angles of attack, the flap-hinge-line blowing results in somewhat higher values of CL than the leading-edge blowinn, Also, if the leading edge is dropped to 300 on the flap-hinge-line blowing colfiguration, it performs as well as, or better than, the leading-edge blowing configuration at the higher

/

1 angles of attack.

1: i

Figure 4 presents the propulsive-related lift as a function of thrust coef- ficient for the distributed leading-edge blowing concept and a conventional , .

USB concept. The propulsive-related lift is the combination of the jet-reaction , + lift CL,~, and the additional circulation lift due to 2ower C L ~ . The con- ventional USB concept (ref. 2 ) used rectangular exhaust nozzles h a - . ? i n g an aspect , .

racio of 6. The distributed leading-edge blowing concept produce2 much more I " .

propulsive-related lift than the USB concept.

! ; L 1 - i ' i .

i . . . , CONCLUDING REMARKS : - .

The distributed blowing concepts, with blowing at either the leading edge or flap hinge line, have both good propulsive-related lift and low drag due to lift because of the uniform spanwise propulsive thrust. The leading-edge blow- ing concept provides low-speed lift characteristics which are competiti~, with the flap-hinge-line blowins concept and does not require additional leading-edge treatment for prevention of abrupt stall.

REFERENCES 1 . Sleeman, William i., Jr.; and Hohlweg, William C.: Low-Speed Wind-Tunnel Investigation of a Four-Engine Upper Surface Blown Model Having a Swept NASA TN D-8061, 1975.

Wing and Rectangular and D-Shaped Exhaust Nozzles.

2. McCormick, Barnes W., Jr.: Aerodynamics of V/STOL Flight. Academic Press, Inc., 1967.

(a) Flap-hinge-line blowing. Cv = 1.9.

(b) Leading-edge blowing.

C,, = 2.0.

Figure 2 . - Effect of flap deflection on l i f t and drag c o e f f i c i e n t s for model with distributed blowing concepts. Tail o f f ; 6f = 45O-oO; c, = 2 .

Figure 3.- Variation of CL with a for the distributed blowing concepts. tif = 45O-15O; C , , = 2.0.

Figure 4.- Variation of propulsive-related lift with thrust coefficient for the distributed leading-edge blowing concept and a conventional USB concept.

CRUISE AERODYNAMICS OF USB NACELLE/WING GEOMETRIC VARIATIONS* John A. Braden, j ~ b n P. Hancock, and Kanrreth P. Burdges Lockheed- Georqia Company SUMMARY Experimental results are presented on aerodynamic effects of yeometric variations i n USB nacel le confiqurations at hiqh-speed cruise conditions. Test data includes both force and pressure measurements on two- and three-dimensional models powered by upper-surface blowing nacelles of varying geometries. Experimental resu Its ure provided on variations in nozzle aspect ratio, nozzle boattail anqle and multiple-nacel le installations. The nacal les are ranked according to aerodynamic drag penalties as we1 l as overall i~stailed drag penal- ties. Sample effects and correlations are shown for data obtained with the pressure model.

INTRODUCTION Use of upper surface blowing (USB) engine installations, illlistrated in figure 1 , has been demonstrated as a viable means of STOL hiqh-lift augmentation by both industry and qwernment sponsored research over the past several years. Such studies have shown the system to be attractive for STOL application from a number of viewpoints. These include generally favorable acoustic characteristics for the terminal area environment, reasonably practical structural compatibility with the airframe, and acceptable flexibility for integrat- ing with basic operational systems or sub-systems. These, of course, are in addition to the recognized potential for good STOL performance. h contrast to the low-speed accounta- bility of the USB system, a comparable data base for the high-speed cruise regime has been lacking. To f i l l this need, the Lockheed-Georgia Company, under contract to the NASA, has conducted experimental irwestigations wherein USB nacelle/wing geometries and oper-

ating conditions are systematkally varied i n the 0.5 s Mo s 0.8 cruise regime. The basic

goal of this parametric investigation i s to define those geometric properties and operating conditions indic~otive of minimum cruise-drag penalties and from which more refined USB configurations can evolve.

* Work performed under contract to the NASA; Contract No. NAS1-13871; "Cruise

Performance of Upper Surface Blowing Conf igurutions . "

The present paper provides a brief over-view of the experimental program, currently still in progress, along with preliminary findings believed to he of tqeneral interest. The experimental work encompasses force-test e\.aluations, surface pressure me~surements, and wake surveys behind powered configurations. A companion analytical effort I s involved in the: basic prugsam, but an evaluation of math model capabilities via experimental correla- tion must awai; a more thorough examination of the test results.

SYMBOLS . , ..

l%i . : :.;! I

Values are given in both SI and U.S. Customary Units. The measurements and calcu- , ' ; I lations were made in U . S . Customary Units.

F i - . I

I .

2 . 2 nozzle exit area, cm (In ) nozzle aspect ratio, w / A N wing semispan, cm (in .)

b/2

4 c wing chord, cm (in .)

l i f t coefficient thrust coefficient incremental drag cceffisient A ' 0

drag , N (ib)

nozzle stagnation pressure, N/m (lb/in2) nozzle pressure ratio freestream Mach number 2 2 freeatream static pressure, E. 'm (Ib/in ) 2 2 freestream dynamic pressure, N/m (Ib/in ) 2 . 2 wing area per semispan, cm (In ) gross thrust, N (Ib)

I . * . - i - u -.A . d -- C_11., ! ..

I w nozzle ~ i d t h , cm (in .)

X chordwise distance from wing leading edge, cm (in.)

angle of attack, deg , .\ . ..

, . . I . ...

