Experimental investigation on a cropper delta wing with edge blowing

H E C O L L E G E O F A E R O N A U T I C S

C R A N F I E L D

E X P E R I M E N T A L I N \ ^ S T I G A T I O N ON A C R O P P E D

D E L T A WING WITH E D G E BLOWING

b y

A . J . A l e x a n d e r

REPORT N O . 1 6 2

June. 1963.

T H E C O L L E G E O P A E R O N A U T I C S C R A N P I E L D

Experimental I n v e s t i g a t i o n on a Cropped D e l t a Wing with Edge Blowing

A . J . Alexander, M.Sc,, P h . D . , A.F.R.Ae.S.

SUMMARY

Low speed wing, tunnel tests have been made on a 70 cropped delta wing with edge blowing both in the plane of the wing said at a downward deflection angle of 30 .

The tests include aix-ooii5)onent force and moment measurements, the distribution of static pressure at four ohordwise stations, and quantitative measurements of the flow in the leading edge vortex.

At a constant incidence, blowing increases the size and strength of the leading edge vertices and moves the vortex cores outwards and ufJwards. Blowing also tends to sujïpress the secondary separation due to the entrainment effect

of the jet. Blowing fvova. the streamwise tips and trailing edge was relatively

ineffective and most of the tests were made with blowing from the swept leading edges only, witji tips and trailing edge sealed.

A C IB

The lift magnification due to blowing, -r;— , decreased with decreasing incidence and increasing C„, but for small C and high incidence, values of almost three were reached with leading edge blowing in the plane of the wing, With a suitable blowing distribution the conical nature of the flow is not

disturbed and these lift increases are obtained with only a small centre of pressure movement, making the scheme particularly attractive for tailless aircraft.

o

10 to the swept leading edge, values of about two were obtained at low 0 and ^ %

high incidence, Providing conicauL flow is maintained, will be a function

a ^^

of ^ only and will increase as — increases. Thus the use of leading edge

a. A.

blowing with full engine thrust on highly swept wings will insrease cruise lift to örag ratios and reduce landing and taks-off speeds and distances by a substantial margin,

CONTENTS

Summary

List of symbols

1. Introduction 1

2. Model and experimental method 2

3« Discussion of results 4

4* Conslusions 19

Acknowledgements 21

5. References 22

oc. geometrie wing i n c i d e n c e

0 angle of s i d e s l i p

e j e t sweep a n g l e , see f i g , 7

fli edge droop a n g l e , see f i g , 7

b wing span = 1,62 f t . o r o o t chord = 3«35 f t , aerodynamic mesin c h c r d = o v. /o = 2,60 f t . p , q malnstreajn s t a t i c a M dynamic p r e s s u r e s p s t a t i c p r e s s u r e 8 wing semi-span x , y , z body a x e s , see f i g , 8

y , z spanwise p o s i t i o n and h e i g h t of vortex core H t o t a l head measured i n l e a d i n g edge v o r t e x K c o t a n g e n t of l e a d i n g edge sweep angle 0 o r i g i n of body axes a t 0,50 co

S wing a r e a = 3«60 s q . f t ,

P - P Q

0 static pressure coefficient =

, -, . , ^«. . J. total momentum ejected

0 blowing momentum coefficient = g l i f t

0 , l i f t c o e f f i c i e n t = -^—s

L q^.S

LQy l i f t increment due t o blowing

0_ drag c o e f f i c i e n t = — - - | D q , 0 ^o ^ . ^ « ««• . ^ c r o s s winft f a r c e 0_ c r o s s - w i n d f o r c e c o e f f i c i e n t = o "" 0 q . 3

0/ rolling moment coefficient about x axis = ^ ^ , * q .S,b.

o

O pitching moment coefficient about y axis (O.5O Co) = ^—-—5^-'"°°'^" ^o

O yawing moment coefficient about z axis (O.5O Co) = '^ ^ q >.

1» Introduction

Maiy attempts have been made in recent years to iaiprove the low speed characteristics of aircraft. Conventional trailing edge flaps give useful increases in lift uiiich can be supplemented by a boundary layer control system but trimming the resulting large pitching moments reduces the effectiveness somewhat. The problem is particularly severe on highly swept tailless aircraft where lift coefficients are low in general and the use of trailing edge controls makes trailing edge flaps ineffective,

The redaction of the lift curve slope with aspect ratio is predicted by various theories both for attached flow and vo-th leading edge separation * ' but it is clear from the theory and from experimenbal evidence that the

existence of leading edge separation and the associated leading edge vortices contributes to the lift and reduces the adverse effect of decreasing aspect ratio,

Since the leading edge vortices, or tip vortices in the case of low aspect ratio unswept wings, play such an impox'tant part in determining the flow pattern, it is clearly desirable to control their development in order further to increase their favourable influence, A promising method of control is to emit from the appropriate edges a jet in the form of a thin sheet. This jet (vortex) sheet rolls up in a manner similar to the rolling-up of the free vortex sheets and increases the strength of the resultant vortex, thus increasing the non-linear lift.*

(4) Tests using this device have been made on a low asi:iect ratio straight wing^ , a 40 swept v/ing , both with tip blowing, and with leading edge blowing from a 70 delta wing . In all these tests, edge blowing in the plane of the wing resulted in increased lift at constant incidence. The effect of blowing on the

unswept wing was to change the spanmse lift distribution from elliptic to approximately constant loading due to the ability of the jet sheet to support a pressure difference. In the case of the win^s vri.th swept leading edges, the size and strength of the leading edge vortices was increased giving increased non-linear lift. In at least, the increase in lift was obtained with little change in the longitudinal static stability over the greater part of the usable C^ range,

* Blowing increases the strength of the leading edges vortices at a given incidence and hence the adverse pressure gradient along the vortex core. Thus the incidence at which vortex breakdown occurs will be less with edge blowing. See Para,3.7 for discussion.

- 2

Although blowing increased the lift at constant incidence in the above

' \

t e s t s t h e l i f t m a g n i f i c a t i o n -;;— was s m a l l and t h e aim of t h e p r e s e n t t e s t s was t o a c t e n d euid amplify t h e e x i s t i n g e x p l o r a t o r y work and, i f p o s s i b l e , t o

improve on t h e r e s u l t s . The p o s s i b i l i t y of u s i n g l e a d i n g edge blowing t o improve l i f t - t o - d r a g r a t i o s was a l s o i n v e s t i g a t e d , and t h e t e s t s covered a reinge of i n c i d e n c e s and blowing momentum c o e f f i c i e n t s a p p r o p r i a t e b o t h t o c r u i s e and t a k e - o f f c o n d i t i o n s . C r u i s e v a l u e s of C c o u l d be o b t a i n e d , but only a t low si^eeds, and the e f f e c t of Mach Mimber c o u l d not be a s c e r t a i n e d from t h i s s e r i e s of t e s t s ,

2 , Model and experimental method

The model i s shown mounted i n t h e wind t u n n e l i n f i g , l . I t i s a 70

swept d e l t a wing w i t h cropped t i p s , of chord equal t o one t h i r d of t h e r o o t chord, and has an a s p e c t r a t i o of 0 , 7 3 . '^he main body i s a hollow gunmetal c a s t i n g a i d d e t a c h a b l e b r a s s edges form a c o n t i n u o u s blowing s l o t round t h e p e r i p h e r y (of c o n s t a n t ^^ddth 0.040 i n . ) except foi' a small r e g i o n near t h e apex. The model i s of rhombic c r o s s - s e c t i o n , the t o t a l edge angle bn both l e a d i n g edges and t i p s i s 20 , t h e t r a i l i n g edge angle iB 15 • P r e s s u r e p l o t t i n g s t a t i o n s were l o c a t e d a t 0.35c . 0.49c . 0.65c and 0.87c from t h e apex, anti each s t a t i o n c o n s i s t e d _{o o* o o -t^ »} of two c o n t i n u o u s t u b e s each spanning h a l f of t h e wing. T h i r t y - s i x s t a t i c p r e s s u r e h o l e s ( 0 , 0 2 0 i n . d i a . ) were d r i l l e d a t each s t a t i o n t o enable t h e

spanwise s t a t i c pcéBsxxce d i s t r i b u t i o n t o be a c c u r a t e l y d e s c r i b e d .

The t e s t s were made i n t h e College of Aeronautics 8 f t . x 6 f t . low speed wind t u n n e l , vrith t h e model supported on a Warden type six-component b a l a n c e . High p r e s s u r e a i r v/ a s f e d t o t h e model a t t h e balance v i r t u a l c e n t r e through t h e hollow support s t r u t , and c o n s t r a i n t s were kept t o a minimum by t h e use of a f l e x i b l e c i r c u l a r r i n g - m a i n f e e d , ( f i g . 5 ) . A r o t a r y s e a l a t t h e c e n t r e enabled the model t o be yawed,

The r a t e of mass flow of a i r t o t h e model m., was measured using sharp-edged o r i f i c e p l a t e s i n t h e main f e e d p i p e . There was, of c o u r s e , no l o s s of a i r as with an a i r b e a r i n g system. The j e t t o t a l head d i s t r i b u t i o n was measured j u s t o u t s i d e t h e s l o t and i t s momentum c a l c u l a t e d on t h e a.ssumption of i s e n t r o p i c flow. With blowing confined t o t h e swept p a r t of the l e a d i n g edge, t h e d i r e c t con^onent of t h r u s t c o u l d be measured on t h e balance and t h e j e t momentum c a l c u l a t e d .

Owing to the difficulty of measuring the total head distribution accurately and of making allowances for induced effedts on the balance measurement of thrust, there is a possible - 5fo inaccuracy in the values of 0 „ ,

In order to taper fend direct the jet, thin grooved perspejc strips were inserted in the slot (fig,4)» Some preliminary tests with directed blc^dng showed that in order to obtain a given direction of the jet sheet, it was necessary to break up the jet into a large number of small jets each capable of individual direction. The individual jets recombined veiry close to the slot into a homogeneous sheet. The perspex strips were 0.5in. wide and 0,040in. deep with 0.020in, wide grooves having 0.040in. spacing out at the

appropriate angle to the leading edge. By varying the depth of the saw out

it was possible to vary the momentum ejected and thus obtain an approximately linear increase in momentum from the apex to the leading edge-wing tip

junction (fig,8),

ITX this series of tests, all fcrces are referred to wini axes and moments to body axes (fig.9). Pitching and yawing moments are referred to a point on the mod.el centre line at one half root chord.

In tests with blowing, the high pressijre air caused the flexible ring main, (fig,5), to distort slightly and this induced additional forces and moments. To ccorrect for these changes the model was removed periodicailüy and a calibrator (fig,2) was attached to the balance strut. The calibrator consisted of a short length of pipe feeding air to the adjustable gap between two flange©. The gap was set such that with any given model configuration the rate of mass flow was the same for a fixed control pressure. Since the air was emitted radialDy from the virtual centre of the balance, the only forces and moments present should be those due to the distortion of the ring main. In order to allow fca: the slight

inaccuracies in manufacture, the balance measurements were taken with the calibrator in a fixed position and then rotated through 130 . The average of the two sets of readings was taken to be the balance constraint correction due to blowing. Typical corrections, for a control pressure of 15 p.s.i, gauge at the orifice plates

(C = 0.178) with blowing from all edges are given below, coefficients based on V«, = 100 ft/seo,

4

--0.51 l b s ,

-0,06 l b s .

+0,55 l b s ,

+1,17 l b s ,

+0.22 l b s ,

+1.17 l b s .

°L

%

°0

Ce

0

m

C

-.007

- , 0 0 1

+.008

+,017

+,002

+.017

l i f t

Drag

Cross-wind force

Rolling moment

Pitching moment

Yawing moment

The moments given here are referred t o the balance v i r t u a l centre and

wind axes,

Wind tunnel constraint corrections have not been applied t o the present

r e s u l t s since no suitable corrections ai'e available, but conventional corrections

are small, Ao = 0.5** for a = 25° and AC = 0,009 for C = 0,45. Corrections

have been applied t o C_, C , C , t o allow for the drag of the incidence wire and

the small exposed pai^t of the main feed pipe,

In order t o explore the vortex i n d e t a i l , a five tube pitch-yawmeter ( f i g , 5 )

was used T«4iich enabled traverses to be made in a spanwise plane. The head was

o similar to the five tube probe described in ref,7. The apex angle was 70 and the outside diajneter 0.125in,

5» Discussion of results

As a preliminary to the main wind tunnel programme, a sho* t series of tests wajs made, both with and without blowing at all edges, in order to provide some basic information on both the rig and the effects of blowing.

Wind sjjeed was varied betv/een 50 ft/sec and 200 ft/sec with no appreciable Reynolds number effect, hence tests without blowing were made at 200 ft/sec. to

obtain the greatest accuracy but tests with blowing were made at lower wind speeds in order to achieve a reasonable range of C , A few tests were made at a wind speed of 50 ft/sec. giving a maximum C ^ value of 1.55 but balance readings were less accurate at this low speed and most of the tests with blowing were made at 100 ft/seo. when the maximum value of C „ was 0,59»

Comparative t e s t s were also made without blowing (a) with the s l e t open

and (b) sealed t o prevent ary airflow through the s l o t . In general, s e a l i n g

the s l o t had l i t t l e effect on the balance readings and the only appreciable

changes were observed at the highest incidence on the r o l l i n g and yawing moments,

5 . 1 . Lift

The jet sheet, originating at the leading edges and tips with blowing, rolls up to form the leading edge vortices in a manner similar to the rolling up of the free vortex sheets without blowing, although the pressure boundary condition is changed since the jet sheet can now support a pressure difference. This rolling up of the jet sheet has occurred in all tlie repoi-ted tests using slot blowing from streamwise tips and is a stable type of flow, A sharp change in lift and pitching moment occurs, however, in passing through zero incidence v/hen the vortex moves from one surface to the other. Without blowing, of course, the flow is attached at zero incidence even with sharp leading edges, but with blowing tlie flow is apparently stable only when the jet sheet has rolled up,

causing an abrupt change when the wing passes through zero incidence. The effect can be seen in fig, 10 vdiere the lift curve shows a discontinuity near a = 0.

In the present tests near zero incidence with blov/ing it was possible to find en incidence at which the jet sheets oscillated from one surface to the other, producing a sinusoidal lift. The frequency of the oscillation was about one cycle per second at 100 ft/seo. and is thought to be associated with downwash lag in the follovdng manner. Since the downy/ach angle e^ ~ — - , and with blov/ing C does not tend to zero as a tends to zero, it is possible to find an incidence c^^whioh is

less than the downwash angle for a given amount of blowing { \ ^ ®n)« •'•f ''-he wing

is at positive incidaice OL and the blowing is turned on, the jet sheet will tend to roll up over the top sxa:*face. As soon as the flow is established the aveaage induced downwash will be greater than the incidence and the wing effectively at negative incidence so the sheet will move to the other surface. Thus an

oscillatory motion is set up viiich would be undesirable in an aircraft, although it seems unlikeJIy that this condition would be approached in practice.

Lifti-inoidence curves without blovdng are shown in fig. 11. At low incidence

the camber effect of the edge droop (fig.6) increases the lift, but at higher incidence, with leading edge separations, the effect is small. The non-linear

6

-curves are typical of sharp-edged, highly-swept v/ings, although near zero incidence the curve is close to the H.T. Jones lift cxxrve slope for an aspect

+ o ratio 0,75 wing with attached flow, SMeslip angles of up to - 5 had no effect on lift. That sideslip has no appreciable effect on lift is fortuitous considering the non-linear nature of the problem. The changes in lift on the two halves of the wing are of opposite sign and, while not affecting overaü.1 lift, give large rolling moments (see Para,5.5).

The lift increment with blowing from all edges is plotted sigainst

momentum coefficient in figs, 12 ( ^ = O ) and 15 ( ^ = 50 ) , At ^ = 0 for small

values of C ^ and constant a the lift increased quite quickly, but above a C

of about 0,2 the rate of increase of lift was reduced. Por a given value of C ^ the lift increment increased with incidence. The way in -vrfiich lift changes with blowing and incidence can be explained in terms of the movement of the

leading edge vortices. Application of edge blowing moves the leading edge vortices upward and outward particular!ly at small incidence (figs. 45, 4 6 ) . The outward movement is accomplished in two stages. First3y, the entrainment effect of the jet reduces and finally eliminates the secondary separation, causing the main vortex to move towards the edge. Secondly, at higher blov/ing pressures, the jet sheet will penetrate fiorther into the mainstream before rolling up, and thus move the vortex core outboard again. At a sufficiently large value of C^ the vortex core will move off the wing. This upward and outward movement of the vortex core with increasing C^ means that the effect of the strong LE vortex is reduced somewhat since its height is greater than without blowing, particularly at

small incidence, and the outward movement reduces the area of wing over which these low pressures are felt; hence the rate of increase of lift decreases with increasing

Cj, at constant a . Again as incidence, or mare strictly •^ , increases the vortex

core moves inboard without blowing so that the G^ at -wiiich — s — begins to level off

'^ AC L

is increased. Thus the condition for maximum relative benefit (—p;—) will be smsü-l /^

Oy to limit the spanwise movement of the vortex and large t; so that change in vortex height due to blowing is small.