..iJ 8 boattail angle, deg

jet turning angle measured statically, deg 7 thrust efficiency for wing/r.acelle combination,

T~~~~ / T~~~~

Subscripts: A, Aero aerodynamic F frictifm . .

IN1 interference

:I ISOL isolated

1 local

M, MEAS measured nacelle *' , TOT total EXPERIME NTAL OBJECTIVES The objective of this experimental program i s to establish a transonic experimental data base covering a wide dr~riation of nacelle geometric parameters. An extensive airay of nacelle/wing geometric cmfigurations were developed so that experimental evaluation of the transonic force and surface pressure charucteristics cculd be made. The geometric configurations were tested over a wide range of Mach number, angle of attack, and nozzle pressure ratio to establish an ex:ensive data base from which summery "effects" arc deve- loped. Tho primay geoms?r.ic variations in nacel la configurations, illustrated in figure 2, for which aerodynamic "effects" data are qenerated and arodynamically ranked are: o nozzle exit aspect ratio o nozzle boattail angle

I

r

I . .

..

6 .

. : . . . ! - o chordwise /vertical/spanwise positioning o size and number of nacelles Secondmy experimental objerti*:ss include the efftcts of:

t

o wing Fweep i o wing camber modifications

Q

f 4 o inlet flow-field effec+s f i

.'.I o jet deflectors I

1 i!

o wing/nacellc. f i l leting and streamlining !

For the present discus*ion, only selected combinat;cns of these ge~metric variations w i l l be csnsidered. In parti "ular, nozzle aspect ratio, boattail angle and mu1 tiple-nacei le interference vli l l be coverell.

I CFF rUNNEL TESTS Test Facility The experimental progra~n i s being conducte.' in the Lockheed-Georaia Compre~sible Flow Facility (CFF), which i s a variable porosity, blowdown wind tunnel r . 2 ~ r c 3). Oper- ating ranges ore 0 . 2 ta 1.2 Mach number and up to 144 x 1 0 6 / ' ~ Reyn~!f;s ,*umber in the trun- sottic speed range. The test sertion, which i s 50.8 cm ('LO in.) ;.vide, 7 1 . 1 cm (28 ir,.) !ligh and 182.9 crn (72 in.) long, can be equipped with porous or solid walls to match particular test requirements. Model engine air supply i s provided by an independent 2.068 M N , ~ ~ (300 lb/in2) source. For p w e r e d force testing, air ir . :oplied to the mcdel throuph the force balanLe by a bellows arrangemetlt.

Test Conditions

The Mach number range of in+erest in tile experimental s f f c d i s n.5 Mo 0.7 +o

the unswept wing and 0.6 s Mo s r3.80 for the swept wing moooi . Maximum nozzle prt-, -

sure rat;= (HiPo) up to about 3.0 were tested over most 3f thr-se speed ranges. Mod51 ..:: . I 1 , ' * . a y l e of attack was v a i e d from 0 degrees to 5 deg-ees encompassing a l l normal cruise .

settings. F!ow visualization using titanium dioxide combined with oil was employed c?xten- .I . .

sively to help understand the force and pressu : test results.

. ", 1 ' 4 , I - , ~ . - -. ,, -??.v - . r . 1 . I . ' . ..

, ---, 1 ;

I -- i

, -

* . I - - * .), -. 4.

Models The test configurations are composed of a large number of interchangeable nozzles wing. There are two basic wings, swept and unswept, which have a wing conversion from 2-D to 3-D configurations. The two-dimensimal pressure models span the 50.8 cm (21 in.) horizortal width of the tunnel. For 3-D force

I . i

s mounted vertically on the balance system i n the tunnel floor.

The semi- span-wing (b/2 = 50.8 cm (20 in .)) then spans 70 percent OF the tunnel height (71.1 cm ! .;: (28 in .)). Examples of the two force models, which illustrate unswept and swept 3-D wings,

1 . .s

are provided in f k -es 4 and 5, respectively.

f Figure 6 illustrates the build-up of a two-dimensional pressure model, using the un- swept wing, alcng with the traversing woke role. A supercritical-type of wing section i s , , used for both the unswept and 25 degree swept wing. This airfoil section hos a streamwise thickness I-atio of 16 percent tor the unswept wing application and 14 percent for the swept . - . -i , wing case; the wing chord I s 17.8 crn (7 i n .) in both instances. The wing design Mach 1 4 .

numbers are about 0.7 and 0.8, respectively. 1 .

i I !

. , As illustrated, the nozz.G . r ru?ply at 2.065 M N / ~ ~ (30C lb/in2) i s routed through i the undetviiq duct and into u plenum formed by a faired-over forebody. The air i s exhausted I i through a choke plate and exits fran the raor71e at pressure m t l ~ s (H/Po) ranging rip to a maxi- , * ..

I ; mum of 3.0. The some forebody i s used with a number of interchangeable nozzles which :lave : ! C , I different exit shapes, but the same discharge area. Unpowered c3nfigumtions may be bui t up

i .i

t y suistitution of a flaw-thrwgh inlet for the h i r e d farebody.

1 i i L 1. ' Model instrumentation includes surface pressure taps at 5 spanwise positions on the wing aitd a1o.g the nozzle upper surface. A trwersing wake rake provides the capability i i 1 for i m p 1 ing jet profiles and f ~ i evaluating momentum losses in the wing/nacel le wake. , RESULTS AND DISCUSSION . ,

I r ! i

The forct -test phase of :he exper:mental program has vscently been completed and . . . . -. 1 .

I . i pressure testing i s currently in progress. For this reason, the present discu*sions en.phcrsize ' .

I t results with limited reference to pressure tes:ing, except where available data s a

j

Aerodynamic Ranking i i

i ! i i