With 0= 50 , P i g 13 shews t h a t w i t h l e a d i n g edge droop t h e e f f e c t of i n c i d e n c e i s roughly t h e same. The l i f t v a l u e s include t h e d i r e c t j e t l i f t (C ;i s i n 0COS a ) t h i s being approximately t h e d i f f e r e n c e between t h e ^ = 0 and

^ = 5 0 r e s u l t s a t the l a r g e r v a l u e s of C ; i .

P r e s s u r e p l o t t i n g r e s u l t s ( P a r a . 5 » 7 ) w i t h blowing from a l l edges showed t h a t t h e m a j o r i t y of t h e l i f t increment due t o blowing came from t h e forward swept p a r t of the wing and t h a t Up and t r a i l i n g edge blov/ing was r e l a t i v e l y i n e f f e c t i v e . A l l f u r t h e r t e s t s were t h e n donfined t o l e a d i n g edge blovdng only w i t h t h e undrooped ( 0= o) e d g e s . Simple t h e o r e t i c a l c o n s i d e r a t i o n s showed t h a t i n order t o m a i n t a i n c o n i c a l flow w i t h blowing i t was necessary t o i n c r e a s e t h e momentum e j e c t e d l i n e a r l y along t h e swept edge from a zero v a l u e a t t h e apex, This was achieved using t h e grooved perspex s t r i p s ( f i g , 4 ) . V/ith thesy s t r i p s i t was a l s o p o s s i b l e t o t e s t t h e e f f e c t of sweeping the j e t s h e e t r e l a t i v e t o t h e edge ( s e e P a r a , 2 ) . P i g . l 4 s h o / s t h e l i f t increment due t o blov/ing w i t h t a p e r e d l e a d i n g edge blowing o n l y , t h e j e t emerging normal t o t h e l e a d i n g edge ( 0 = O). Comparison w i t h P i g . 1 2 shows t h a t t h e l i f t increment a t a g i v e n value of C y i s i n c r e a s e d , p a r t i c u l a r l y a t small G^ v a l u e s . At l a r g e C/i v a l u e s (C ^ " 0 . 5 ) , t h e g a i n s are s m a l l s i n c e i n b o t h c a s e a G^ i s based on t h e t o t a l momentum e j e c t e d and t h e v o r t e x movement (and hence l i f t increment) depends on t h e l o c a l On v a l u e which i s l a r g e r f o r l e a d i n g edge blowing only ( i . e . not bloving fro{j t i p s and t r a i l i n g edge) f o r a given o v e r a l l C yj .

I n an attempt t o a n a l y s e t h e r e s u l t s o b t a i n e d w i t h 9 = 0 , t h e r e a i I t s of f i g . 14 were p l o t t e d i n non-dimensional form i n f i g . 1 4 a , g i v i n g an approximately l i n e a r r e l a t i o n

between —x— and —— f o r f i x e d •^. P i g . 14b sho: s t h a t a r e l a t i o n of t h e form ^ L _ C^

— o ~ /-^ = 0»20 + 1,24 R ; ) ^ c o u l d e x i s t . _{j^2 K K}

1 1

Thus AC = 0.20 C'^ K + 1.24(ft .C . K ) ^ . Adding t h i s t o Mangier and S m i t h ' s r e l a t i o n f o r t h e l i f t we h a v e ;

± ± 2

8

-This r e l a t i o n i s unsatisfactory i n t h a t the e f f e c t s of blowing diminish

with aspect r a t i o vdiile the effects of the normal secondary separation do not.

Using a r e l a t i o n of the formj

- ^ / ^ = 0.20 + 1.57(|) t f i g . l 4 b )

and the Brown and Michael expression for the l i f t we havet

± _1_ 2. i_

C^ = (277U+ 0.20 C2)K + (5,0a + 1,37 Cf) a^ k^

Here, at l e a a t , the effects of blowing and tlie unblovm non-linear effects are

comiDarable, A more exact r e l a t i o n must await the r - s u l t s of a t h e o r e t i c a l

investigation now i n progress,

Pigs. 15-20 show the effect of sweeping the j e t sheet, for small C^ values,

e = 0 ° , 20°, 50°, 60°, 70°, 80°, The increase i n l i f t with C^ tends t o be l i n e a r

up t o C ' s of at l e a s t 0.03 for a >0. The beneficial effect of increasing

a ( o r —) i s c l e a r l y v i s i b l e . Apart from 0= 6O , where the blowing d i s t r i b u t i o n was

poor, the effect of increasing 6 i s f a i r l y small up t o 6= 70 . At 6 = 80 , hovrever,

with f u l l leading edge blowing (fig.20) there i s a marked f a l l - o f f i n l i f t at p. given

0 , This effect was investigated and a t t r i b u t e d t o the j e t sheet clinging t o the

leading edge by means of the Coanda effect instead of emergini^ at the angle of the

grooves as was the c ase up t o 0 = 7 0 , This effect delayed the r o l l i n g - u p of the

j e t sheets and hence reduced the strength of the leading edge v o r t i c e s ,

In order t o mitigate t h i s effect, a further set of t e s t s was made with 0= 80

but with blowing from the f i r s t 6.5 i n . of the leading edge slot only, (Prom 5 i n . t o

9,5 i n . from the apex). The G^ range vms very much reduced owing t o the smaller

s l o t area (smaller mass flow) but f i g , 2 1 shows a considerable improvement over f i g . 2 0

for the higher incidence,

The lift augmentation ratio, „ , is plotted agiinst jet sweep single 6 in fig,22 for a 0„ value of 0,01, corresponding approximtitely to the cruise value. This shows clearly that the effect of iixjreasing 6 is small up to

o

70 but is large at greater angles when the jet sheet clings inltial3y to the leading edge. The "starred" values of 0 = 80 were obtained with blowing close to the apex.

5.2 Qra^

Ik"ag i s p l o t t e d a g a i n s t l i f t i n f i g . 2 5 f o r 0 = 0 and 30 and C v a l u e s of 0 and 0 . 1 7 8 , blovidng from a l l e d g e s . Blowing i n c r e a s e s t h e drag a t s m a l l i n c i d e n c e due t o a s m a l l forward component of t h r u s t , b u t drag i s l e s s vjith blowing a t h i g h e r v a l u e s of l i f t due t o a r e d u c t i o n i n t h e l i f t - d e p e n d e n t d r a g , The e f f e c t on drag of s e a l i n g t h e s l o t i n the unblown c a s e , and of s i d e s l i p

+ o

between - 5 b o t h w i t h and w i t h o u t blowdng, was n e g l i g i b l e .

With blovri-ng, a t c o n s t a n t i n c i d e n c e , t h e d r a g w i l l c o n s i s t of f o u r p a r t s i C^ = C-rv I-. - 0 c o s ( l l 0 - 0 ) c o s a + A C^ t a n a + A C^

D. - . D no blow n ^ ' K, D

blowing ^

The f i r s t term i s t h e drag w i t h o u t b l o w i n g , t h e second i s t h e d i r e c t t h r u s t component due t o l e a d i n g edge blowing, t h e t h i r d i s t h e iixsrement of induced drag due t o t h e i n c r e a s e d l i f t , ard. t h e l a s t term i s a small t h r u s t due t o t h e f a c t t h a t most of t h e i n c r e a s e d l i f t due t o blowing a c t s on t h e forward p a r t of t h e wing c a u s i n g i n c r e a s e d s u c t i o n s on t h e forward f a c i n g s u r f a c e s . The f i r s t term c a n be o b t a i n e d from f i g . 25 and t h e second and t h i r d terms are e a s i l y c a l c u l a t e d , so i t i s of i n t e r e s t t o p l o t t h e remaining term. I t i s , e f f e c t i v e l y , a r e d u c t i o n i n induced d r a g and oan be c o n v e n i e n t l y p l o t t e d i n t h e form - -—- V s A C .

°L

h

Values for a = 5°, 10°, 15°, 20° with 0 = 0°, 20°, 50°, 60°, 76°, 80° are plotted in figs. 24-27. The values of - ^ 0 are small and there is a good deal of scatter

VLIEGTU;

Bir' 10 Bir'

-but the results for 6 = 80 are probably less reliable than the others.

The important fact ia, that the values of Ac are always negative i.e. with leading edge blowing there is a small increment of thrust over and above the direct jet thrust measured wind off. Vifith an aircrsift designed to utilise its full thruat as leading edge blowing, this would bring a further increase in I/D, quite apart from obvious gains due to lower induced drag at a given lift coefficient,

3.3» Cross-vd.nd force

Cross-wind force is plotted against the sideslip angle, /3, in figs. 28 and 29-Por the range of sideslip tested ( ^= - 5 , 0 +5 ) the variation is apparently linear, although no tests were made at intermediate points. The results of

+ 0

r e f . 8 suggest t h a t the v a r i a t i o n w i l l be l i n e a r between p = - 5 for the

iixïidence range t e s t e d . Por ^ = 0 the blowing r e s u l t s d i f f e r only s l i g h t l y

from the unblown r e s u l t s , due t o asymmetries i n the blowing giving a small force

a t zero s i d e s l i p . Similar observations may be made for 0= 50 . I t i s

concluded that the effect of symmetrical edge j e t s on cross-wind force i s anxall,

5 . 4 . Pitching moments

Pitching moments about half root chord are p l o t t e d against l i f t i n

o

f i g s , 30 - 32, Again, sealing the s l o t vidthout blovri.ng and sideslipping up to 5

had very l i t t l e e f f e c t , V/ithout blowing, ^ = 0 , the pitching moment varied l i n e a r l y

with l i f t ,

The effect of blowing from a l l edges and of drooping the edges can be seen i n

f i g , 5 0 . With ^ = 0 , the effect of blovri.ng on pitching moments was small kt high

incidence but larger at small incidence, Thexe was also a discontinuity near a = 0

due t o the vortex moving from one surface t o the other corresponding t o tlie

discontinuity i n l i f t (see P a r a , 3 l ) , Drooping the edges 50 moved both the

aerodynamic centre aM centre of pressure t o the r e a r , almost 0 , 1 o , Blowing with

o

^ = 30 aggravated t h i s effect ajnd introduced considerable n o n - l i n e a r i t i e s , although

these are probably due mainly t o non-uniform blowing.

KLg, 31 shows the effect of increasing 0^^ for the case d - ^ = 0 with

leading edge blowing only. In t h i s case the effect of blowing i s t o move the

aerodynamic centre and centre of pressure forward as the major effect of the

blowing i s f e l t on the foirward part of the wing*. Pig.32 shows changes due t o

increasing 0 with ^ = 0 and 0 „ = .014. Here the tendency i s for the centre

of pressure t o move back with increasing 0 since sweeping the j e t delays the

r o l l i n g up of the sheet somewhat,

3 . 5 . Rolling moments

In f i g s . 3} and 34 the r o l l i n g moment about the model c e n t r e l i n e i s

p l o t t e d against s i d e s l i p angle j3 , and the corresponding & values

(ö-sgr ) are p l o t t e d against l i f t i n f i g , 3 5 . The v a r i a t i o n of C^ appears t o

be l i n e a r vyithin the range -5 < 0 < + 5 and evidence from r e f . 8 supports t h i s ,

The r o l l i n g moments induced on a d e l t a wing i n s i d e s l i p are due mainly t o an

a^mmetric p a t t e r n of the leading edge v o r t i c e s ( f i g , 4 l ) . The vortex on the

advancing edge remains t i g h t l y r o l l e d but t h a t on the r e t r e a t i n g edge becomes

more diffuse and weaker. Thus for negative s i d e s l i p there i s a p o s i t i v e r o l l i n g

moment which increases with increasing incidence. The r e s u l t s for <f> = G^^ = 0

are shovm i n f i g . 5 5 j only a t the highest incidence was there an appreciable

oheinge due t o sealing the s l o t ,

Without blowing, the r o l l i n g moment due t o s i d e s l i p i s caused mainly by a

weaJcening of one leading edge vortex r e l a t i v e t o the other. With blowing, however,

the r o l l i n g up of the leading edge vortex sheets i s p a r t l y controlled by the edge

j e t s , vdiich ai'e comparatively unaffected by s i d e s l i p , and should r e s i s t t h i s

deformation of the flow p a t t e r n . This should tend t o reduce the magnitude of

0 ^ , and hence &, at a given l i f t . Prom the r e s u l t s for ^ = 0, however, i t

appears t h a t t h i s i s only t r u e at inoiderces above 15 (C_ > 6 , 5 ) .

• At C /x values corresponding t o the landing case ( ''* 0 , 4 ) , the forward movement

of the aerodynamic centre i s 2-3f<> of the root chord and s t a b i l i t y considerations

may l i m i t the usable C^ under these conditions,

12

-The effect of drooping the edges, fi = 30 , (fig.35) is to reduce •&

at a given lift. Without blowing, at least part of the reduction is due to the decrease in span due to the drooped edge and with blowing, the further reduction is the result of the direct jet reaction producing lift but no extra rolling moment.

5 . 6 , Yawinp: moments

Yavdng moments about h a l f mean chord are p l o t t e d a g a i n s t s i d e s l i p angle /3 for /ZJ = 0 and 50 i n f i g s , 56 and 3 7 . Again the v a r i a t i o n was l i n e a r i n t h e range of s i d e s l i p t e s t e d . Corresponding values of n are p l o t t e d a g a i n s t

l i f t i n f i g . 3 8 .

The asymmetric p r e s s u r e d i s t r i l i u t i o n due t o s i d e s l i p , (see P a r a . 5 . 5 ) , w i l l produce a yawing moment on a wing w i t h t h i c k n e s s due t o the p r e s s u r e d i f f e r e n t i a l on the side a r e a . As w i t h the r o l l i n g moments, s e a l i i i g the s l o t has an a p p r e c i a b l e e f f e c t a t t h e h i g h e s t i n c i d e n c e . Por fd = 0 , t h e e f f e c t of blowing from a l l edges i s s m a l l ,

Drooping t h e edges 50 i n c r e a s e s the side area by equal amounts f o r e and a f t of the h a l f chord p o s i t i o n , but s i n c e t h e chordwise loading i s c o n c e n t r a t e d towards t h e apex the e f f e c t of edge droop i s t o i n c r e a s e the magnitude of t h e yawing moments comijared with tl-ie /2S = 0 c a s e . There i s a change i n C /9= 0 due t o asymmetric blowing but n i s s c a r c e l y a f f e c t e d . There i s a r e d u c t i o n i n n a t a g i v e n l i f t w i t h blowing due t o t h e d i r e c t j e t l i f t ( s e e P a r a . 5 . 5 ) .

v

5 . 7 . Pressixre measurements and flow v i s u a l i s a t i o n

The spanvdse v a r i a t i o n of s t a t i c iaressure was measured a t four chordwise s t a t i o n s , 0.33c , 0.49c , 0.63c and 0.87c . Without blowincr, t h e r a n g e of

o o o o

incidence was a= 2 , 5 , 10 , 15 , 20 , 25 and w i t h blowin^T, a= 5 , 10 , 15 . The r e s u l t s are shown i n f i g s . 39-43. The p r e s s u r e s measured on t h e s t a r b o a r d wing are on t h e l e f t hand s i d e of t h e f i g u r e s , i . e . the wing i s viewed from the stream d i r e c t i o n ,

Without blowing ( f i g s . 5 9 - 4 l ) the p r e s s u r e d i s t r i b u t i o r s a r e t y p i c a l of wings with s h a r p , h i g h l y swept leading edges. The s u c t i o n peak which occurs beneath t h e v o r t e x core can be c l e a r l y seen, p a r t i c u l a r l y a t t h e more forward s t a t i o n s , and small s u c t i o n peaks are evident at angles of incidence a s low as two degrees ( f i g . 5 9 ) . The s l i g h t b l u n t n e s s of the leading edge, due t o t h e s l o t , does not appear t o prevent the flow s e p a r a t i n g vidthout blo^ying, even i t very low i n c i d e n c e . At higher i n c i d e n c e , t h e extent of the secondary s e p a r a t i o n i s i n d i c a t e d by a r e g i o n of roughly c o n s t a n t p r e s s u r e outboard of t h e main v o r t e x c o r e . Reasonable agreement i s obtained vdth the r e s u l t s of r e f . 9 , i n vriiich surface s t a t i c p r e s s u r e s were measured on a 70 t r u e d e l t a vdng.

P i g . 4 0 shows t h e chordvdse v a r i a t i o n of s t a t i c p r e s s u r e , without blov/ing, a t constant i n c i d e n c e , a = 55 . The tlovi remains approximately c o n i c a l i n

form back t o as l e a s t 0.65c , i . e . t o t h e leading e d g e - t i p j u n c t i o n . At 0.87c , however, t h e r e has c l e a r l y been a c o n s i d e r a b l e r e d u c t i o n i n l i f t due t o t h e e f f e c t of t h e t r a i l i n g edge. I t i s not thought t h a t v o r t e x breakdown i s occurring a t t h i s incidence a t t h e r e a r s t a t i o n . Ref. 10 shows t h a t v o r t e x breakdov/n occurs behind the t r a i l i n g edge of a 70 swept p l a t e a t a = 25 , although t h e e f f e c t of cro^yping a d e l t a wing has not been i n v e s t i g a t e d . The s u c t i o n peak r i s e s s t e a d i l y vri.th incidence a t — = 0.87 and does not shav the r e v e r s a l of t h i s t r e n d a t high _c

o

incidence which i n d i c a t e s v o r t e x breakdovm i n r e f . 10. Pig.2)1 shews the e f f e c t on t h e p r e s s u r e d i s t r i b u t i o n of f i v e degrees of s i d e s l i p . '^he basic flow p a t t e r n appears t o be unchanged, but t h e v o r t e x on the advancing edge i s s t r o n g e r and has moved s l i g h t üy i n b o a r d , while t h e v o r t e x on t h e r e t r e a t i n g edge is weaker but does not appear t o have moved. These r e s u l t s ai"e g e n e r a l l y i n agreement with r e f . 1 1 ; a t higher angles of s i d e s l i p the v o r t e x on the r e t r e a t i n g edje nioves towai'ds t h e leading edge and the secondary s e p a r a t i o n i s no longer v i s i b l e .

•^'he incx-ease i n v o r t e x s t r e n g t h due t o edge blovdng can be seen i n f i g , 4 2 . I n crder t o maintain c o n i c a l flow w i t h blowing, the inementum e j e c t e d should i n c r e a s e l i n e a r l y from a zero value a t the apex and t h i s ia almost achieved using t h e

grooved perspex s t r i p s (C „ = O.O96,j0= 0 ) , V/ith an open s l o t , blovdng from a l l edges, t h e momentum d i s t r i b u t i o n was more . n e a r l y constant i . e . t o o much a i r was e j e c t e d near the apex, and r e s u l t e d i n a very non-uniform chordwise p r e s s u r e

14

-d i s t r i b u t i o n . Even w i t h blowing, the loa-d at t h e rearmost p r e s s u r e p l o t t i n g s t a t i o n , 0.87c , was s t i l l small (see P a r a , 3 . 1 ) . With blowing from t h e leading edges only t h e r e wdll be a sudden r e d u c t i o n i n the s t r e n g t h of the leading edge v o r t e x sheet a t the leading edge wing t i p j u n c t i o n . This would tend t o r e d u c e t h e incidence a t which v o r t e x breakdown occiirs but i t i s not thought t o have occurred i n the p r e s e n t t e s t s . Further information i s d e s i r a b l e , however, and a more d e t a i l e d i n v e s t i g a t i o n i n t o t h e e f f e c t s of blowing on v o r t e x breakdown i s t o be made s h o r t l y .

P i g , 4 3 compares the r e s u l t s obtained with and without leading edge

blowing only w i t h t h e t h e o r i e s of Brovm and Michael and Mangier and Smith . The s i m i l a r i t y between t h e p r e s s u r e d i s t r i b u t i o n s for t h e experimental r e s u l t s w i t h blowing and t h e t h e o r e t i c a l v a l u e s vriLthout blowing i s t o some extent f o r t u i t o u s and i s dependent on the value of C , but i t i s i n t e r e s t i n g t o note t h a t t h e e x i s t e n c e of t h e secondary s e p a r a t i o n i s mainly r e s p o n s i b l e f o r t h e

( 1 2 ) . discrepancy between t h e o r y and experiment f o r t h e spanwise p o s i t i o n of the v e r t e x The s e p a r a t i o n which occurs without blowing i s suppressed g r a d u a l l y by t h e

entrainment e f f e c t of the edge j e t and can be e n t i r e l y suppressed vidth q u i t e small values of C (about 0,05 a t a= 2 0 ° ) ( s e e a l s o r e f . 1 2 ) ,

The movement of t h e v o r t e x core w i t h incidence (O = O) i s shovm i n f i g . 4 4 (9 13)

and t h e p r e s e n t r e s u l t s are compared w i t h other expex-imental values ' and with t h e t h e o r i e s of Brovm and Michael and Mangier and Smith . 'The p o s i t i o n s were estimated using a t u f t g r i d and using t h e f i v e - t u b e pitch-yavraeter. The experimental r e s u l t s ax'e i n r e a s o n a b l e agreement except for t h e spanvdse p o s i t i o n s of t h e v o r t e x given by r e f . 13 which appear t o be i n e r r o r . The t h e o r y of Brown and Michael p r e d i c t s

(2 3)

t h e h e i g h t of t h e v o r t e x core reasonably w e l l but b o t h t h e o r i e s , ' , a r e s e r i o u s l y i n e r r o r w i t h r e g a r d t o the spanwise p o s i t i o n . Some r e a s o n s for t h e discrepancy are d i s c u s s e d i n r e f . 1 2 . The e f f e c t of leading edge blowing only i s shown i n f i g u r e s 45 and 46 where t h e r e s u l t s with blowing, (C„ ~ O.O5 and 0 . 1 6 = 0,50 ) a r e obnpared v d t h t h e msan v a l u e s obtained vri.thout blowing. At zero incidence w i t h blowing t h e r e is q u i t e a powerful v o r t e x , but without blowing t h e r e i s no v o r t e x on t h i s symmetrical wing. Thus f o r a given C^ the e f f e c t of blowing on t h e vortex p o s i t i o n i s g r e a t e s t a t low incidence and decreases a t higher incidence where a

v o r t e x would e x i s t vdthout blov/ing, A change i n 0 from 0 t o 50 a t c o n s t a n t CM does not change t h e v o r t e x p o s i t i o n g r e a t l y as would be expected from t h e l i f t r e s u l t s ( f i g , 2 2 ) . At much l a r g e r v a l u e s of C^ ( Q = 0 , C^ = O.56) t h e v o r t e x core i s w e l l aboiie the vdng and at incidences below 15 i s outboard of t h e l e a d i n g edge.

A s e r i e s of t r a v e r s e s i n a spanwise plane were made with the f i r e t u b e p i t c h -yavraiet«r ( f i g , 3 ) a t one i n c i d e n c e , vdth and without blowing, t o show t h e changes i n flow p a t t e r n . Tlie apparatus was a l s o used t o l o c a t e t h e p o s i t i o n of t h e v o r t e x

core as a check on t h e t u f t o b s e r v a t i o n s . R e s u l t s for a = 10 , On = 0 and 0.048 are H-Po

conipared i n f i g s . 47 and 48, Values of are p l o t t e d a g a i n s t non-dimensional ^o

semi-span ^ ; t h i s g i v e s a good i n d i c a t i o n of t h e l o s s e s occurring and shows t h e s t r u c t u r e w e l l . The extent of the leading edge v o r t e x i s defined by t h e v a l u e 1.0 of t h e v a r i a b l e ,

Without blowing ( f i g , 4 7 ) t h e v o r t e x core and t h e r e g i o n of secondary s e p a r a t i o n

H-PQ

can be c l e a r l y seen. The minimum value of • ^ recorded was - 0 , 7 9 . vriiioh

q ' indicates a high axial velocity although the actual value could not be calculated

owing to the difficulty of measuring pressure in the core with the five-tube probe. With blowing, C^^ = 0,048, (fig,48) a much larger region is affected,

The height of the vortex core is increased and it is moved outboard. The secondary separation has been eliminated and it is possible to trace the effect of the jet for

** o

almost a full turn of the sheet, Mirdmum value of — — is now -2,5, indicating a

%

considerable increase in velocity over the unblown case,

3.8, Some practical applications of leading edge blowing

While it is possible to tolerate, at relatively lov/ speeds, an aircraft composed of g. number of separate items (fuselage, vdngs, tail, engines etc.) at much higher speeds the case for a fully integrated aircraft is unanswerable. In the early stages, the application of a new idea is not easy to foresee and is usually fraught with engineering difficulties, e.g the jet flap, but it was felt worthwhile to

include some thoughts on the possible uses of leading edge blowing if only as a basis for further disoussion,

16

-The use of leading edge blovdn^^- presupposes a very highly swept vdng and t h e experimental r e s u l t s show t h a t c o n s i d e r a b l e g a i n s i n i^ressure l i f t a r e achieved without any d i r e c t t h r u s t l o s s , at l e a s t a t low speeds. Very l i t t l e information on t h e effect of Mach number i s a v a i l a b l e , although some unpublished work c a r r i e d out a t t h e College of Aeronautics vdth s l o t blovdng from t h e shoulder of a h a l f cone ahovted t h a t even a t a Mach number of 2 , 0 , a much l a r g e r v o r t e x was produced liy t h e blowing and a p p r e c i a b l e i n c r e a s e s i n l i f t o b t a i n e d . Owing t o t h e small s c a l e of the experiment, however, i t was not p o s s i b l e t o measure the

i n c r e a s e of l i f t a c c u r a t e l y , although 'J^ vms of t h e order of one a t small C v a l u e s . Provided the vdng and v o r t e x sheet a r e vrell i n s i d e the mach c o n e , the

e f f e c t s of mach number w i l l probably not be excessive a t small a although since the p r e s s u r e on t h e t o p surface cannot f a l l below a b s o l u t e vacuum t h e r e i s a d e f i n i t e

l i m i t t o t h e i n c r e a s e i n n o n - l i n e a r l i f t t h a t can be achieved. Thus i t seems l i k e l y t h a t a p p r e c i a b l e i n c r e a s e s i n l i f t vd.ll s t i l l be achieved even for Mach nxunbers around tvro and s i n c e l i f t t o d r a g v a l u e s are g e n e r a l l y low on h i g h l y swept vdngs, even small improvements are w e l l worthwhile,

I n o r d e r t o feed a i r t o t h e leading edge the choice seems t o l i e between c o n v e n t i o n a l l y mounted j e t engines ( i . e , a t t h e r e a r ) vdth a l a r g e amount of ducting and i t s a t t e n d a n t l o s s e s and space problems, and engines mounted much c l o s e r t o t h e l e a d i n g edge vd.th a f i s h t a i l nozzle forming t h e leading edge s l o t , The l a t t e r can be fux'ther sub-divided i n t o a l a r g e number of sm.all j e t engines spaced a t i n t e r v a l s dovim t h e leading edge or perhaps four l a r g e j e t s mounted i n t h e fuselage a t t h e nose (two per s i d e ) and exhausting near t h e vdng apex,

P r o p u l s i o n would s t i l l be obtained from d i r e c t j e t t h r u s t and t h i s would mean t h a t t h e j e t exhaust could not be i n c l i n e d more t h a n say 20 from t h e a i r c r a f t c e n t r e l i n e i n t h e plane of the vdng except during l a n d i n g , Ovd.ng t o t h e tendency f o r t h e j e t t o c l i n g t o the leading edge vihen i t i s highly svrept (see P a r a , 5 , l ) , i t would be p r e f e r a b l e t o c o n c e n t r a t e t h e blovdng i n t h e r e g i o n of t h e vri.ng apex, Thus from an aerodynamic viewpoint, fuselage mounted j e t s exhausting c l o s e to t h e wing apex a r e p r e f e r r e d .

Moving t h e engines from t h e v i c i n i t y of the t r a i l i n g edge t o t h e apex involves a major r e d i s t r i b u t i o n of t h e weight and, as a r e s u l t , t h e passenger load would have t o be moved a f t . A t e n t a t i v e layout suggests t h a t an a p p r e c i a b l e p a r t of t h e fuselage w i l l p r o j e c t a f t of t h e t r a i l i n g edge i n s t e a d of forward of the apex a^ vri.th t h e p r e s e n t "Concord" a i r c r a f t . J e t exhausts emanating from the vdng apex may cause some h e a t i n g problems on t h e v/ing upper s u r f a c e , but the f a c t t h a t t h e y w i l l r o l l up i n t o the leading edge vortex core and hence do not a c t u a l l y touch t h e vdng, combined with t h e temperature drop i n t h e v o r t e x core caused by the r o t a t i n g flow, suggests t h a t the h e a t i n g problem may not be s e v e r e . P r e s s u r e f l u c t u a t i o n s on t h e wing s u r f a c e under t h e v o r t e x core would a l s c i n c r e a s e somewhat. Nose mO'onted j e t s w i l l undoubtedly i n c r e a s e t h e n o i s e l e v e l i n s i d e t h e a i r c r a f t u n l e s s a d d i t i o n a l lagging i s provided, but noise i n t h e f a r f i e l d v d l l probably be l e s s owing t o t h e s h i e l d i n g a f f e c t of t h e a i r c r a f t i t s e l f .

I n crder t o g«t some i d e a of p o s s i b l e improvements i n performance vrtiich would be o b t a i n e d v d t h l e a d i n g edge blovdng c l o s e t o the apex, a simple comparison has been

made between two a i r c r a f t of the same s i z e and weight and roughly r e p r e s e n t a t i v e of t h e supeEsonio a i r l i n e r ,

The c o n v e n t i o n a l a i r c r a f t considered has a g r o s s weight of 500,000 l b s , and a o

maximum t o t a l t h r u s t of 100,000 l b s . I t i s assumed t o c r u i s e a t o= 4 and have a l i f t / d r a g r a t i o of e i g h t (C, = 0 , 1 , C_ = 0,0125), The unconventional a i r c r a f t i s the same s i z e but has i t s engines mounted i n t h e nose, exhausting i n t h e form of a sheet near t h e vdng apex i n t h e plane of t h e vdng a t an angle of 15 t o t h e a i r c r a f t c e n t r e

l i n e , ^Q Lg

Par c r u i s i n g , the assuaiption i s made t h a t •••-•" ' = 1 , 0 and henee f o r t h e same l i f t (CT = 0 , 1 ) O i s reduced vd.th a consequent r e d u c t i o n i n l i f t dependent d r a g . ( i t i s f u r t h e r assumed t h a t blowing does not i n c r e a s e t h e vreive d r a g ) . I t should be noted here t h a t t h e l i f t incx-ement due t o blovdng depends on t h e momentum e j e c t e d over the v/ing i . e . depends on t h e g r o s s and not t h e net t h r u s t , (A value of g r o s s t h r u s t = 1,5 X n e t t h r u s t has been t a k e n h e r e ) . Under t h e s e c o n d i t i o n s i t can be shown t h a t a thrxist of only 90go of t h e t h r u s t of the conventional a i r c r a f t i s

18

-sufficient to fly the aircraft at o = 3*3 and the lift/drag ratio is increased

to 8,97, an increase of 12?^, even without a further allowance for the decrease in fuel and engine weight. More detailed calculations were not made owing to the

vincertainty of the basic assumption -^ = 1.0, but this is probably not greatly

U

i n e r r o r and more d e t e d l e d c a l c u l a t i o n s vri.th allowances for decreases i n f u e l and A C i ^

engine vreight based on an a c c u r a t e value f o r -r;— a r e escpected t o show i n c r e a s e s i n l i f t / d r a i g r a t i o s of t h e same o r d e r .

At t a k e off, -7;— v d l l be l a r g e r t h a n for c r u i s e , owing t o t h e higher

i n c i d e n c e , a = 15 . The p l a i n wing i s assumed t o give a C_ of 0.5 a t t h i s incidence without blovdng, r i s i n g t o C^. = O.65 vd.th G^ = 0.195» Thus t h e t a k e - o f f speed i s reduced by about 12yo» On landing (weight = 170,000 l b s . ) , C^ r e a c h e s 0,79 a t

On = 0,417 g i v i n g a r e d u c t i o n i n landing speed of 20JiJ, assuming t h a t some form of

svd.velling nozzle v d l l enable t h e engines t o be r u n at f u l l t h r u s t , t h e net t h r u s t being reduced t o an a p p r o p r i a t e v a l u e by d e f l e c t i n g t h e j e t s away from the a i r c r a f t , *

Thus I t seems t h a t t h e a p p l i c a t i o n of leading edge blowing t o a supersonic g i l r l i n e r of c o n v e n t i o n a l s i z e and weight would show u s e f u l g a i n s i n c r u i s e and t ak e -o f f performance and very s u b s t a n t i a l g a i n s i n landing performance, viAiich i s i n maiy c a s e s t h e l i m i t i n g f a c t o r i n p r e s e n t day d e s i g n s . No attempt should be made t o minimise the d i f f i c u l t i e s involved i n the use of leading edge blovd.ng, e s p e c i a l l y on a l a r g e s c a l e , but none

would appear t o be i n s u p e r a b l e , e s p e c i a l l y when compared v d t h t h e d i f f i c u l t i e s involved i n c o n v e n t i o n a l d e s i g n s ,

4. Conclusions

lent speed vdnd tunnel tests have been made on a 70 cropped delta vri.ng

to investigate the effects of edge blowing.

The main conclusions are as follows)

1. Edge blovdLng, in particular leading edge blowing, increases the lift by increasing the size and strength of the leading edge vortices at a given

incidence. The increase in lift is due mainly to an increase in the non-linear c ontribu t ion,

2. The edge jet sheets always roll up to form leading edge vortices giving steady flow patterns vdth the vd.ng at incidence. At a sufficiently small incidence, however, the jet sheets oscillate from one surface to the other due to a dovmwash lag effect. This is unlikely to be a practical limitation since the incidences are of the order of 0.25 .

3. The effect of bloY«lng from the streamvdse tips and trailing edge was small, and best results were obtained ;vith leading edge blowing only. The use of small grooved perspex strips enabled both the distribution and direction of the blovd.ng to be controlled and except for very highly swept jets there vra.s very little reduction in lift augmentation vdth increasing 9,

4. Maximum values of the lift augmentation, - g — , were obtained for small values of C (<0,05) and large values of inciderjce. The maximum value achieved was 2.8 for «• = 20 and Ö = 0 , 20 vdth blowing in the plane of the wing i.e. no direct jet lift.

5. There was no appreciable Reynolds number effect in the range 0.8 x 10 to 3.2 X 10^.

/• + o

V / l ' ' - - ^

20

-7 . The e f f e c t of edge blovdng i n t h e plane of the wing on t h e l a t e r a l d e r i v a t i v e s t and n a t a givien l i f t was s m a l l . The e f f e o t on p i t c h i n g moment depended on O/i auid 6, but f o r p r a c t i c a l values of C^ i t was not e x c e s s i v e ,

8 , TXxe t o t h e higher s u c t i o n f a r c e s a c t i n g mainly on t h e forward f a c i n g sxxrfaces, t h e v a l u e s of t h r u s t were s l i g h t l y higher than vrauld have been expected from c o n s i d e r a t i o n of t h e d i r e c t j e t t h r u s t component,

9 . If l e a d i n g edge blovd.ng were used t o reduce landing speeds using a i r b l e d from t h e j e t engine compressors, a b l e e d of 10% of the mass flow vrauld reduce landing speeds by about t e n k n o t s , i . e . roughly t h e r e d u c t i o n obtained by t h e use of blovdng boundary l a y e r c o n t r o l vdth t r a i l i n g edge f l a p s on more c o n v e n t i o n a l

planforras. Ducting would be a problem because of the l e n g t h involved ( a p p r o x , 2 0 0 f t , ) and t h e r e l a t i v e l y l a r g e area (about 5?" of wing c r o s s s e c t i o n a l a r e a ) .

10, I f t h e t o t a l j e t t h r u s t vrere a v a i l a b l e i n t h e form of l e a d i n g edge s l o t blovdng, p r e f e r a b l y o l o s e t o t h e apex, r e d u c t i o n s i n t a k e - o f f and landing speeds of 12?'o and 20/i r e s p e c t i v e l y could be achieved, assuming t h a t t h e d i r e c t i o n of the j e t oould be s u i t a b l y o o n t r o l l e d . Although the e f f e c t s of Mach number a r e not knovTn a c c u r a t e l y , i n c r e a s e s i n t h e l i f t / d r a g r a t i o of about lO^o would appear f e a s i b l e ,

The vrork d e s c r i b e d i n t h i s r e p o r t formed p a r t of t h e A u t h o r ' s Fh,D, t h e s i s ,

The author i s i M e b t e d t o P r o f e s s o r C M . L i l l e y for h i s adviae and encouragement during t h e oourse of t h e work,

Thanks are a l s o due t o Mr, S,H, L i l l e y , Mr. D. Horn and Ur. I.UcRae f o r t h e i r h e l p w i t h the vri.nd t u n n e l programme, and t o Miss R. P u l l e r v^o performed the lengthy computations,

22

-5.

References

1 . J o n e s , R,T, R r o p e r t i e s of low aspect r a t i o p o i n t e d vri.ngs a t speeds below and above t h e speed of sound. N.A.C.A. Report No, 855, 1946.

2.

3.

Brovm, C.E, Michael, W,H. Mangier K.W. Smith J . H . B , Effect of leading-edge s e p a r a t i o n on t h e l i f t of a d e l t a vd.ng. J . A e . S . V o l . 2 1 , 1954, P.69O.

A theory of the flow past a slender delta vdng id.th leading edge separation.

Eroo. Roy, Soo. A. Vol.251, 1959, pp.200-217.

Smith, V.J. Sin^son, C J .

A preliminary investigation of the effect of a thin high velocity tip jet on a lov/ aspect ratio vdng.

A u s t r a l i a n Dept. of Supply. Note ARL.AI65. 1957,

5.

6.

7.

Ayers, R , P , Wilde, M.R. T r e b b l e , W,J,G. Bryer, D,W. Walshe, D,E, G-arner, H.C,

Unpublished College of Aeronautics T h e s i s ,

Low speed v»ing t u n n e l experiments vd.th blovd.ng from t h e l e a d i n g edge of a d e l t a vri.ng vri.th 70 svreepbaok, unpublished M.O.A. Report,

P r e s s u r e probes s e l e c t e d for mean-flow measurements, E x p l o r a t i o n of tiarbulent boundary l a y e r s .

A.R.C. Report 17,997. November 1955.

8. Peokhïun, D.H. low speed vdnd t u n n e l t e s t s on a s e r i e s of uncambered slender p o i n t e d vd.ngs vdth sharp edges.

9* Marsden, D , J , Simpson, R,W, R a i h b i r d , W,J,

The flow over d e l t a vdngs a t low speeds w i t h l e a d i n g edge s e p a r a t i o n , O t l l e g e of Aeronautics Report 114, 1957» 10, lambourne, N,C. Bryer, D,W, The b u r s t i n g of leadingedge v o r t i c e s -some o b s e r v a t i o n s and d i s o u s s i o n of t h e phenomenon. AJR.C. R, and M, 3282, 1962,

1 1 . Harvey, J,K, Some measurements on a yawed slender d e l t a vdng v d t h leading edge s e p a r a t i o n .

A J i . C . R. and M. 516O, 1958,

12. Alexander A , J , Experiments on a d e l t a wing using l e a d i n g edge blovd.ng t o remove t h e secondary s e p a r a t i o n s . College of Aeronautics Report I 6 I , I965,

13» Lambourne, N,C. Bryer, D,W,

Some measxirements of t h e p o s i t i o n s of t h e v o r t i c e s f o r sharp edged d e l t a a n i svirept back wings.

FIG. 1 MODEL MOUNTED IN WIND TUNNEL

FLCXIM-E HOSE 0 - 2 THeuST IN Ibft/inch C, r i o i s 1 SYMBOL O X 0ISTR16ÜTION IDEAL USN& PERSPEX STRIPS OPEN SLOT -^ ^ O . ^ ^ ï

A

X

FIG.e BLOWING MOMENTUM DISTRIBUTION ALONG LEADING EDGE. PIC.S. MODEL MOUNTING.

• • J O *

FIG 6 3 0 " D R O O K D EOCE . FIG.9. FORCE AND MOMENT AXES SYSTEM.

- "

c» »o 10

FIG. 10. JET SHEET INSTABILITY NEAR ZERO INCIDENCE

BLOWING FROM ALL EDGES. FIG. 7 . DEFINITION OF • t 4

° 5 10 -t* 15 lO 25

FIG.II. VARIATION OF LIFT WITH INCIDENCE. C)i» O.

0 5 I 1 1 1 1

1-FIG. 12. VARIATION OF LIFT INCREMENT DUE TO BLOWING WITH MOMENTUM COEFFICIENT. BLOWING FROM ALL EDGES. • = 0 '

J^ / ^ ^ / ^ / /

J

f\

y

STMBOL e • X 1 + .( s' lO* 15' JO' 25'

FIG. 13. VARIATION OF LIFT INCREMENT DUE TO BLOWING WITH MOMENTUM COEFFICIENT BLOWING FROM ALL EDGES ^ s S O ' .

FKS. M. VARIATION OF LIFT INCREMENT DUE TO BLOWING

WITH MOMENTUM COEFFICIENT L E . BLOWING ONLY. ^ - • > 0 .

FIG.14b. VARIATION OF - ^ WITH ( • ^ ) ^ O ^ C ^ ) '

Cp'2

K

AC. _{^ S a C u '}

FIG. 16. VARIATION OF g F T INCREMENT DUE TO BLOWING WITH MOMENTUM COEFFICIENT. L E BLOWING ONLY. • -O^ « •20°.

0 0 2 C(i 0 0 3

FIG IB. VARIATION OF LIFT INCREMENT DUE TO BIOWING WITH MOMENTUM COEFFICIENT L E . BLOWING ONLY.

FIG. 19. VARIATION OF LIFT INCREMENT DUE TO BLOWING WITH MOMENTUM COEFICIENT L.E. BLOWING ONLY. * » 0 ' , 9i70">.

O 06

A.ISf 20°

0 0 4

FIG.21. VARIATION OF LIFT INCREMENT DUE TO BLOWING WITH MOMENTUM COEFICIENT, BLOWING FROM APEX. < ^ . 0 , 9= 8 0 °

O-OJ C(i. 0 0 4

FIG.20. VARIATION OF LIFT INCREMENT DUE TO BLOWING WITH

MOMENTUM COEFFICIENT L E . BLOWING ONLY. < > - 0 " , 9 - 8 0 ° FIG. 2 2 . VARIATION OF LIFT AUGMENTATION

f L 5 ^ WITH JET SWEEP ANCLE 9 . ^ • O , C^sOOI.

FIG.25. V A R I A T I O N OF -—^ WITH AC, „ L E . BLOWING ONLY. C(_ LB

PIG. 21. VIMIATION OF DRAG WITH LIFT. BLOWING FROM ALL

EDGES. 0 O 2 6 0 O 2 O» + A 0 0 r • • 1 -o a/"^ A 4 . ^ O C

y

y.

k )6 O O ? ^ . e 0 1 0 SYMBOL B T + X O 1 « o« 2 0 ° 5 0 ° 6 0 ° 7 0 ° 8 0 ° 8 0 ° 2 •• 0 . -ACg -ACr

FIG.26. VARIATION OF - ^ WITH A C L B L . E . BLOWING ONLY.

- ^ 0

- A C r

FIG. 27. VARIATION OF - ^ — " WITH AC, „ L E . BLOWING ONLY. _{CL LB}

-° 1°.^" 1 > B 1 . - . . 1 » ^ ^ V . C)l.OI78

^-O^^

- ^ - " ^ ' i ^ . ^ _ • • 0 ° 2 ° ^ ° A F I G . 2 8 . V A R I A T I O N O F C R O S S - W I N D FORCE WITH SIDESLIP A N G L E ^ , ^ » O " B L O W I N G F R O M A L L E D G E S .

FIG 2 9 VARIATION OF CROSS WIND FORCE WITH SIDE SLIP ANGLE fl * • 3 0 * BLOWING FROM A L L E D G E S .

- O I O -©—o—G . ó -" -" • * - • « . " ^ ^ - ^ S ^ ,

—^rc

^•._ '-= ^ C . l * . 0 * l * . 3 0 * 0 0178 • O

FIG.30. VARIATION OF PITCHING MOMENT ABOUT MID ROOT CHORD WITH LIFT. BLOWING FROM ALL EDGES.

C « - 0 0 4 • t. x X • C|i=0 o u

^i

X SYMBOL • O • T + X " ^ > o° 2 0 ° 5 0 * 6 0 ° 7 0 ° • O * " • ^ ^ . C ^ i s O

FIG. 32. VARIATION OF PITCHING MOMENT ABOUT MID-CHORD WITH LIFT. L.E. BLOWING ONLY . «• » O.

i t . O °

1/^

/ / 1 1 • / / / / ^—''• / / / _ ' t" 1 < ^ . I O ° 1 1 — 1 1 1 1 1 1 y ^ * . 1 S ° / / /* / / i 1 1 1 1 - = 2 0 ° t I „^C^.o-sa. / / , » - » C / < - 0 3 l 8 ^C)l'OI35 1 < ^ ^ C | J . 0 - 0 5 J ^ 0 ^ < j . C ) < . 0 0 2 4 ^C)|.0

FIG.31. VARIATION OF PITCHING MOMENT ABOUT MID ROOT CHORD WITH LIFT L.E. BLOWING ONLY. • = 9 = 0 .

1 1

l ^ ^ ^ ^ s . ^

.o°— 3 5° 3iO^ 1 5 ; a 0° fl • 5°

= i i ^ ^ ^

yo% C>i.O

r9M

— — - . ^ " V ^ s ' ' - * — * « . ^ " ^ Q l O " ^ ^ ^ « i . . ^ * * * ^ X 15 * ^ * ^ ^ ^ i 20'

r"°

-5» O* ^ + 5 °

FIG.33. VARIATION OF ROLLING MOMENT WITH SIDESLIP ANGLE ,8. BLOWING AT ALL EDGES. ^ > 0 * - 0 2 0 ^Q "W-V •> . rl ^ ' " ~ ' -^ < -^ , , -^ ,""'''•--_{' ^ • ^ ^ n}

•

^

" ' • • ' = ' - • . . 1 ~*~-i:Sia CK 0 0-178 «.o» • 4 "*^ ^^^,, ^^^» * . 3 0 ° 0 )»^ •^•s:; 1 . o 0 2 0 0 40 0 6 0 C|_ c e o 1 0 0 1-20

FIG. 35, VARIATION OF {^ ( ' S ^ - ^ * ' T " LIFT. BLOWING FROM ALL EDGES. <* s O" « 3 0 * .

0 ° « o 4

FIG 34 VARIATION OF ROLLING MOMENT WITH SIDESLIP ANCLE JB. BLOWING AT ALL CDCCS 4 . 3 0 ° 0 02 Cn ^ ^ ^ . ï__r7 *"**^-^*?"*'*'^. i . O

1 - ^

K • o"._ 3 10" I 2 t f ^ ^ 2 5 °

FIG.36. VARIATION OF YAWING MOMENT WITH SfDCSLIP ANCLE S. BLOWING FROM

*Vl

: ^ ^ ^ > 2 ° 15° 110° .15° ^^o°_ -5° ' + 2 S *

FIG.37 VARIATION OF YAWING MOMENT WITH SIDESLIP ANGLE 0. BLOWII«: FROM ALL EDGES. • • l O ^ - C p / * /^. C ^ . O jStO o 2 0 l-OO I 2 0 FIG.3». VARIATION OF n, ^ | ^ " ) ALL EDGES, ^ x o " 1 3 0 ? O 6 0 c,^ O BO

WITH L I F T . BLOWING FROM

/ /

l^

z

1 / \)S.

_{K ^}

FIG.40. VARIATION OF SPANWISE STATIC PRESSURE DISTRIBUTIONS WITH x . ^ ^O

FIG. 41. SPANWISE VARIATION OF UPPER AND LOWER SURFACE STATIC PRESSURE DISTRIBUTIONS, ^/c • 0 - 4 9 .

F I G . 4 3 . SURFACE STATIC PRESSURE DISTRIBUTIONS. COMPARISON WITH THEORY.

2 . S ( ) O 1 q / ^ X ° X

/^ov7/

// / f / ! / / • / / / ' ' / 1 1 / / / t _{O S si I O}

FIG.45a S 45b. VORTEX CORE POSITIONS WITH AND WITHOUT LEADING EDGE BLOWING.

^ ' ^ ' ^ ^ s o

-FIC.47 VARIATION OF TOTAL HEAD THROUGH VORTEX. Q l « 0 , ^ % ^ OSO.

FIG.46a. VORTEX CORE POSITIONS, WITH AND WITHOUT LEADING EDGE BLOWING.

V--^v <-l"° »'SO° 0^^Cfi«O 092 0,1.0 048 4, 0 7 Yi_ 0 8 S

RG. 4 6 b. VORTEX CORE POSITIONS WITH AND WITHOUT LEADING EDGE BLO%VING.